CanSat 2018 Critical Design Review (CDR) Outline Version 1cansatcompetition.com/docs/teams/Cansat2018_5002_CDR_v12.pdfCanSat 2018 CDR: Team 5002 4 Acronyms HS Heat shield CDH Communications

CanSat 2018 CDR: Team 5002 1

CanSat 2018

Critical Design Review (CDR)

Outline

Team #5002

Manchester CanSat Project


Presentation Outline

Presenter: Iuliu Ardelean

Section Presenter

Presentation Outline Iuliu Ardelean

Systems Overview Lawrence Allegranza France

Sensors Subsystem Overview Iuliu Ardelean

Descent Control Design Iuliu Ardelean

Mechanical Subsystem Design Lawrence Allegranza France

Communications and Data Handling Subsystem Design Lawrence Allegranza France

Electrical Power Subsystem Lawrence Allegranza France

Flight Software Design Lawrence Allegranza France

Ground Control System Design Lawrence Allegranza France

CanSat Integration and Testing Lawrence Allegranza France

Mission Operations and Analysis Iuliu Ardelean

Requirements Compliance Iuliu Ardelean

Management Iuliu Ardelean

This review follows the sub-sections listed below:


Team Organization

Team Member Responsibility

Iuliu Ardelean (IA) CE, SE, GCS

Nicole Zieba (NZ) CDH, SE, GCS, PM

Lawrence Allegranza France (LAF) I&T, CDH

Xisco Jover (XJ) EPS, CDH, SE

Robert Stana (RS) FSW, SE

Alex Shelley (AS) ME, DCS

Davis Joseph (DJ) ME, DCS

Julia Stankiewicz (JS) ME, DCS

Zair Chaudhry (ZC) ME, DCS

Nacho Salsas Leon (NSL) ME, DCS

Iuliu Ardelean

Chief Engineer

Mechanical Subsytem

Alex Shelley

3rd Year

Davis Joseph

4th Year

Julia Stankiewicz

2nd Year

Zair Chaudhry

3rd Year

Nacho Salsas

3rd Year

Electronics Subsytem

Iuliu Ardelean

3rd Year

Xisco Jover

3rd Year

Nicole Zieba

4th Year

Lawrence Allegranza France

4th Year

Robert Stana

3rd Year

Integration and testing


Ground Control Station

Iuliu Ardelean

Nicole Zieba

Project Manager

Kate Smith

Faculty Advisor

Matt Hogg

Leadership Mentor


Acronyms

HS Heat shield

CDH Communications and Data Handling

EPS Electrical Power Subsystem

FSW Flight Software

GCS Ground Control Station

ME Mechanical Subsystem

SE Sensors Subsystem

DCS Descent Control Subsystem

CE Chief Engineer

PM Project Manager

I&T Integration and Testing

CONOPS Concept of Operations

GUI Graphical User Interface

A Analysis

I Inspection

T Testing

D Demonstration

TBC To be confirmed

TBD To be determined

RE# Top Level Requirement

SL System Level

SSL Subsystem Level

IDE Integrated Development Environment

RTC Real Time Clock

I2C Inter-Integrated Circuit

SPI Serial Peripheral Interface

ADC Analog to Digital Converter

EEPROM Electrically Erasable Programmable Read-

only memory

MCU Microcontroller Unit

CCU Central Control Unit (Mechanical)


System Overview


Presenter: Lawrence Allegranza France


Mission Summary

Objectives:

1. Build a CanSat with an atmospheric-sampling Probe, a single hen’s Egg, a Heatshield and a Parachute.

2. The CanSat shall be launched in a sounding rocket to an altitude of 675-725 meters.

3. After release from rocket payload bay, the CanSat shall deploy the Heatshield, and descend to an

altitude of 300 meters without tumbling.

4. At 300 meters, the CanSat shall release the Heatshield and deploy a Parachute. The Heatshield will

descend on its own at a rate of 5 m/s.

5. The Probe shall collect and transmit atmospheric data to a Ground Control Station in real-time,

throughout its operation phase.

6. The Proble shall land leaving the egg intact, after which it will continuously operate an audio beacon.

7. The Ground Control Station shall receive and display CanSat data.

Selectable Bonus and Rationale:

• Camera Bonus selected because of abundant team members experience.

External Objectives:

• Continue to deliver Manchester CanSat Project weekly, educational, space-related Workshops towards

University of Manchester STEM Students.

• Develop a UK CanSat Competition.

• Inspire other UK Universities and Academic Institutions to adopt the Manchester CanSat Project model

to create a network of CanSat societies across the UK.


Summary of Changes Since PDR


There are 4 changes since the PDR that are worth mentioning here.

Most other changes were done to improve weight – which has always been our biggest constraint.

PDR CDR RATIONALE

Multiple

Independent

Mechanisms

Single

Unified

Mechanism

Lighter weight.

Primitive Mounting Plate PCB Improved robustness of the electric circuit and

overall design.

Two MCUs One MCU Lighter Weight.

Adafruit Serial JPEG

Camera

with NTSC Video

Modified

SQ11 Pawaca

Camera

High Definition. Reduced circuit complexity and

overall weight. Possible to make ‘legal’ for the

competition.


System Requirement Summary


RE# Description A I T DRE1 Total mass of the CanSat (probe) shall be 500 grams +/- 10 grams. X X

RE2 The aero-braking heat shield shall be used to protect the probe while in the rocket only and when deployed from the rocket.

It shall envelope/shield the whole sides of the probe when in the stowed configuration in the rocket. The rear end of the

probe can be open

X X

RE3 The heat shield must not have any openings. X

RE4 The probe must maintain its heat shield orientation in the direction of descent. X

RE5 The probe shall not tumble during any portion of descent. Tumbling is rotating end-over-end.

RE6 The probe with the aero-braking heat shield shall fit in a cylindrical envelope of 125 mm diameter x 310 mm length.

Tolerances are to be included to facilitate container deployment from the rocket fairing.X X

RE7 The probe shall hold a large hen's egg and protect it from damage from launch until landing. X X

RE8 The probe shall accommodate a large hen’s egg with a mass ranging from 54 grams to 68 grams and a diameter of up to

50mm and length up to 70mm.X

RE9 The aero-braking heat shield shall not have any sharp edges to cause it to get stuck in the rocket payload section which is

made of cardboard.X X X

RE10 The aero-braking heat shield shall be a florescent color; pink or orange. X X

RE11 The rocket airframe shall not be used to restrain any deployable parts of the CanSat. X X

RE12 The rocket airframe shall not be used as part of the CanSat operations. X X

RE13 The CanSat, probe with heat shield attached shall deploy from the rocket payload section. X

RE14 The aero-braking heat shield shall be released from the probe at 300 meters. X X X

RE15 The probe shall release a parachute at 300 meters. X X X

RE16 All descent control device attachment components (aero-braking heat shield and parachute) shall survive 30 Gs of shock. X X

RE17 All descent control devices (aero-braking heat shield and parachute) shall survive 30 Gs of shock. X X

RE18 All electronic components shall be enclosed and shielded from the environment with the exception of sensors. X

RE19 All structures shall be built to survive 15 Gs of launch acceleration. X X

RE20 All structures shall be built to survive 30 Gs of shock X X

RE21 All electronics shall be hard mounted using proper mounts such as standoffs, screws, or high performance adhesives. X



RE# Description A I T DRE22 All mechanisms shall be capable of maintaining their configuration or states under all forces X

RE23 Mechanisms shall not use pyrotechnics or chemicals. X

RE24 Mechanisms that use heat (e.g., nichrome wire) shall not be exposed to the outside environment to reduce potential risk

of setting vegetation on fire.X X X

RE25 During descent, the probe shall collect air pressure, outside air temperature, GPS position and battery voltage once per

second and time tag the data with mission time.X X X X

RE26 During descent, the probe shall transmit all telemetry. Telemetry can be transmitted continuously or in bursts. X X

RE27 Telemetry shall include mission time with one second or better resolution. Mission time shall be maintained in the event

of a processor reset during the launch and mission.X X

RE28 XBEE radios shall be used for telemetry. 2.4 GHz Series 1 and 2 radios are allowed. 900 MHz XBEE Pro radios are also

allowed.X

RE29 XBEE radios shall have their NETID/PANID set to their team number. X X X

RE30 XBEE radios shall not use broadcast mode. X X X

RE31 Cost of the CanSat shall be under $1000. Ground support and analysis tools are not included in the cost. X X

RE32 Each team shall develop their own ground station. X

RE33 All telemetry shall be displayed in real time during descent. X X

RE34 All telemetry shall be displayed in engineering units (meters, meters/sec, Celsius, etc.) X X

RE35 Teams shall plot each telemetry data field in real time during flight X X

RE36 The ground station shall include one laptop computer with a minimum of two hours of battery operation, XBEE radio and

a hand held antenna.X

RE37 The ground station must be portable so the team can be positioned at the ground station operation site along the flight

line. AC power will not be available at the ground station operation site.X

RE38 Both the heat shield and probe shall be labeled with team contact information including email address. X

RE39 The flight software shall maintain a count of packets transmitted, which shall increment with each packet transmission

throughout the mission. The value shall be maintained through processor resets.X X



RE# Description A I T DRE40 No lasers allowed. X X

RE41 The probe must include an easily accessible power switch. X X X

RE42 The probe must include a power indicator such as an LED or sound generating device. X X X

RE43 The descent rate of the probe with the heat shield deployed shall be between 10 and 30 meters/second. X X

RE44 The descent rate of the probe with the heat shield released and parachute deployed shall be 5 meters/second. X X

RE45 An audio beacon is required for the probe. It may be powered after landing or operate continuously. X X X

RE46Battery source may be alkaline, Ni-Cad, Ni-MH or Lithium. Lithium polymer batteries are not allowed. Lithium cells must be

manufactured with a metal package similar to 18650 cells.X X X

RE47An easily accessible battery compartment must be included allowing batteries to be installed or removed in less than a

minute and not require a total disassembly of the CanSat.X X X

RE48Spring contacts shall not be used for making electrical connections to batteries. Shock forces can cause momentary

disconnects.X

RE49A tilt sensor shall be used to verify the stability of the probe during descent with the heat shield deployed and be part of the

telemetry.X X X

Bonus

1

Camera: Add a color video camera to capture the release of the heat shield and the ground during the last 300 meters of

descent. The camera must have a resolution of at least 640x480 and a frame rate of at least 30 frames/sec. The camera

must be activated at 300 meters.

X X X X

Bonus

2

Wind Sensor: A radio transmitter shall be added to transmit the wind speed by changing its 10 frequency. The frequency

change shall be 1 Hz per 0.1 meter/sec. The transmitter must be custom designed and built. It cannot be a commercial

product. The frequency must be in the 433 MHz ISM band or if a team member has an amateur radio license, an amateur

radio band can be used. The transmitter must be able to be set to 8 different frequencies in the 433 MHz ISM band with 25

KHz separation. The transmitter must turn off after the probe lands to minimize interference. The team can use a commercial

receiver.

X X X X


System Concept of Operations

0. Pre -Launch

1. Launch

2. HS Deployment

3. CanSatDescent

4.1. HS Release

4.2. Parachute

Deployment

5. Probe Descent

6. Landing 7. Recovery 8. Data Handover

ROLES & RESPONSIBILITIES

Mission Control Officer: NZ.

Ground Station Crew: IA, LAF, NZ, RS.

Recovery Crew: AS, NSL, DJ, ZC, JS, XJ.

CanSat Crew: IA, LAF, RS, AS, NSL, DJ, ZC, JS, XJ.




0. Pre -Launch

• CanSat Switched on

• Telemetry transmitting start

1. Launch

• CanSat insertion in Rocket Payload Bay

• Rocket ignition and ascension

• Apogee Reached

2. HS Deployment

• Rocket and nose cone separation

• CanSat deployed from rocket Payload Bay

• HS deploys

• Rocket and nose cone descent

3. CanSatDescent

• CanSat descent with Heatshield deployed

4.1. HS Release

• 300 m altitude sensed by Probe

• HS released

• Release captured by Camera




4.2. Parachute

Deployment

• 300 m altitude sensed by Probe

• Parachute deployed

5. Probe Descent

• Probe descent with Parachute

• HS tumble down on its own

• Descent captured by Camera

6. Landing

• Telemetry transmitting Stop

• Audio Beacon Activation

7. Recovery

• Audio Beacon Operational

• All systems recovered (including HS)

• CanSat switched off

8. Data Handover

• Data formatted and saved to USB

• USB handed over to officials

• Data analysis and reduction for PFR



Payload Physical Layout

248.5 mm

120 mm

409 mm

Deployed Heat Shield PayloadStowed Payload

61.5 mm

2.5

mm

2.5

mm

Stowed Payload inside Rocket

290 mm




Heat Shield Release +

Parachute Deployment InitiationParachute Deployed




A

B

D

H

F

J

C

E

G

K

12

13

9

14

15

L

M

N

P

12

3

4

5

6

7

8

10

11

A - Parachute Hatch Hinge (Nylon) J – Parachute Hatch Spring 2 – 118 mm 10 – 115 mm

B – Parachute Stowed (RS Nylon) K – GPS & Buzzer 3 – 54 mm 11 – 78 mm

C – Stickers L - Support Column (CF) 4 – 23.5 mm 12 – 190 mm

D – HS Rod (Carbon Fibre) M – 15 mm Rod (Nylon) 5 – 15 mm 13 – 290 mm

E – On/Off Switch N – Pivot Rod (ABS) 6 – 30 mm 14 – 222.1 mm

F – Battery P – Nose Cone Springs 7 – 130 mm 15 – 248.5 mm

G – Camera Q – Egg 8 – 30 mm

H - Servo 1 – 175 mm 9 – 100 mm

Q







Egg with padding

Heat Shield Assembly

Stowed Parachute

in Parachute Bay

Electronics Board

Central Control Unit

(CCU) in Camera Bay

Camera

Battery

Electronics Cover



Launch Vehicle Compatibility

DAVIS

Available Volume (as per Competition Requirements):

Diameter : 125 mm

Height : 310 mm

CanSat Volume:

Diameter : 120 mm

Height : 248.5 mm

Clearance : More than 2.5 mm throughout

No sharp protrusions

Dimensions account for ease of fit and deployment.

The heat shield envelopes the sides of the probe to protect it.

A test launch has been performed at the University of Manchester to

verify Launch Vehicle Compatibility. The results confirm that the

CanSat is indeed compatible.

61.5 mm

2.5

mm

2.5

mm

Stowed Payload inside Rocket



Sensor Subsystem Design

Iuliu Ardelean


Sensor Subsystem Overview

Selected Component Function

Adafruit 10 DOF IMU Determine Pressure, Temperature, Altitude and Tilt.

Adafruit Ultimate GPS v3.0 Determine GPS Position

Modified SQ11 Pawaca Camera Modified 720p Video Camera

ADC + Voltage Divider Battery voltage

TEENSY MICROCONTROLLER

ADAFRUIT

10 DOF IMU

ADAFRUIT

ULTIMATE

GPS

MODIFIED

SQ11

CAMERA

I2C SERIAL DIO

VOLTAGE

DIVIDER

FROM

BATTERY

ADC

SD CARD

BREAKOUT

SPI

RTC

NO LASERS.


Sensor Changes Since PDR


PDR CDR RATIONALE

Serial JPEG TTL Camera Modified SQ11 Pawaca Camera - Would have required

additional recorder.

- Better video resolution (from

SD to HD and FHD)

- Lighter



Sensor Subsystem Requirements

RE# Description VERIFICATION

A I T D

RE1 Total mass of the CanSat (probe) shall be 500 grams +/- 10 grams. X

RE6 The probe with the aero-braking heat shield shall fit in a cylindrical envelope of 125 mm diameter x 310

mm length. Tolerances are to be included to facilitate container deployment from the rocket fairing.

X

RE15 The probe shall release a parachute at 300 meters. X X

RE18 All electronic components shall be enclosed and shielded from the environment with the exception of

sensors.

X

RE21 All electronics shall be hard mounted using proper mounts such as standoffs, screws, or high performance

adhesives.

X

RE25 During descent, the probe shall collect air pressure, outside air temperature, GPS position and battery

voltage once per second and time tag the data with mission time.

X X X

RE31 Cost of the CanSat shall be under $1000. Ground support and analysis tools are not included in the cost. X

RE46 Battery source may be alkaline, Ni-Cad, Ni-MH or Lithium. Lithium polymer batteries are not allowed.

Lithium cells must be manufactured with a metal package similar to 18650 cells.

X

RE49 A tilt sensor shall be used to verify the stability of the probe during descent with the heat shield deployed

and be part of the telemetry.

X X

B1 Video Camera X X

B2 Wind sensor X X

SE1 Pressure Sensor should be accurate. X


Probe Air Pressure Sensor

Summary

Name Weight/

Size

Cost Power Operational

Environment

Accuracy

/Error

Resolution Drift Interface Other

Adafruit 10

DOF IMU

2.8 g 40 GBP 1 mA x 3.6 V 30 – 110 kPa 12 Pa 2-6 Pa 100 Pa/yr I2C All in

one38x23x3

mm

3-32 uA x 3.6 V 9000 to – 500 m 1 m 0.17-0.5 m

Pressure data will be collected and processed with the help of the Adafruit BMP180 Library.

In order to calculate altitude, the altitude standard equation can be used as follows:


Probe Air Temperature Sensor

Summary

Name Weight/Size Cost Power Operational

Environment

Accuracy

/Error

Interface Other

Adafruit 10 DOF

IMU

2.8 g 40 GBP 1 mA x 3.6 V -40 to +85 degC

(0 to +65 full

accuracy)

2ºC I2C All in one

38x23x3 mm 3 – 32 uA x 3.6 V

Temperature data will be collected and processed with the help of the Adafruit BMP180 Library.


GPS Sensor Summary

Component Weight/Size Cost Power Operational

Environment

Accuracy/

Error

Interface Other

Adafruit Ultimate

GPS Breakout

8.5g 40 GBP 20mA x 3.3V 515 m/s 3 meters Serial Warm/cold start:

34 seconds25.5mm x 35mm x

6.5mm

GPS data will be collected and processed with the help of the Adafruit GPS Library.


Probe Voltage Sensor Summary

A simple Voltage Divider and the inbuilt Teensy ADC

can be used to measure the voltage of the battery.

The Teensy ADC can only measure values of up to 3.3 V,

therefore the need for a Voltage Divider. The resistances

values will be so that 𝑅1 = 2 ∗ 𝑅2 . Preliminary testing

suggests 𝑅1 = 2 ∗ 𝑅2 = 24 𝑘𝑂ℎ𝑚 are suitable values.

The Teensy’s ADC is 10-bit, hence the

resolution/accuracy that we get is of the order of 9 mV.

In code, the system is easy to implement:

𝑉𝑖𝑛 = 𝑎𝑛𝑎𝑙𝑜𝑔𝑅𝑒𝑎𝑑 𝑎𝑛𝑎𝑙𝑜𝑔𝑃𝑖𝑛 ∗3.3

1023∗𝑅2 + 𝑅1𝑅2


Tilt Sensor Summary

Name Weight/Size Cost Power Operational

Environment

Accuracy/

Error

Interface Other

Adafruit 10 DOF

IMU

2.8 g 40 GBP 1 mA x 3.6 V 0-360 3º I2C All in One

38x23x3 mm 3-32 uA x 3.6 V

Data will be collected and processed with the help of the Adafruit LSM303 Library.


Bonus Objective Camera Summary

The SQ11 Pawaca Camera offers HD and FHD

(greater than 640 x 480) @ 30fps color video camera

functionality.

This camera has an inbuilt SD Card slot on which the

video can be stored. This means there will no longer be

necessary to add a video recorder or and SD Card

Breakout Board.

The camera will be disassembled, to save weight (from

40 to 5 grams) and to replace the mechanical switch

with an electrical one. Hence it will be possible to

activate the camera at 300 meters, by operating the

electrical switch, right after the HS release, etc.

The camera’s battery will be removed, and the camera

will be powered from the voltage regulator.

The main challenge associated with this modified

camera is the really delicate soldering job required.

Selection Rationale: Previous Experience.


Receiver• Frequency down conversion


Bonus Objective Wind Sensor


Radio• Teensy uses GPS data to calculate wind speed, outputting digital

voltage signal according to speed.

• Through serial-to-parallel and DAC, variable capacitor is controlled,

changing the frequency 1 Hz for every 0.1 m/s change.

• C4, C3, and L2 makeup the Colpitts Oscillator tank.

• C3 is set on the ground to set to one of 8 channels.

• C4 is controlled by the voltage from the DAC.

• Frequency mixer used to get frequency in 433 MHz range ISM band

Counter-rationale

(Rationale for not choosing this bonus):

The team has more experience working with

video cameras, making this bonus less attractive.


Descent Control Design

Iuliu Ardelean


Descent Control Overview

Descent Control System

•This will include 2 main subsystems – a heat shield (already set to

a predetermined deployment angle) and a parachute.

•The heat shield envelopes the sides of the probe to protect it.

Deployment and release order

• Probe is released at an altitude of 675-725 meters and the aero breaking

heat shield, covering the whole probe, opens which results in descent rate

being kept between 10-30 m/s.

• At 300 meters heat shield is dropped and the parachute is deployed almost

simultaneously.

•The descent rate is decreased to 5 m/s which is slow enough for the egg to

remain intact after landing.

The next slide shows required diagrams.

Heat shield projected surface

area0.084 𝑚2

Parachute projected surface

area0.243 𝑚2

Fabric for the Parachute (rip stop nylon)

6 strings for the Parachute

Fabric for the Heat Shield (rip stop nylon)

4 carbon fibre rods

4 zip ties

4 rod pivots

1 3D printed Nose Cone

4 springs

Necessary components



Descent Control Overview

1st stage – Apogee to 300 m

1st event – deployment of the

heat shield

2nd stage – 300 m to Landing

(events take place simultaneously)

2nd event – parachute

deployment

3rd event – heat

shield separation

4th event – heat

shield free flight

0th stage – Undeployed configuration



Descent Control Changes Since

PDR

PDR CDR Rationale

Torsion springs hold the carbon fibre

rods in place for the HS.

3D printed pivots hold the carbon

fibre rods in place for the HS. They

pivot about some arms on the nose

cone.

The torsion springs attachment method to the CF rods

was weak and had to be improved. ABS rod holders

were designed and held in place to act as a pivot. They

were attached to the nose cone with extensions

springs to maintain tension so the natural position was

in the deployed position.

Circular shape of the Nose Cone Squared shape of the Nose Cone

Upon realisation the shape of the Nose Cone has been

changed from circular to square as it is easier to attach

a squared HS fabric compared to a circular geometry.

Lower attachment of Heat Shield

fabric – Circular shape

Lower attachment of Heat Shield

fabric – Square shape

Upon realisation the shape of the HS fabric has been

changed from circular to square as it is easier to stitch

and attach a squared HS fabric compared to a circular

geometry

A solenoid retracts from inside the

HS attachment point to release the

HS. Springs push the nose cone

away once released.

A servo with a length of metal rod

retracts to release the HS. The

springs have been removed.

It was decided that one dedicated mechanism for HS

deployment, HS release and parachute release was

preferable to multiple mechanisms. This was decided

as it reduces weight, reduced the complexity of

electronics and the flight software and is easier to

implement. Springs were removed upon realisation

during testing, as separation took place regardless of

their existence.


The following 3 slides describe changes since PDR and prototype testing.



PDR

PDR CDR Rationale

The HS is deployed through an

independent bay that has an ‘X’

shaped device which deploys the HS

when activated with a solenoid.

A central system that uses a servo

with a metal rod also controls the

deployment of the HS, as well as the

release of the HS and parachute.

The rotating gate which unhooked the rods to deploy

the heat shield has been replaced with a much simpler

string system. Upon testing the rotating gate

mechanism was deemed unreliable. PLA and ABS

both are not stiff enough to rotate fully if a force is

applied on one side of the gate, the gate flexed and

deformed unhooking only three out of four rods. The

actuator was too large and heavy to fit in the

deployment bay

The nose cone has four hollow pillars

to hold four springs to aid release.

There are four points for attaching

torsion springs. A central pillar is

used for release.

The hollow pillars have been

dismissed as the springs aren’t

necessary. They are now just pillars.

The torsion springs have been

replaced by 3D printed rod pivots that

pivot about arms on the nose cone.

The overall shape of the nose cone

has changed as the circular edges

have been removed for straight lines.

The pillars remain to provide a stable base for the rest

of the CanSat and it also stops additional stresses on

the servo rod against the nose cone attachment point.

The torsion spring idea proved to be completely

unreliable and unrealistic. It has been replaced by a

more reliable and smoother method. There have been

many changes to the nose cone to further reduce the

weight of the CanSat. The circular edges were

removed to save weight and to ease the difficulty of HS

manufacture.

A servo opens the parachute bay

hatch to release the parachute.

A spring has been used so that the

natural position is open. A string

holds it closed. Therefore, when the

string is released, the parachute bay

hatch will open. Height of the bay has

also increased.

The use of a servo for only parachute bay hatch

opening was undesirable due to weight, centre of mass

and electronics issues. By using a spring, it reduces

weight and reduces the strain on the electronics

systems.




PDR

Date Test Outcome

08/03 Parachute Descent Rate Parachute surface area proven to be correct as

descent rate average at 5 m/s.

11/03 Whole probe test

Heat Shield proven to be released successfully.

Heat Shield descent rate proven to be 5 m/s post

separation.

Parachute structural strength proven to be insufficient.



Descent Control Requirements

•

•

•

•

•


A I T D


RE2The aero-braking heat shield shall be used to protect the probe while in the rocket only and when deployed

from the rocket. It shall envelope/shield the whole sides of the probe when in the stowed configuration in the

rocket. The rear end of the probe can be open

X X


RE4 The probe must maintain its heat shield orientation in the direction of descent. X

RE5 The probe shall not tumble during any portion of descent. Tumbling is rotating end-over-end. X X

RE6The probe with the aero-braking heat shield shall fit in a cylindrical envelope of 125mm diameter x 310mm

length. Tolerances are included to facilitate container deployment from the rocket fairing.x

RE9The aero-braking heat shield shall not have any sharp edges to cause it to get stuck in the rocket payload

section which is made of cardboard.X





RE14 The aero-braking heat shield shall be released from the probe at 300m. X X

RE15 The probe shall release a parachute at 300m. X X

RE16All descent control device attachment components (aero-braking heat shield and parachute) shall survive

30Gs of shock. X X


Descent Control Requirements

•

•

•

•

•


A I T D

RE17 All descent control devices (aero-braking heat shield and parachute) shall survive 30Gs of shock. X X

RE18All electronic components shall be enclosed and shielded from the environment with the exception of

sensors.X

RE19 All structures shall be built to survive 15Gs of launch acceleration. X X

RE20 All structures shall be built to survive 30Gs of shock. X X


RE38 Both the heat shield and probe shall be labelled with team contact information including email address. X

RE43 The descent rate of the probe with the heat shield deployed shall be between 10 and 30 meters/second. X X

RE44 The descent rate of the probe with the heat shield released and parachute deployed shall be 5 meters/second. X X

RE44The descent rate of the probe with the heat shield released and parachute deployed shall be 5

meters/second.X X


Payload Descent Control Hardware

Summary

Stowed Configuration (Heatshield)

• All subsystems (heat shield and parachute) will be controlled by a main servo, situated in the lower bay of the main

probe.

• The servo will be placed in a casing containing already set indents where each of the subsystem’s deployment

mechanisms will be introduced.

• The servo will be connected to a prolongation (metal rod) of its mechanical arm. This arm will be responsible for

holding the subsystems from separating.

• Component sizing was given in DCS Overview. The CanSat in stowed configuration is sized to fit inside the Rocket

Payload Bay.

Servo

Casing

Heat shield release

Heat shield deployment

Parachute deploymentMetal rod

Stowed Configuration Section View


The following 6 slides describe DCS Hardware summary.



Summary

Deployment Method (Heat Shield)

Spring

3D printed parts

• Aforementioned servo will rotate the arm at a calculated angle of 10 deg.

• This will cause the fist of the indents to be free of the arm extension causing the string to be released. The string

goes around all four rods and then hooks back on to itself to maintain stowed position.

• For simplicity, the string is hooked to itself in the stowed position, before attaching the HS fabric over the rods.

• The springs found at the base of the heat shield will exert a rotating motion on the carbon fibre rods, causing the

full deployment of the heat shield

• The heat shield has been designed to have a maximum opening projected surface area of 0.084 m^2

Servo armHeat shield deployment

mechanism cavity (CCU)


HS ring secured in

the cavity (CCU)

String

Hook



Summary

Deployed Configuration (Heat Shield)




Summary

Stowed and Deployed Configuration (Parachute)

Stowed Deployed




Summary

Deployment Method (Parachute)

• Again, on receiving a PWM signal from the MCU (triggered by altitude), the servo will rotate the arm another 85

degrees. This will simultaneously release the parachute’s deployment mechanism and the heat shield.

• On loosing its tension, the string attached to the parachute’s payload bay cover will not have sufficient force to

resist the moment created by the tension spring.

• The spring will open the bay’s cover, where the parachute will be ejected through a separate spring mechanism

system.

Spring

String Spring




Summary

• The Descent Control hardware only uses a passive system.

• The heat shield will have a sufficient projected area to achieve a descending rate of 13 m/s with the probe

mounted, and 5 m/s on free flight. The projected areas for HS and Parachute are in the table below. Full

details on slides: 16 and 59.

• The squared shape of the heat shield will aid in keeping static stability and the nadir motion.

• The location of the C.G., being below the C.P., will prevent tumbling and help recover the nadir direction.

• The springs located in the base of the heat shield and attached to the 3D printed parts (rod pivots) will apply

sufficient moment force to maintain the heat shield deployed throughout all its descent stages.

• Heat shield selected to be orange for ease of retrieval

• Zip ties will be used to attach HS fabric (rip stop nylon) to the carbon fibre structural rods

Component Sizing & Key Considerations

Top View Side View Rod pivot + Spring


Heat shield projected surface area 0.084 𝑚2 Parachute projected surface area 0.243 𝑚2

Descent Stability Control Design

45CanSat 2018 CDR: Team 5002

The heat shield uses a passive design, as it saves weight and simplifies the system.

The shape is simplified from a truncated cone to a truncated pyramid for ease of

stitching and attachment to nose cone and rods. It allows aerodynamic stability by

passing the flow around to the payload smoothly and behaves like a finned design.

The centre of gravity (cg) is placed lower than the centre of pressure (cp) of the heat

shield to maintain the nadir direction.

In order to prevent tumbling there is a considerable difference between cg and cp

improved by removing the deployment gate and parachute release servo replaced by a

single servo and battery both placed at the bottom of the payload.

Battery Servo

Truncated

pyramid



Descent Rate Estimates•

•

•

1. 𝑾 = 𝑫 =𝟏

𝟐𝝆𝒗𝟐𝑪𝑫𝑺 𝑫 – drag force acting on the probe

𝑾 – weight of CanSat/Probe

𝝆 – air density

𝒗 – terminal velocity

𝑪𝑫 – drag coefficient

𝑺 – projected surface area of descending object

2. 𝑺 = 𝝅𝒓𝟐 ⟶ 𝒓 =𝟐𝑫

𝝆𝑪𝑫𝝅𝒗𝟐𝒓 – radius of projected surface area




•

•

Assumptions:• Weight of the falling object is equal to drag when it travels with constant velocity (terminal velocity),

• Density of air is assumed to be 𝟏. 𝟐𝟐𝟓𝒌𝒈

𝒎𝟑,

• No wind or air currents (depending on weather conditions the heat shield size can be adjusted).

the heat shield envelopes

the probe completely to protect it

1. 𝒎𝟏 = 𝟎. 𝟓 𝒌𝒈

2. 𝒎𝟐 = 𝟎. 𝟒𝟏𝟕 𝒌𝒈

3. 𝒎𝟑 = 𝟎. 𝟎𝟖𝟑 𝒌𝒈

4. 𝝅 = 𝟑. 𝟏𝟒𝟏𝟓

5. 𝒈 = 𝟗. 𝟖𝟏𝒎

𝒔𝟐

𝑪𝑫𝒑- drag coefficient of parachute

𝑪𝑫𝒉𝒔- drag coefficient of heat shield

Outputs:

𝑺𝒑- projected area of the parachute

𝑺𝒉𝒔 - projected area of Heat Shield

𝒎𝟏- mass of CanSat (prior to separation of the Payload)

𝒎𝟐- mass of Probe following separation from the Heat Shield

𝒎𝟑- mass of Heat Shield following release from the Payload

𝒈- gravitational acceleration




•

•

Heat Shield

(before separation)

• Estimated drag coefficient (𝑪𝑫𝒉𝒔)*: 0.55

𝟏𝟎𝒎

𝒔< 𝒗 < 𝟑𝟎

𝒎

𝒔

𝟐𝒎𝟏𝒈

𝟑𝟎𝟐 × 𝝆𝑪𝑫𝒑< 𝑺𝒉𝒔 <

𝟐𝒎𝟏𝒈

𝟏𝟎𝟐 × 𝝆𝑪𝑫𝒑

𝟎. 𝟎𝟏𝟔𝒎𝟐 < 𝑺𝒉𝒔 < 𝟎. 𝟏𝟒𝟔 𝒎𝟐

• Chosen surface projected area of HS is

𝟎. 𝟎𝟖𝟒𝒎𝟐

* Estimate based on drag coefficients of the hemisphere and cone

** Estimate based on HS falling down without turning upside down

Parachute• Estimated drag coefficient (𝑪𝑫𝒑): 1.1

• Required velocity (𝒗𝟐): 5 𝒎

𝒔

𝑺𝒑 =𝟐𝒎𝟐𝒈

𝝆𝒗𝟐𝑪𝑫𝒑= 𝟎. 𝟐𝟒𝟑𝒎𝟐

• Area of the spill hole is chosen to be 3%

of the total parachute projected area

Heat Shield

(after separation)

• Estimated drag coefficient (𝑪𝑫𝒉𝒔′)**: 0.55

𝒗 =𝟐𝒎𝟑𝒈

𝝆𝑪𝑫𝑺ℎ𝑠= 𝟓. 𝟑𝟔𝟏

𝒎

𝒔




•

•

Velocity of CanSat before separation

𝑣1 =2 · 𝑚1 · 𝑔

𝜌 · 𝐶𝐷ℎ𝑠 · 𝜋 · 𝑆ℎ𝑠= 13.2

𝑚

𝑠

Velocity of Probe with deployed

parachute

𝑣2 =2 · 𝑚2 · 𝑔

𝜌 · 𝐶𝐷𝑝 · 𝑆𝑝= 5

𝑚

𝑠

Velocity of Heat Shield after separation

𝑣3 =2 · 𝑚3 · 𝑔

𝜌 · 𝐶𝐷ℎ𝑠′ · 𝜋 · 𝑆ℎ𝑠= 5.3

𝑚

𝑠



Mechanical Subsystem Design



Mechanical Subsystem Overview

Key Aspects & Materials Heat ShieldKey Aspects & Materials

Nylon Heat

Shield

Egg Container

(ABS)

Parachute Bay

(ABS)Nylon Rods

Camera Bay (ABS)

Carbon Fibre

Rods

Nose Cone

(ABS)

248.5 mm

120 mm

Heat Shield

Heat Shield is attached via a solenoid

rod extending through a hole at the top

of the nose cone. Once the solenoid

retracts it’s rod, the heat shield

releases from the probe.

ProbeOnce the heat shield

is released, the

parachute deploys

from its bay. The

parachute is held to

the probe via two

holes in the

parachute bay plate.


Mechanical Subsystem

Changes Since PDR


PDR CDR Rationale Prototype Testing

3D printed material uses

PLA.

3D printed material uses

ABS.

This was due to problems encountered

involving warping and poor strength.

Sparse setting is still used.

PLA wasn’t strong enough and

parts kept snapping on the

prototype.

Diameter of CanSat is 90

mm (excl. nose cone &

heat shield)

Diameter of CanSat is 100

mm (excl. nose cone & heat

shield)

It was realised that more room was

required and there was enough space to

implement it.

N/A

A 3D printed part is used

to cover the egg and it’s

protection material

(sponge).

The dimensions have been

changed by increasing the

height and width to provide

extra room for padding.

There have been reinforcing

struts placed on the interior

surfaces. Holes have been

placed on multiple faces

across the egg cover. They

will be covered with stickers.

Reinforcing struts used to reduce the

effects of warping. Holes have been used

to reduce the weight of the egg cover as it

was the heaviest part. They will be

covered with stickers. This change saved

20 g of weight.

Environmental testing during

the test launch proved the egg

protection system did not work

but this was under extreme

circumstances. Drop tests will

confirm whether the method

works.

120 mm nylon spacers

used for structural

integrity in the main bay.

100 mm carbon fibre rods

and 15 mm spacers at each

end.

This change was implemented to both

reduce weight and increase the structural

integrity of the system. This change saved

20 g of weight. It also increased height.

The spacers survived the

environmental testing during a

prototype launch.

There have been a significant number of changes to the CanSat design. The majority of them have been to reinforce the

structural integrity and to reduce the overall weight. There are only a few design changes - such as the removal of the

solenoid and allowing one servo to complete HS deployment, HS release and parachute release - that have been done to

have smoother operation.



Changes Since PDR



Parachute hatch is flat.

15 mm laser cut ‘feet’ have

been placed at the top of the

CanSat at the parachute

end.

This was implemented so that the

CanSat was not lying on the

bulkhead bolt inside the rocket.

The feet survived the

environmental testing during a

prototype launch.

Torsion springs hold the

carbon fibre rods in place

for the HS.

3D printed pivots hold the

carbon fibre rods in place for

the HS. They pivot about

some arms on the nose

cone.

The torsion springs attachment

method to the CF rods was weak and

had to be improved. ABS rod holders

were designed and held in place to

act as a pivot. They were attached to

the nose cone with extensions

springs to maintain tension so the

natural position was in the deployed

position.

When building the prototype, it

became apparent that the torsion

springs wouldn’t work. The rod

pivots have been tested

independently and during a test

launch and worked effectively.

A servo opens the

parachute bay hatch to

release the parachute.

A spring has been used so

that the natural position is

open. A string holds it

closed. Therefore, when the

string is released, the

parachute bay hatch will

open. Height of the bay has

also increased.

The use of a servo for only parachute

bay hatch opening was undesirable

due to weight and electronics issues.

By using a spring, it reduces weight

and reduces the strain on the

electronics systems.

The Central Control Unit (CCU)

has been tested independently and

during a test launch and worked

effectively.

A 3D printed mounting

plate is used to place the

electronics onto.

A PCB holds all of the

electronics.

This change was employed for

multiple reasons. It not only reduced

the weight of the CanSat, but it also

made it easier to directly mount

electronics.

The PCB has yet to be tested. It

will be tested independently and

during a test launch.



Changes Since PDR



Camera is located in

camera bay but

attached to the

underside of the

payload bay plate.

Camera is located in

camera bay but attached

to the top of the camera

bay plate.

As the choice of camera changed, it was

easier to mount it on the bottom plate rather

than the location seen on the PDR.

The camera has been tested

independently and has worked

effectively. It will have an

environmental test during the next

prototype launch.

A solenoid retracts

from inside the HS

attachment point to

release the HS.

Springs push the nose

cone away once

released.

A servo with a length of

metal rod retracts to

release the HS. The

springs have been

removed.

It was decided that one dedicated mechanism

for HS deployment, HS release and

parachute release was preferable to multiple

mechanisms. This reduced weight, reduced

the complexity of electronics and the flight

software and is easier to implement.

The CCU has been tested



The HS is deployed

through an

independent bay that

has an ‘X’ shaped

device which deploys

the HS when activated

with a solenoid.

A central system that

uses a servo with a

metal rod also controls

the deployment of the

HS, as well as the

release of the HS and

parachute.

It was decided that one dedicated mechanism

for HS deployment, HS release and

parachute release was preferable to multiple

mechanisms. This was decided as it reduces

weight, reduced the complexity of electronics

and the flight software and is easier to

implement.

The CCU has been tested



Sponge is used to

protect the egg

against shock and

vibrations.

A combination of

sponge, cotton balls and

a bag are used to

protect the egg.

During the prototype launch, the egg was

placed inside a bag filled with cotton balls to

provide better padding and to protect the

electronics if the egg is to break.

Environmental testing during the

test launch proved the egg

protection system did not work

but this was under extreme

circumstances. Drop tests will

confirm whether the method

works.



Changes Since PDR



The nose cone has four

hollow pillars to hold four

springs to aid release.

There are four points for

attaching torsion springs.

A central pillar is used for

release.

The hollow pillars have been

dismissed as the springs aren’t

necessary. They are now just

pillars. The torsion springs

have been replaced by 3D

printed rod pivots that pivot

about arms on the nose cone.

The overall shape of the nose

cone has changed as the

circular edges have been

removed for straight lines.

The pillars remain to provide a stable

base for the rest of the CanSat and it also

stops additional stresses on the servo rod

against the nose cone attachment point.

The torsion spring idea proved to be

completely unreliable and unrealistic. It

has been replaced by a more reliable and

smoother method. There have been many

changes to the nose cone to further

reduce the weight of the CanSat. The

circular edges were removed to save

weight and to ease the difficulty of HS

manufacture.

The nose cone has been

tested independently and

during a test launch and

worked effectively.

GPS and buzzer located

on PLA mounting plate.

The GPS and buzzer have

been moved into the parachute

bay and are covered with a

laser cut plywood sheet.

It was realised that the GPS may struggle

to find signal on the side of the CanSat so

it was moved to the parachute bay so it

faces upwards. The buzzer joined as it too

big to be mounted on the PCB due to

space constraints.

The GPS wasn’t available

for the environmental test

but has worked sufficiently

during independent tests.

The buzzer has worked for

both situations.

1 mm carbon fibre plates

used to separate bays.

3 mm plywood is used to

separate the bays.

The plywood is easier to manufacture and

is still relatively lightweight.

The plywood plates have

been tested independently

and during a test launch

and worked effectively.



Mechanical Sub-System

Requirements


A I T D


RE2The aero-braking heat shield shall be used to protect the probe while in the rocket only and when deployed


rocket. The rear end of the probe can be open

X X


RE6The probe with the aero-braking heat shield shall fit in a cylindrical envelope of 125mm diameter x

310mm length. Tolerances are included to facilitate container deployment from the rocket fairing.x

RE7 The probe shall hold a large hen’s egg and protect it from damage from launch until landing. X X

RE8The probe shall accommodate a large hen’s egg with a mass ranging from 54 grams to 68 grams and a

diameter of up to 50mm and a length of up to 70mm. X

RE9The aero-braking heat shield shall not have any sharp edges to cause it to get stuck in the rocket payload

section which is made of cardboard.X





RE14 The aero-braking heat shield shall be released from the probe at 300m. X X

RE15 The probe shall release a parachute at 300m. X X

RE16All descent control device attachment components (aero-braking heat shield and parachute) shall survive

30Gs of shock. X X

RE17 All descent control devices (aero-braking heat shield and parachute) shall survive 30Gs of shock. X X


Mechanical Sub-System

Requirements


A I T D


RE19 All structures shall be built to survive 15Gs of launch acceleration. X X

RE20 All structures shall be built to survive 30Gs of shock. X X

RE21All electronics shall be hard-mounted using proper mounts such as standoffs, screws, or high performance

adhesives.X

RE22 All mechanisms shall be capable of maintaining their configuration or states under all forces. X

RE23 Mechanisms shall not use pyrotechnics or chemicals. X

RE24Mechanisms that use heat (e.g. nichrome wire) shall not be exposed to the outside environment to reduce

potential risk of setting vegetation on fire.X X


RE38 Both the heat shield and probe shall be labelled with team contact information including email address. X

RE40 No lasers allowed. X X

RE41 The probe must include an easily accessible power switch. X X

RE42 The probe must include a power indicator such as an LED or sound generating device. X X

RE47An easily accessible battery compartment must be included allowing batteries to be installed or removed in less

than a minute and not require a total disassembly of the CanSat.X X

B1

Camera: Add a colour video camera to capture the release of the heat shield and the ground during the last 300

meters of descent; the camera must have a resolution of at least 640x480 and a frame rate of at least 30

frames/sec

X X

Payload Mechanical Layout of

Components


1) Launch configuration

2) Deployed configuration

3) Parachute is released

4) Released configuration of probe and heat shield

Red – Parachute Bay

Yellow – Electronics Bay

Light Blue – Egg Container Bay

Dark Blue – Camera/Release/Deployment Bay

1

2

3

4



Components


A

B

D

H

F

J

C

E

G

K

12

13

9

14

15

L

M

N

P

12

3

4

5

6

7

8

10

11

A - Parachute Hatch Hinge (Nylon) J – Parachute Hatch Spring 2 – 118 mm 10 – 115 mm

B – Parachute Stowed (RS Nylon) K – GPS & Buzzer 3 – 54 mm 11 – 78 mm

C – Stickers L - Support Column (CF) 4 – 23.5 mm 12 – 190 mm

D – HS Rod (Carbon Fibre) M – 15 mm Rod (Nylon) 5 – 15 mm 13 – 290 mm

E – On/Off Switch N – Pivot Rod (ABS) 6 – 30 mm 14 – 222.1 mm

F – Battery P – Nose Cone Springs 7 – 130 mm 15 – 248.5 mm

G – Camera Q – Egg 8 – 30 mm

H - Servo 1 – 174 mm 9 – 100 mm

Q



Components


Rocket Body

Tube

Heat Shield

(Ripstop Nylon)

Stowed

Parachute Stowed

(Ripstop Nylon)Parachute

Bay Feet

(Plywood)

Nose Cone

(ABS)

Rod

Pivots

(ABS)

CCU (ABS)

Servo Head Extension (ABS)

Cotton Balls

Egg

Sponge


HS Attachment

(Details Slide 63)

Egg Protection Mechanical Layout

of Components


The egg protection structure design has not changed much from the PDR. However, in order to save

weight, material has been removed at some points. The method for mounting the egg cover onto the egg

protection base has remained the same. The dimensions have also changed, as seen below, to provide

more cushioning for the egg during descent.

A

B

D

C

Figure 3 shows the change in dimensions of

the egg cover, such that (A) = 106 mm, (B) =

70 mm, (C) = 70 mm, (D) = 50 mm. This has

assumed that the maximum egg side will be

used. This gives padding of 10 mm either size

of the egg and 18 mm above and below it. It

was decided more cushioning was required.

Figure 2 shows the weight reductions made to the egg cover. They saved a

mass of 20 g. In order to adhere to the requirements, the exposed areas will

be covered with stickers, which can be seen in Figure 1.

The diagonal trusses are included to maintain structural integrity of the

structure. This is completed to ensure that the egg protection structure will

not fail under extreme conditions.

Figure 1 Figure 2

Figure 3


Egg Protection Mechanical Layout

of Components


Figure 1 Figure 2Figure 3

Figure 4 Figure 5

1) Nut and bolts are undone to allow free movement of the cover hatch. A lip is

used to secure the hatch at the bottom, as depicted in Figure 4.

2) The cover hatch is removed by pulling out and lifting up at the same time. The

egg is now exposed. The white bit simulates the cotton surrounding the egg

inside the bag.

3) It is now possible to remove the egg from its bed.

Figure 4 is included in to show the lip method used to restrict the bottom of the hatch.

Figure 5 is included to show how the egg is fully restricted to avoid damage.



Heat shield Release Mechanism

• The Heat Shield is locked in position by a metal rod that

passes through HS attachment point; the position of the

rod is controlled by a servo; the rod and servo are part

of the Probe.

• At 300 m, the servo receives a signal to fully retract the

rod which results in Heat Shield being released

• One servo controls three processes: Heat Shield

deployment, parachute deployment and Heat Shield

release; this design is simple and saves weight.

• After prototype testing it was decided that compression

springs are no longer needed for the Heat Shield to

release promptly; the drag created by parachute

deployment will be sufficient.

Figure 1: Shows the Central Control Unit (CCU) is located in

the camera bay. This figure is shown for context for location.

Figure 2: The camera and a nylon spacer have been made

transparent to show the metal rod holding the HS attachment

point in place. The HS attachment point has been circled in

red.

Figure 3: As the servo rotates about its axis, it drives the rod

through one degree of freedom. It then release the HS which

can be seen at the bottom of Figure 3.

Figure 1

Figure 2

Figure 3


Probe Parachute Release

Mechanism


• The Parachute is being stowed in a parachute bay

which is kept closed by a string until the CanSat

reaches 300 m.

• A plate in the parachute bay has two designated

holes used to attach the Parachute to the Probe.

• The string is attached to a washer through which

passes servo-rod mechanism mentioned in the

previous slide.

• At 300m servo receives the signal to retract the rod

and the string is released. This results in the

parachute bay being opened and the Parachute

deploys.

Figure 1

Figure 2

Figure 3

Figure 1: The CCU holds the parachute bay line with a

washer.

Figure 2: As the servo rotates and retracts the metal rod,

the washer is released and tension is lost.

Figure 3: The other end of the line hold the parachute bay

hatch closed when the line is in tension. The natural position

for the hatch is open due to the tension on the hinge from

the spring.

Probe Parachute Release

Mechanism


Figure 4

Figure 6

Figure 5

Figure 4: The parachute is pushed down onto a piece of

cardboard that covers a spring that is glued to the probe.

This figure shows the parachute stowed. It is attached to the

bulkhead of the parachute bay. The parachute has been

folded so that it can unravel successfully.

Figure 5: As the door is opened, the cardboard is free to

push the parachute out. This is done to ensure release.

Figure 6: The parachute then releases and slows the

descent rate of the CanSat.

Structure Survivability


Criteria PDR CDR Rationale

Electronics

component

mounting

methods

The electronics are screwed or

glued onto the 3D printed PLA

mounting plate.

The electronics will now be

mounted onto a PCB that has

replaced the mounting plate.

However, the GPS and the

buzzer are now located in the

parachute bay and will be

secured with strong double sided

tape.

The PCB will provide a solid method for

mounting the electronics and also

reduces the likelihood of short circuits. It

is also quicker to mount the electronics

onto. In our test launch, the double sided

tape survived very well so it was

considered adequately reliable.

Electronic

component

enclosures

The main electronics bay is

covered using a thin sheet of

hard plastic. The camera, servo

and two solenoids are covered

within a 3D printed bay.

The main electronics bay is

covered using a thin sheet of

hard plastic. The camera, servo,

GPS, buzzer and switch are all

covered within a 3D printed bay.

Some of the parts were moved for other

reasons and some parts have been

dismissed/replaced. It is regarded as

adequate by the team.

Accelerations and

shock force

requirements and

testing

N/A

Most of the parts being used

have been tested during a

prototype launch and have

survived a fall from ~200 m

without parachute due to a

failure.

The remaining components that haven’t

been tested for shock force requirements

and testing will be tested over the next

two prototype launches and dedicated

systems tests.

Structure Survivability


Criteria PDR CDR Rationale

Securing electrical

connections

Electrical connections will be

secured through tape and

soldering.

As a PCB is now being used,

most of the components won’t

need electrical connections

being secured. The servo

used for the CCU is being

screwed into a bulkhead.

Where necessary, connections

will be secured further with

epoxy.

As the majority of components are

mounted onto a PCB, there is no worry

about components becoming

disconnected. For parts that are liable to

it, epoxy will be more than sufficient as it

is a very strong bond.

Descent control

attachments

A servo and a solenoid were

the only methods for fastening

descent control attachments.

Only a servo is the electronical

components that is used to

connect attachments.

Tensioned string is used to

hold some parts in place.

Springs hold the rod pivots on

the HS and a spring pulls the

parachute bay hatch open.

The servo mechanism (CCU) has been

tested numerous times and has proved to

be reliable. However, on some occasions

it has proven to struggle with the various

loads placed upon it. To improve it, the

team will grease the rod and look at

alternative methods to ease the forces

placed upon the servo. The string

methods for HS deployment and

parachute release have been tested

numerous times also. These tests have

been proven the method to be reliable. As

long as the springs don’t yield, they can

be considered a sound method.


Mass Budget

Mass of all Electronic Components

Subsystem Component Mass (g) Justification

CDH/SE/FSW Teensy 3.2 5.4 Measured

CDH RTC 3 Measured

CDH Xbee 6.6 Measured

CDH Buzzer 3.3 Measured

SE 10DOF IMU 3.3 Measured

CDH SD Card & Breakout 2.6 Measured

SE Camera 5 Measured

EPS Battery 33 Measured

SE Adafruit Ultimate GPS 9.1 Measured

ALL PCB 30 Estimated

DCS/ME Servo 9 Measured

EPS On/Off Switch 5 Measured

Total 115.3 g



Mass Budget

Mass of Structural Elements (Probe 1/3)


M/DC

Egg Containment

Egg 68 Datasheet

Sponge 5.8 Measured

Cotton Balls 2 Measured

Egg Bag 4 Measured

15 mm Spacers (x6) 11.2 Measured

30 mm Spacers (x3) 9 Measured

100 mm Carbon Fibre Spacers (x3) 8 Measured

Parachute Bay Door

Release

Parachute 10 Datasheet

Spring 0.7 Measured

String 0.5 Measured

Hinge 1.8 Measured

Electronics Cover 8 Measured

SG90 Servo Head Extension 0.6 Measured

Stickers (x7) 4 Estimated

Washer 0.2 Measured



Mass Budget



M/DC

3 mm Plywood Plates

Payload Bay Plate 8.1 Measured

Camera Bay 9 Measured

Parachute Bay 7.5 Measured

Parachute Bay Hatch 6.1 Measured

Parachute Bay

Electronics Cover4 Measured

Parachute Pusher 3.6 Measured

Parachute Bay Feet (x4) 1.9 Measured

3D Printed Parts (ABS

on ‘sparse’ setting of

70%)

Parachute Bay 24.64 Measured

Egg Container Cover 28.14 Measured

Egg Container Base 7.98 Measured

Camera Bay pt. 1 6.93 Measured

Camera Bay pt. 2 8.68 Measured

CCU Base 9.24 Measured

CCU Pin Head 0.42 Measured

CCU Servo Head

Extension0.35 Measured



Mass Budget



M/DC

M3 Bolts (x15) 8.4 Measured

M3 Nuts (x6) 1.8 Measured

M1.6 Bolts (x7) 1.8 Measured

M1.6 Nuts (x4) 0.4 Measured

Glue 3 Estimated

Total 275.8 g



Mass Budget

Mass of Structural Elements (Heat Shield)


ME/DCS

Nose Cone (ABS - 100% Fill) 49.9 Measured

Rip stock Nylon 12.1 Measured

Rod Pivots (x4) 9 Measured

Carbon Fibre Rods (x4) 8.4 Measured

M3 Bolts (x4) 3.0 Measured

Springs (x4) 2.8 Measured

Pivot M3 Bolts (x4) 4.9 Measured

Zip Ties (x4) 2 Measured

Glue 3 Measured

Cotton (for sewing) 1 Measured

Total 96.1 g



Mass Budget

Mass Margins

Heat Shield 5% 4.8 g

Probe – Structure 5% 13.8 g

Probe – Electronics 5% 5.8 g

Total Margin 15% 24.4 g

System Total

Margins 24.4 g

Probe (Structure + Electronics) 391.1 g

Heat Shield 88.9 g

Total Inc. Margins 504.4 g

Total Exc. Margins 480 g

Currently, the mass is within/under the mass budget

requirement. Once a second prototype has been fully

assembled after recent changes made, the team will asses

if more weight needs to be added.

More weight can be added in many forms. One key

method is increasing the size of the egg cover and giving

the egg more padding. Not only will this provide better

protection, but the egg cover is the second heaviest part

on the CanSat.

Another method for adding weight is to change the 3D

printed parts to 100% infill rather than 70%. This is a

simple, yet effective method.

In case the egg is smaller than expected (as the largest

egg has been assumed), then ballast will have to be

added. Ballast can be added underneath the protection

sponge for the egg. This will be planar in nature to reduce

the effect of reducing the amount of protection for the egg.

The table to the left shows the overall mass margins for

the entire system.

When considering requirement 1, the prototype satisfies it

when margins are included. However, when excluded, the

CanSat is underweight but this can be resolved as

mentioned above.


TOTAL BUDGET


Communication and Data Handling

(CDH) Subsystem Design



CDH Overview

Probe System Overview

Teensy

3.2 A

Temp.

Tilt

GPS

Air

Pressure

Camera

SD Card A

XBee Pro

S2C

SD Card B

Taoglas

Patch

Antenna

GCS

BAT+

SPI

3.3 V

Digital

Pin

SPI

Serial Serial

3.3V

3.3VI2C

I2C

I2C

CDH Component Overview

Component Function

Teensy 3.2 Probe Microprocessor

XBee Pro S2C Probe Radio

DS1338 RTC RTC for the system

BAT+


CDH Changes Since PDR

•


PDR CDR RATIONALE

Using two microcontrollers:

Teensy 3.2 and Arduino Nano

Using only single

microcontroller

(Teensy 3.2)

• Selected Camera has onboard SD card

interface hence Arduino Nano not

required

• Camera and Servo can be powered by

5V from voltage regulator

• Camera and Servo can be controlled

using Teensy 3.3V digital pin and PWM

pin respectively

RTC: Adafruit DS1307 RTC: Adafruit DS1338 • 3.3V compliant

SD Card breakout: 2x 5V SD

breakout

SD Card breakout: 1x

3.3V SD breakout

• 3.3V compliance

• Change from two microcontrollers to

one and new camera has in-built SD

card storage



CDH Requirements


A I T D

RE1 Total mass of the CanSat (probe) shall be 500 grams +/- 10 grams. X X

RE6 The probe with the aero-braking heat shield shall fit in a cylindrical envelope of 125 mm diameter x 310 mm

length. Tolerances are to be included to facilitate container deployment from the rocket fairing.X X



adhesives.X

RE25 During descent, the probe shall collect air pressure, outside air temperature, GPS position and battery voltage

once per second and time tag the data with mission time.X X

RE26 During descent, the probe shall transmit all telemetry. Telemetry can be transmitted continuously or in bursts. X

RE27 Telemetry shall include mission time with one second or better resolution. Mission time shall be maintained in

the event of a processor reset during the launch and mission.X

RE28 XBEE radios shall be used for telemetry. 2.4 GHz Series 1 and 2 radios are allowed. 900 MHz XBEE Pro

radios are also allowed.X

RE29 XBEE radios shall have their NETID/PANID set to their team number. X X

RE30 XBEE radios shall not use broadcast mode. X X


RE42 The probe must include a power indicator such as an LED or sound generating device. X X

RE45 An audio beacon is required for the probe. It may be powered after landing or operate continuously. X X

RE46 Battery source may be alkaline, Ni-Cad, Ni-MH or Lithium. Lithium polymer batteries are not allowed. Lithium

cells must be manufactured with a metal package similar to 18650 cells.X

RE49 A tilt sensor shall be used to verify the stability of the probe during descent with the heat shield deployed and

be part of the telemetry.X

B2 A radio transmitter shall be added to transmit the wind speed by changing its frequency. The frequency change

shall be 1 Hz per 0.1 meter/sec. The transmitter must be custom designed and built. It cannot be a commercial

product. The frequency must be in the 433 MHz ISM band, or if a team member has an amateur radio license,

an amateur radio band can be used. The transmitter must be able to set to 8 different frequencies in the 433

MHz ISM band with 25 kHz separation. The transmitter must turn off after the probe lands to minimize

interference.

X X X X


Probe Processor & Memory

Selection

DEVICE CHOSEN RATIONALE

Teensy 3.2 • Low mass

• Exceptional memory capability

• Easy to programme with Arduino IDE

Processor

SpeedCost Weight Data Interfaces

Non-volatile Memory

Options

Volatile Memory

Options

72 MHz £19.80 4.8 g

USB 1

EEPROM

(2KB)

Flash

(256 KB)SRAM (64 KB)

Serial 3

SPI 1

I2C 2

Memory storage requirements:

• RTC value and packets transmitted will be saved in microcontroller non-volatile EEPROM so that in the

event of processor reset, mission time will still be known and telemetry can resume coherently

• Store telemetry data in non-volatile memory (SD Card A – 32GB sufficient for maximum expected data

volume)

• Videos taken during descent are stored in SD Card B (on-board camera SD Card B – 32GB sufficient for

maximum expected data volume)

Arduino Nano is no longer needed because new camera selected has on-board SD card and can be

powered by a single 3.3 V digital output pin from the Teensy.


Probe Real-Time Clock



DS1307 • Light weight

• Affordable

• 2 sec/day drift is reasonable

Weight Size Cost Power Accuracy/Error @ 25 °C Interface

2.3 g 26x22x5 mm ~₤2Coin Cell Battery ~23 ppm

2 sec/dayI2C

Duracell LR44

Duracell LR44 specifications: 9g, 1.5 V, 105 mA-hr

Payload Real-Time Clock: DS1307

• Hardware clock

• On-board microcontroller time will be synced with the RTC readings to give an accurate <Mission

Time> value for telemetry

• 2 sec/day accuracy

• Reset tolerance:

▪ RTC readings will be saved in microcontroller EEPROM so that in the event of processor

reset, mission time will still be known and telemetry can resume accurately



Probe Antenna Selection


FXP70 Freedom Multi

Standard Antenna

• Higher gain

• Low profile

Gain VSWR Mass Size Polarization

5 dBi ≤ 1.5:1 1.2 g 27 x 25 x 0.8 mm Horizontal, vertical

Antenna Considerations

Antenna is 2. 4 GHz because this is the XBee Pro S2C’s frequency.

Antenna’s “face” will be mounted flush to “bottom” (downward facing during flight) of CanSat.



Probe Radio Configuration

XBee Considerations

• NETID will be team No. 5002

• Broadcasting mode will not be used to transmit data

• Transmission will be handled by code in the µ-Controller

• No transmissions during launch.

• Transmissions are made at a rate of 1Hz during the descent phase of the mission.

• Transmissions cease after CanSat has landed.


XBee Pro S2C • High sensitivity

• Team member experience with device

• Tx supply current is manageable by Teensy 3.2

Gain VSWR Mass Size Polarization

5 dBi ≤ 1.5:1 1.2 g 27 x 25 x 0.8 mm Horizontal, vertical



Probe Telemetry Format

Data will be transmitted at a rate of 1 Hz in bursts.

The telemetry data file will be named:

CANSAT2018_TLM_<TEAM ID>_<TEAM_NAME>.csv

Telemetry data shall be transmitted with ASCII comma delimited fields followed by a carriage

return in the following format:

<TEAM ID>,<MISSION TIME>,<PACKET COUNT>,<ALTITUDE>, <PRESSURE>,

<TEMP>,<VOLTAGE>,<GPS TIME>,<GPS LATITUDE>,<GPS LONGITUDE>,<GPS

ALTITUDE>,<GPS SATS>,<TILT X>,<TILT Y>,<TILT Z>,<SOFTWARE STATE>,<BONUS>

Example toy packet:5002,100,100,600.0,101000.0,20.0,8.7,100,32.2,-43.2,600.0,7,45.0,45.0,45.0,DEPLOYED



Probe Telemetry Format

Quantity Description Variable Type Maximum Size in Bytes

<TEAM ID> Four digit team identification number Integer 2

<MISSION TIME> The time since glider power up [sec] Integer 2

<PACKET COUNT> Count of transmitted packets Integer 2

<ALT SENSOR> Altitude with one meter resolution [m] Float 4

<PRESSURE> Measurement of atmospheric pressure [Pa] Float 4

<TEMP> Sensed temperature with one degree resolution [°C] Float 4

<VOLTAGE> Voltage of the CanSat power bus [V] Float 4

<GPS TIME> Time generated by the GPS receiver Integer 2

<GPS LATITUDE> Latitude generated by GPS receiver Float 4

<GPS LONGITUDE> Longitude generated by GPS receiver Float 4

<GPS ALTITUDE> Altitude generated by GPS receiver Float 4

<GPS SATS> # of GPS satellites being tracked by GPS receiver Integer 2

<TILT X> Tilt sensor X axis value. Float 4

<TILT Y> Tilt sensor Y axis value. Float 4

<TILT Z> Tilt sensor Z axis value. Float 4

<SOFTWARE STATE> Current operating state of the software String 18

<BONUS> N/A (Videos stored on camera SD card) N/A N/A



Electrical Power Subsystem Design

Presenter Name(s) Go Here


EPS Overview

Component Function

Switch Manual Power On/Off

RTC Battery CR 1225 Battery for time keeping

GPS Battery CR 1220 Battery for time and location

Energize L522 Ultimate Lithium Main Battery

LM7805 Voltage Regulator

Teensy 3.2 MCU

Adafruit 10 DOF IMU Press, Temp, Alt, Tilt

Modified SQ11 Camera Video Bonus

Servo 9g Release Mechanisms

XBEE S2C Pro Transceiver

SD Card Breakout Board Data Logging Onboard

Adafruit Ultimate GPS v3.0 GPS

Audio Beacon Audio Beacon

EPS Overview:

The CanSat probe is now powered by a single 9V

Energizer L522 Ultimate Lithium Battery.

The probe now contains only one Microcontroller, the

Teensy 3.2.

The Teensy is powered from a 5V regulator along with

all other electronics, except the SD Card and the

XBee, which are powered from a 3V3 output pin of the

Teensy.



EPS Overview

VOLTAGE REGULATOR

AUDIO BEACON

SD CARD

SERVO

TEENSY 3.2.

10 DOF IMU

GPS

CAMERA

XBEE

RTC

SWITCHVOLTAGE

DIVIDER


EPS Changes Since PDR


PDR CDR RATIONALE

2 MCUs 1 Teensy 3.2. Less weight

2 SD Card Breakout as standalones 1 SD Card Breakout as standalone

1 SD Card Breakout inbuilt on the

Camera

Camera comes with inbuilt SD

Card Slot

3 Release Mechanisms 1 Release Mechanism Significantly less weight

Adafruit Serial JPEG Camera Modified SQ11 Pawaca Camera Better quality video

Better weight

Components attached by various

methods onto PLA plate

PCB Reliability

Ease of removal of components

(modularity)

Duracell 9V Alkaline Energizer 9V Ultimate Lithium Improved power capacity.

Less weight.



EPS Requirements

RE# Description Verification

A I T D


RE6 The probe with the aero-braking heat shield shall fit in a cylindrical envelope of 125 mm diameter x 310

mm length. Tolerances are to be included to facilitate container deployment from the rocket fairing.X

RE18 All electronic components shall be enclosed and shielded from the environment with the exception of

sensors.X


adhesives.X

RE25 During descent, the probe shall collect air pressure, outside air temperature, GPS position and battery

voltage once per second and time tag the data with mission time.X X


RE41 The probe must include an easily accessible power switch. X

RE45 An audio beacon is required for the probe. It may be powered after landing or operate continuously. X X X

RE46 Battery source may be alkaline, Ni-Cad, Ni-MH or Lithium. Lithium polymer batteries are not allowed.

Lithium cells must be manufactured with a metal package similar to 18650 cells.X X X

RE47 An easily accessible battery compartment must be included allowing batteries to be installed or removed in

less than a minute and not require a total disassembly of the CanSat.X X X

RE48 Spring contacts shall not be used for making electrical connections to batteries. Shock forces can cause

momentary disconnects.X

Bonus Camera: Add a colour video camera to capture the release of the heat shield and the ground during the

last 300 meters of descent. The camera must have a resolution of at least 640x480 and a frame rate of at

least 30 frames/sec.

X X X X

Bonus Wind Sensor and Radio Transmitter X X X X


Probe Electrical Block Diagram

NOTES

The GPS, Adafruit and RTC

are 5V compatible. The Servo

requires 5V power supply. The

SD card and XBee are powered

by the Teensy as they need

3.3V.

The audio beacon will be used

to indicate that the CanSat is

powered and operational.

The switch is an easily

accessible external switch. The

Teensy 3.2 can be accessed

through an umbilical cord

without disassembling the

CanSat.

No spring contacts will be

used for making electrical

contacts to batteries. A PCB will

be used for the circuit

connection.


Probe Power Source


Battery selection: Energizer L522 Ultimate Lithium

• This battery was chosen as it has a higher mAh rating than the previous one (Duracell 9V

Alkaline)

NOTE: The chosen battery source is Lithium. The cell is indeed manufactured with a metal package

similar to 18650 cells. The chosen battery is not easy to damage and does not represent a fire hazard.

Voltages: 9V power supply, and using a voltage regulator to get 5V for the circuit. A 3.3V is

supplied to specific components through the Teensy

Current capacity (Tested in Real Life):

• 230 mA without camera.

• 363 mA with camera (just last one minute of the flight).

How much current battery can generate: 1000 mA (Datasheet)



Probe Power Budget


Component NameCurrent

[Amp]

Voltage

[Volt]

Operational

Power [W]

Duty Cycle

[%]

Duty Cycle

[Hrs]

Duty hour

[sec]

Required

Capacity

[W-hr]

Required

[A-hr]

Adafruit 10DOF IMU 0.001 5 0.005 12% 00:14:24 864 0.0012 0.00024

Adafruit Ultimate GPS

Breakout 0.02 5 0.1 12% 00:14:24 864 0.024 0.0048

Zigbee/802.15.4

Modules 0.12 3.3 0.396 12% 00:14:24 864 0.09504 0.0288

Modified SQ11

Pawaca Camera 0.075 5 0.375 1% 00:01:00 60 0.00625 0.00125

Teensy 3.2 USB 0.0003 5 0.0015 100% 02:00:00 7200 0.003 0.0006

Servo 9g 0.25 5 1.25 100% 02:00:00 7200 2.5 0.5

SD Breakout 0.1 5 0.5 12% 00:14:24 864 0.12 0.024

Audio Beacon 0.035 3.3 0.1155 12% 00:14:24 864 0.02772 0.0084

TOTAL 0.56809

All data is from component data sheet.

Power source (Energizer 9V Ultimate Lithium) offers at least 0.75 A-hr (Datasheet). (Margin of 40%)

In practice, the true Duty Cycle will not be this long.


Flight Software (FSW) Design



FSW Overview

•

•The architecture of the program relies on Teensy 3.2.

C programming language is used as it

offers more in depth programming flexibility compare to

higher level languages (e.g. C#)

C is the most common language among

team’s programmers so it makes sense to use

the existing skillset available

The IDE used is ‘Arduino’- a very simple and easy to

understand IDE which should provide all the functionality that

we need (simpler to get everyone involved in the development)

Tasks of the software:

• Calibrate

• Ensure everything runs smoothly (running checks)

• Power

• Sensor failures

• Critical mission points

• Handle (process) data

• Store system data to EEPROM – ensures state recovery in

caseof sudden power loss

Altitude = 300m (On descent)

Heat Shield Released

Parachute Deploys

Ground

Impact

Audio

Beacon

Activates

Launch

C/C Sensing and Transmitting

Apogee reached

Payload detached

HS engaged


FSW Changes Since PDR


PDR CDR Rationale

Arduino Nano

&

Teensy 3.2

Teensy 3.2 Weight savings

Camera has built-in SD card

New features

• EEPROM backup

• Force Sampling

New way of testing

• Simulation



FSW Requirements

•

•

•

•

•

ID RequirementVerification

A I T D


RE14 The aero-braking heat shield shall be released from the probe at 300 meters X

RE15 The probe shall deploy a parachute at 300 meters. X

RE25During descent, the probe shall collect air pressure, outside air temperature, GPS position and battery

voltage once per second and time tag the data with mission time. X X

RE27Telemetry shall include mission time with one second or better resolution. Mission time shall be

maintained in the event of a processor reset during the launch and mission.X X

RE31Cost of the CanSat shall be under $1000. Ground support and analysis tools are not included in the

cost. X

RE39The flight software shall maintain a count of packets transmitted, which shall increment with each

packet transmission throughout the mission. The value shall be maintained through processor resets. X X

RE42 The probe must include a power indicator such as an LED or sound generating device. X X X

RE45 An audio beacon is required for the probe. It may be powered after landing or operate continuously. X X

RE49A tilt sensor shall be used to verify the stability of the probe during descent with the heat shield

deployed and be part of the telemetry. X X

BONUS

Add a color video camera to capture the release of the heat shield and the ground during the last 300

meters of descent. The camera must have a resolution of at least 640x480 and a frame rate of at least

30 frames/sec. The camera must be activated at 300 meters.

X X X

BONUS

A radio transmitter shall be added to transmit the wind speed by changing its 10 frequency. The

frequency change shall be 1 Hz per 0.1 meter/sec. The transmitter must turn off after the probe lands

to minimize interference.

X X X


Probe CanSat FSW State Diagram

Info:

Mission State:

• Not Deployed

• Deployed

• HS Released

• Landed

[C]

represents the limit of

possible retries in case of

negative results for a check

System Recovery:

EEPROM memory will be

read in order to recover

the state(settings) of the

software in case of

sudden processor

resets.

Payload

Switched On

Turn on all systems

All systems operational?

(sensor & radio) [C]

No

Yes

Calibrate Functional Sensors

Take Sensor Measurements

Is altitude > 350m?

Yes

No


Is altitude <= 350m?

No

[C=0]


[C]

Is altitude <= 301m?

No

No [C=0]

Deploy

Heat Shield

+

Store exact

deployment

data

Take Sensor

Measurements

Is altitude <

10m?

Activate Audio

Beacon

& Stop

measurements Yes

No

Yes

Yes

Handle Packet

(ensure 1s interval)

Handle Packet

(1s intervals)

Is time since the last

transmission < 1s

Yes

Kill Time (1-time since last

transmission)

Transmit and

store data

NoIn Out

Is payload released

[C]

No

[C=0]

Handle Packet


No

Engage shield

(Power Actuator)

Yes


Handle Packet




Software Development Plan

Development sequence (chart given below):

1. Test each component to identify and address any challenges with thatcomponent

2. Integration testing to identify and address any challenges which may occur in regards tocomponents

compatibility with othercomponents

3. Weekly development sessions of the FSW

The development sequence is a part of the project plan. It will be finished as and when manufactureand build finishes in

order to permit system level testing(simulation).

FSW Development Iteration: (Plan > Implement > Test)

Component

Testing

Integration

Testing

FSW

Development: Iterations

Simulation

DONE

ONGOING

We’ve already started an exhaustive system testing in a simulative environment (created by the system itself using dummy

variables), which we will continue until the beginning of the mission.

The results are looking great at the moment. However, we will keep a defensive programming style if any uncovered cases

are to be found.




Software development team (RS, IA, LAF, NZ)

We are using GitHub (web-based Git repository manager) to store and manage our source-code.

Every FS team-member was instructed on how to use Git commands. This allowed us to keep track of all the changes

made in the code and of course we have the possibility to return to previous versions of it if something goes wrong.

Benefits:

▪ Keep track of all the code changes (recover previous versions if needed)

▪ “Issue Tracker” - a tool to create and track issues/development steps, which has useful functionalities like: developer

assignation, deadlines

The team conducts weekly meetings to discuss planning(creating new issues and setting new deadlines), rather than

presenting individual progress which is already done in GitHub by giving commits(modifications) descriptions and by

commenting/closing issues.

This allows us to focus more on planning, by saving a lot of time with the functionality provided by GitHub.




FS Strategy and Progress (Keyword: FAILPROOF)

We planned on creating a software (completely free of bugs) with a low dependency between components, such that at the

end of the mission we will be able to tell exactly which components failed physically.

And we are confident to say that we did it!

Key Features:

• Force Sampling (we are able to collect at least ten altitude samples per second, hence we should be able to release the

heat shield and deploy the parachute at an altitude very close to 300 meters )

• Every component is independent of all the other components, besides the processor (enhances analysis)

• Retrieving data:

➢ Main: radio transmission

➢ First backup: SD card

➢ Second backup: EEPROM (at state change packets will be stored in the microcontroller’s non-volatile memory)

This approach (reduced component dependency & multiple backups) will help us a great deal with the mission analysis by

ensuring a continuous flow of information (as long as the microcontroller works) in order to answer a lot of the ”WHY?”

questions which will be raised at the end of the mission

A well structured, abstract documentation(explanation with references to the actual code) will be written in order to

demonstrate that the software is fail-proof. This information will be used for proving our mission analysis.



Ground Control System (GCS) Design




GCS Overview

Probe XBee

GCS XBee

SMA to RP-SMA Adapter

Laptop (GUI)

2.4 GHz Yagi


GCS Changes Since PDR


NO CHANGES MADE TO GCS GROUND STATION



GCS Requirements

ID RequirementVerification

A I T D

RE26 During descent, the probe shall transmit all telemetry. Telemetry can be transmitted continuously or

in bursts. X X

RE28 XBEE radios shall be used for telemetry. 2.4 GHz Series 1 and 2 radios are allowed. 900 MHz

XBEE Pro radios are also allowed. X

RE30 XBEE radios shall not use broadcast mode. X X X

RE31 Cost of the CanSat shall be under $1000. Ground support and analysis tools are not included in the

cost. X X

RE32 Each team shall develop their own ground station. X

RE33 All telemetry shall be displayed in real time during descent. X X

RE34 All telemetry shall be displayed in engineering units (meters, meters/sec, Celsius, etc.) X X

RE35 Teams shall plot each telemetry data field in real time during flight X X

RE36 The ground station shall include one laptop computer with a minimum of two hours of battery

operation, XBEE radio and a hand held antenna. X

RE37 The ground station must be portable so the team can be positioned at the ground station operation

site along the flight line. AC power will not be available at the ground station operation site. X

GCS Design

104

2.4 GHz Handheld Yagi

XBee S2C

GCS LaptopParallax XBee USB Adapter

Board

Mini USB to

USB 2 CableSMA to RP-

SMA Adapter

Specifications

Battery 4 hours (from fully charged)

Overheating MitigationLaptop Cooling Pad

Sun-shielding umbrella

Auto-update MitigationDisable auto update feature

Disable Internet connection

CanSat 2018 CDR: Team 5002Presenter: Lawrence Allegranza France


GCS Antenna


Yagi Antenna

Model: TY-24-17-20

• Higher dBi than other Yagi with comparable beam width

• Larger beam width than grid antenna

Gain Horizontal/Vertical Beamwidth Connector Polarization

17 dBi 25° / 24° N Female Horizontal, vertical

Distance Link Prediction and Margins

Range will need to be tested further under more controlled

conditions.

Most recent test resulted in a consistent signal received at

approximately 500 m. Link was tested in a metropolitan

area and through a window pane, which are two factors

that need to be eliminated/made negligible in the next test.

Margins must still be tested.

Predicted Range (under ideal conditions): 1 km

Predicted Margin: Yagi pointing can be off by

approximately 10° from the ideal pointing orientation in

either horizontal or vertical direction.



GCS Antenna

Antenna PortabilityAntenna will be handheld.

It will be mounted on lightweight PVC

pipe

Antenna Construction (Assembly)To ensure that connection between Yagi

and GCS XBee is as short as possible (to

reduce effect of noise), the XBee will be

mounted on the Yagi/boom, and

connected to the GCS Laptop by a long

USB cable.


GCS Software


The team has developed their own Ground Control Station.

The GCS code was finished at the PDR – no progress.

COTS Software packages:

Python 2.7 – Computational Environment of choice.

Anaconda Python Package– encompasses real time plotting and data manipulation utilities for Python.

XBEE Python Library – encompasses real time access to XBEE through USB interface.

SKLearn Python Library – simple data filtering and data post-processing utilities

Command Software and interface:

No commands are planned to be incorporated, as the whole operation will be automated.

However, commands can be sent from the Ground Control Station to the CanSat at the push of a button.

The GCS Script makes use of the XBEE Python Library to access the XBEE receiver through its USB interface, in order to collect

data in real time.

Telemetry Data Recording:

Data (temperature, pressure, etc.) will be recorded in a .csv file right after being read through the USB interface,

without any processing.

Data from this .csv file will later be processed in Python or MS Excel to show at the PFR.

During flight, the data (temperature, pressure, etc.) is then processed, checked and plotted in their respective plot windows.

.csv file creation:

.csv file creation is a relatively simple and straight forward task. The .csv file is created during the setup of

the GCS Python script, and data is continuously appended to the file, as it arrives in packets to the GCS.


GCS Software


GCS Code Architecture.

Top is flowchart.

Right is implementation.


GCS Software


LEFT:

GCS tested in real life with Yagi antenna.

RIGHT:

GCS with random data coming through USB cable.

Screenshots of GUI as seen on Laptop GCS screen.


GCS Bonus Wind Sensor

110CanSat 2018 CDR: Team 5002Presenter: Lawrence Allegranza France

ISM Band

Receiver

ISM Band Yagi Antenna

USBMicro

processor

• ISM Band Yagi Antenna (additional to 2.4 GHz Yagi)

• Designed ISM receiver connected to Yagi to pick up ISM Band signal

• Frequency down conversion

• Tone (frequency) is decoded.

• Data is received and processed the same way as other telemetry, with

no difference except the USB Port address.

• Wind speed will be plotted (magnitude only, as direction not required)

• Frequency down conversion done within ISM Band receiver

Tone

Decoder


CanSat Integration and Test





Overview

Legend

Subsystem level testing plan

Integrated level functional testing plan

Environmental testing plan

Systems Level

Subsystems Level

Mission

Launch VehicleGround Control Station

Sounding Rocket

CanSat

▪ Sensors

▪ CDH

▪ EPS

▪ FSW

▪ Mechanical

▪ Descent Control

▪ Antenna and XBee

▪ GUI/Display

Mission

Test Procedure

Experimental E

Simulation S

Verification V

Mission integration and testing overview:




Overview

Probe

DCS

FSW

CDH

EPS

SE

Grouped

Electronic

Subsystems

Mech

Probe

Subsystem integration plan:

Antenna

+

Xbee

GCS

Software




Overview

Subsystem tests identified:

These correspond to the test procedures on “Test procedures description” slides.

Test description

SENSORS

Verify all sensors are correctly calibrated and configured for their respective functions

Verify each sensor's range, resolution, sampling rate is as required to provide valid data at a

rate of at least 1Hz

Verify collected data is in suitable format for FSW

Verify tilt sensor readings can be used to exhibit probe stability during descent with

heatshield deployed

BONUS – Verify camera can produce colour footage at 640x480 at least 30fps

EPS

Verify power source provides required power level (current and voltage)

Verify voltage divider provides required voltage to all components

COMMUNICATIONS: CDH

Verify radio can be configured to transmit as required: burst or continuous transmission, API

mode (NOT broadcast mode), correct PAN ID, correct team ID, correct packet

Verify real time clock can retain data following system power loss

Verify SD card can store the maximum expected volume of data (maximum telemetry

packet size multiplied by maximum number of seconds)

Verify full telemetry packet is transmitted correctly by the radio module at a rate of 1Hz

Verify microcontroller regains correct function and retains mission data (correct mission

time using RTC) following power loss

COMMUNICATIONS: GCS

Verify GCS computer is suitable for competition final: portable, battery life is suitable for

maximum expected mission time (>2 hours), with Xbee radio and antenna assembly

Verify GCS software is compatible with GCS computer

Verify serial communication can be established between antenna and GCS computer

Verify GCS can plot/present live data in real time during descent in SI units

COMMUNICATIONS: FSW

Verify with each software design iteration that programming language, functions,

libraries are compatible with chosen microcontroller

Verify software state is valid well defined at every conceivable point in mission

sequence, including environmental variations

Verify FSW successfully counts the number of packets transmitted, including following

power loss

MECHANISMS: Mechanical

Verify that in the event of component design iteration that resulting change in system

specifications (e.g. dimensions, weight) is compliant with requirements

MECHANISMS: DCS

Verify that in the event of component design iteration that resulting change in system

specifications (e.g. dimensions, weight) is compliant with requirements

Verify heatshield (all comprising materials, the interfaces between them, their

interface with the nose cone, and probe attachment components) can withstand

forces required to provide required descent rate

Verify parachute (all comprising materials, the interfaces between them and the

probe attachment components) can withstand forces required to provide required

descent rate

Verify that when assembled the heatshield has no openings or sharp edges

Verify heatshield, with clearances, complies with rocket body dimensions




Overview

Integrated functional level tests identified:


Test Description

PROBE: ENTIRE MISSION

Verify buzzer is of sufficient volume to help indicate the location of the probe

following landing

Verify battery compartment can be removed in less than 1 minute without total

CanSat disassembly

Verify full set of sensor data can be acquired at the required sampling rate

simultaneously

Verify power requirements are satisfied for the full mission sequence and

maximum predicted mission time

Verify that when code from Sensors, CDH and GCS subsystems are integrated

that microcontroller has sufficient memory for sketch, global variables and

EEPROM (non-volatile) memory

Verify fully assembled CanSat system weighs within 10g of 500g

Verify heatshield release mechanism is activated by the FSW at 300 metres

Verify parachute deployment mechanism is activated by the FSW at 300 metres

Verify that when assembled all electronic components except sensors are

enclosed

Verify power switch is easily accessible and reliably activates the system

PROBE: PRE-HEATSHIELD RELEASE + HEATSHIELD RELEASE TRIGGER

Verify that when probe is assembled the heatshield envelops the whole sides of

the probe

Verify heatshield deployment trigger performs as expected as part of the

constructed CanSat structure

Verify that with heatshield deployed the CanSat descends at the required

descent rate

Verify heatshield release trigger perform as expected as part of the constructed

CanSat structure

Verify centre of pressure of dummy probe is well below centre of gravity to

ensure stable descent (i.e. no end-over-end tumbling) and square heatshield

design assists stable descent (i.e. minimal rotation around descent axis)

BONUS – Verify camera can be activated by the FSW




Overview

Integrated functional level tests identified:


Test description

PROBE: POST-HEATSHIELD RELEASE

Verify parachute deployment mechanism performs as expected as part of

the constructed CanSat structure and parachute is released without

snagging

Verify that with parachute deployed the probe descends at the required

descent rate

COMMUNICATIONS: GCS

Verify GCS software handles live data from antenna+Xbee and plots it in real

time in SI units

Verify antenna receives data from assembled CanSat at and above the

required range

COMMUNICATIONS: CDH

Verify full live telemetry transmission is of the correct form for reception at

the GCS




Overview

Environmental tests identified:


Test Description

VIBRATION TEST - CANSAT

CanSat is fixed to an orbit sander to provide up to 14,000 rpm of rotation

equivalent to 233 Hz of vibration, to expose failure of CanSat

structures/components if they vibrate at resonance

THERMAL – CANSAT

Fully assembled CanSat is placed in controlled thermal chamber and heated

to ~80ºC while systems are active

15G LAUNCH ACCELERATION TEST - CANSAT

Fully assembled CanSat is placed in a sounding rocket for a test launch to

simulate launch conditions. This means the integrity of the heatshield, its

deployment mechanism, and probe attachment subsequent to launch can

be verified. This acceleration can be verified with the accelerometer within

the Sensors subsystem.

>30G DROP TEST – PROBE

Probe with heatshield released and parachute deployed is subjected to a

drop from 80cm with a cord attachment to result in a 48G shock

acceleration, to simulate rocket body separation forces. This acceleration

can be verified with accelerometer within the Sensors subsystem.

DIMENSIONS VERIFICATION - CANSAT

Fully assembled CanSat is subject to a fit check using a sheet of plywood

with a hole of diameter 125.5mm, to ensure clearances are sufficient to

ensure CanSat does not snag on rocket body.




Overview

System-level environmental testing: Test launches

• On 11/03/18 MCP performed a test launch using a Loc Precision Minie-Magg sounding

rocket, with a 20 inch payload bay, provided by MACE Space Research Group

• These allow the team to test the integrated system under launch conditions, verifying

compliance of both integrated functional tests and environmental tests.

• The crucial advantage of a test launch is to carry out all environmental tests

simultaneously, to ensure the payload, electronics, structures and mechanisms can perform

on launch day.

• MCP plans to carry out two further test launches, on 22/04/18 and 15/05/18.

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CanSat 2017 CDR: Team ### (Team Number and Name) 119

Subsystem level testing– In this section “Communications” is taken to be the CDH, GCS and FSW

subsystems. “Mechanisms” is taken to be the Mechanical and DCS

subsystems.

Test no. Test type Test description Req No. Pass/Fail Criteria

SENSORS

1 EVerify all sensors are correctly calibrated and configured

for their respective functions25

Sensors can provide correctly calibrated measurements of

known values

2 EVerify each sensor's range, resolution, sampling rate is as

required to provide valid data at a rate of at least 1Hz25 Sensors can provide valid data at 1HZ

3 V Verify collected data is in suitable format for FSW 25 Data should be in a format compatible with FSW

4 EVerify tilt sensor readings can be used to exhibit probe

stability during descent with heatshield deployed49

Check tilt sensor produces orientation data that can clearly

represent an object as stable (sufficiently small drift, noise)

5 EBONUS – Verify camera can produce colour footage at at

least 640x480 at least 30fpsR-Bonus

Camera produces at least 640x480 footage at 30fps when

supplied with voltage/current equivalent to that of the final

system

EPS

6 VVerify power source provides required power level

(current and voltage)-

Measured values of voltage and current are >= 9V and 1A

respectively

7 S,VVerify voltage divider provides required voltage to all

components-

Voltage divider works in theory, and provides 5V supplies a 5V

line

Test Procedure Experimental E Simulation S Verification V

TEST

PASSED

TEST NOT

ATTEMPTED

ATTEMPTED,

NOT PASSED

Test Procedures Descriptions

Subsystem level testing


Test no. Test type Test description Req No. Pass/Fail Criteria

COMMUNICATIONS: CDH

8 V

Verify radio can be configured to transmit as required: burst

or continuous transmission, API mode (NOT broadcast

mode), correct PAN ID, correct team ID, correct packet

29, 30 Test packets with these parameters are sent from Xbee to Xbee

9 EVerify real time clock can retain data following system

power loss27

After removing DC power to RTC time information is retained

when DC power is regained

10 S,V

Verify SD card can store the maximum expected volume of

data (maximum telemetry packet size multiplied by

maximum number of seconds)

25 SD card stores data successfully

11 VVerify full telemetry packet is transmitted correctly by the

radio module at a rate of 1Hz26 Xbee can send a test packet at 1Hz

12 E

Verify microcontroller regains correct function and retains

mission data (correct mission time using RTC) following

power loss

27 Mission data is stored throughout power loss


TEST

PASSED

TEST NOT

ATTEMPTED

ATTEMPTED,

NOT PASSED





TEST

PASSED

TEST NOT

ATTEMPTED

ATTEMPTED,

NOT PASSED


Test no. Test proc Test description Req No. Pass/Fail Criteria

COMMUNICATIONS: GCS

13 V

Verify GCS computer is suitable for competition final:

portable, battery life is suitable for maximum expected

mission time (>2 hours), with Xbee radio and antenna

assembly

36, 37 GCS computer qualifies for use on launch day

14 V Verify GCS software is compatible with GCS computer 33, 34, 35 GCS software runs as required when used on GCS computer

15 EVerify serial communication can be established between

antenna and GCS computer33, 34, 35

Dummy telemetry is received by the GCS software via antenna

and xbee

16 EVerify GCS can plot/present dummy data in real time in SI

units33, 34, 35 Dummy telemetry is plotted live

COMMUNICATIONS: FSW

17 V

Verify with each software design iteration that programming

language, functions, libraries are compatible with chosen

microcontroller

- FSW uploads correctly to microcontroller

18 E

Verify software state is valid and well defined at every

conceivable point in mission sequence, including

environmental variations

- FSW stays in a valid state at all times

19 EVerify FSW successfully counts the number of packets

transmitted, including following power loss39 Mission data is stored throughout power loss


CanSat 2017 CDR: Team ### (Team Number and Name) 122Presenter: Name goes here


MECHANISMS: Mechanical

20 V

Verify that in the event of component design iteration that

resulting change in system specifications (e.g. dimensions,

weight) is compliant with requirements

- New design retains compliance

21 S, EVerify all materials, structures and mechanisms can

withstand forces on the constructed CanSat-

CanSat structure and materials are can withstand forces in

theory, and are intact after physical testing

MECHANISMS: DCS

22 S,V

Verify that in the event of component design iteration that

resulting change in system specifications (e.g. dimensions,

weight) is compliant with requirements

- New design retains compliance

23 E

Verify heatshield (all comprising materials, the interfaces

between them, their interface with the nose cone, and

probe attachment components) can withstand forces

required to provide required descent rate

- Heatshield remains intact during simulated descent

24 E

Verify parachute (all comprising materials, the interfaces

between them and the probe attachment components) can

withstand forces required to provide required descent rate

- Parachute remains intact during simulated descent

25 S,VVerify that when assembled the heatshield has no openings

or sharp edges3, 9 No openings or sharp edges present

26 S,VVerify heatshield, with clearances, complies with rocket

body dimensions6 Heatshield dimensions are compliant


TEST

PASSED

TEST NOT

ATTEMPTED

ATTEMPTED,

NOT PASSED


Integrated Functional Level Testing– In this section the integrated system are split into:

• Probe: Entire mission

• Probe: Pre-heatshield release

• Probe: Post-heatshield release

• Communications: GCS



PROBE: ENTIRE MISSION

27 VVerify buzzer is of sufficient volume to help indicate the

location of the probe following landing41 Buzzer can be heard from a good distance

28 EVerify battery compartment can be removed in less than 1

minute without total CanSat disassembly42, 45 Battery compartment is easily removed

29 EVerify full set of sensor data can be acquired at the

required sampling rate simultaneously47 Sensor data are collected correctly

30 EVerify power requirements are satisfied for the full mission

sequence and maximum predicted mission time25 CanSat remains powered throughout with power to spare


TEST

PASSED

TEST NOT

ATTEMPTED

ATTEMPTED,

NOT PASSED



Integrated Functional Level Testing


PROBE: ENTIRE MISSION (continued)

31 S, E

Verify that when code from Sensors, CDH and GCS

subsystems are integrated that microcontroller has

sufficient memory for sketch, global variables and

EEPROM (non-volatile) memory

-Microcontroller has memory to spare after maximum

expected mission time

32 S, VVerify fully assembled CanSat system weighs within 10g

of 500g1 CanSat is within 10g of 500g

33 S, EVerify heatshield release mechanism is activated by the

FSW at 300 metres14

Heatshield releases at 300 metres, and/or activated by an

equivalent simulated trigger

34 S, EVerify parachute deployment mechanism is activated by

the FSW at 300 metres15

Parachute deploys at 300 metres, and/or activated by an

equivalent simulated trigger

35 S, VVerify that when assembled all electronic components

except sensors are enclosed18 Electronics excluding sensors are enclosed

36 VVerify power switch is easily accessible and reliably

activates the system21 Power switch is compliant


TEST

PASSED

TEST NOT

ATTEMPTED

ATTEMPTED,

NOT PASSED





PROBE: PRE-HEATSHIELD RELEASE + HEATSHIELD RELEASE TRIGGER

37 S, VVerify that when probe is assembled the heatshield

envelops the whole sides of the probe2 Probe is fully enveloped

38 EVerify heatshield deployment trigger performs as

expected as part of the constructed CanSat structure- Heatshield is deployed

39 EVerify that with heatshield deployed the CanSat descends

at the required descent rate43 Dummy probe descends at between 10 to 30 m/s

40 EVerify heatshield release trigger perform as expected as

part of the constructed CanSat structure- Heatshield is released

41 E, S, V

Verify centre of pressure of dummy probe is well below

centre of gravity to ensure stable descent (i.e. no end-

over-end tumbling) and square heatshield design assists

stable descent (i.e. minimal rotation around descent axis)

4, 5CanSat descends in a stable fashion both in theory and in

practice

42 V BONUS – Verify camera can be activated by the FSW R-Bonus Camera operation is controlled by FSW


TEST

PASSED

TEST NOT

ATTEMPTED

ATTEMPTED,

NOT PASSED





PROBE: POST-HEATSHIELD RELEASE

43 E

Verify parachute deployment mechanism performs as

expected as part of the constructed CanSat structure and

parachute is released without snagging

- Parachute is deployed without snagging

44 EVerify that with parachute deployed the probe descends

at the required descent rate44 Dummy probe descends at 5 m/s

COMMUNICATIONS: GCS

45 EVerify GCS software handles live data from antenna+Xbee

and plots it in real time in SI units33,34 GCS software plots live data in real time

46 EVerify antenna receives data from assembled CanSat at

and above the required range- Telemetry is received at test GCS

COMMUNICATIONS: CDH

47 EVerify full live telemetry transmission is of the correct

form for reception at the GCS26,27 Telemetry is valid and transmission successful


TEST

PASSED

TEST NOT

ATTEMPTED

ATTEMPTED,

NOT PASSED



Environmental Testing

Test no. Test Proc Test Description Test subject Rqmts Pass/Fail Criteria

VIBRATION TEST - CANSAT

48 E

CanSat is fixed to an orbit

sander to provide up to 14,000

rpm of rotation equivalent to

233 Hz of vibration, to expose

failure of CanSat components at

if they vibrate with resonance

Structures 16,17

CanSat structures comply with all expected tensile/compressive/torsional

loads during and following test period, including performance of heatshield,

parachute, and their respective attachment components

Mechanisms 22 CanSat mechanisms perform as required following test period

Egg state - Egg is intact after test period

Electronics -Electronics perform at a constant level (excluding the accelerometer) during

and following test period

THERMAL – CANSAT

49 E

Fully assembled CanSat is

placed in controlled thermal

chamber and heated to ~80ºC

while systems are active

Structures -

CanSat structures comply with all expected tensile/compressive/torsional

loads during and following test period, including performance of heatshield,

parachute, and their respective attachment components

Mechanisms - CanSat mechanisms function as required during and following test period

Egg state - Egg is intact after test period

Electronics - Electronics perform at a constant level during and following the test period


TEST

PASSED

TEST NOT

ATTEMPTED

ATTEMPTED,

NOT PASSED





15G LAUNCH TEST - CANSAT

50 E

Fully assembled CanSat is placed in a

sounding rocket for a test launch to simulate

launch conditions. This acceleration can be

verified with the accelerometer within the

Sensors subsystem.

Structures 19,20

CanSat structures comply with all expected dimensional,

tensile/compressive/torsional loads during and following shock

period, including performance of heatshield and parachute

and their respective attachment points

Mechanisms 22 CanSat mechanisms perform as required following test period

Egg state 7, 8 Egg is intact after shock test

Electronics 21Electronics perform as required during and following test

period

>30G DROP TEST – PROBE

51 E

Probe with heatshield released and

parachute deployed is subjected to a drop

from 80cm with a cord attachment to result

in a 48G shock acceleration. This

acceleration can be verified with

accelerometer within the Sensors

subsystem.

Structures 19,20

Probe structures comply with all expected dimensional,

tensile/compressive/torsional loads during and following drop,

including performance of parachute and its attachment point

Mechanisms 22 CanSat mechanisms perform as required following drop test

Egg state 7,8 Egg is intact after drop test

Electronics 21Electronics perform as required during and following test

period


TEST

PASSED

TEST NOT

ATTEMPTED

ATTEMPTED,

NOT PASSED



• After subsystem, integrated functional level, and environmental testing, 40 of 50 requirements are

mapped to a test.

• The remaining 10 requirements primarily require initial verification of designs and ongoing any

subsequent design iterations.

• Responsibility for compliance of the system to these requirements falls to the Integration & Testing

leader (LAF), who will frequently confirm compliance.


DIMENSIONS VERIFICATION - CANSAT

52 S, V

Fully assembled CanSat is subject to a fit

check using a sheet of plywood with a hole

of diameter 125.5mm, to ensure clearances

are sufficient to ensure CanSat does not

snag on rocket body..

Structures 11, 12, 13 CanSat can slide through hole in sheet



TEST

PASSED

TEST NOT

ATTEMPTED

ATTEMPTED,

NOT PASSED


Remaining Requirements

• So 50 (49 + bonus) requirements are mapped to tests.


Rqmt No. Requirement descriptionRelevant subsystem/mission section

Evidence of compliance

10 The aero-braking heat shield shall be a florescent color; pink or orange. DCSSee DCS “Descent Control Overview” slide

23 Mechanisms shall not use pyrotechnics or chemicals. Mech + DCSMech “Physical Layout” and DCS “Payload Descent Control Hardware Summary” slides

24Mechanisms that use heat (e.g., nichrome wire) shall not be exposed to the outside environment to reduce potential risk of setting vegetation on fire.

Mech + DCS Mech “Physical Layout” and DCS “Payload Descent Control Hardware Summary” slides

28XBEE radios shall be used for telemetry. 2.4 GHz Series 1 and 2 radios are allowed. 900 MHz XBEE Pro radios are also allowed.

CDHCDH SLIDE “Probe Radio Configuration” slide

31Cost of the CanSat shall be under $1000. Ground support and analysis tools are not included in the cost.

PROJECT MANAGEMENTSee Management “CanSat Budget –Hardware” slide

32 Each team shall develop their own ground station. GCS See GCS “GCS Software” slide

38Both the heat shield and probe shall be labelled with team contact information including email address

MISSION OPERATIONSSee Mission Operations & Analysis “CanSat Location and Recovery” slide

40 No lasers allowed. SensorsSee “Sensor Subsystem Overview” slide

46Battery source may be alkaline, Ni-Cad, Ni-MH or Lithium. Lithium polymer batteries are not allowed. Lithium cells must be manufactured with a metal package similar to 18650 cells.

EPS EPS “Probe Power Source” slide

48Spring contacts shall not be used for making electrical connections to batteries. Shock forces can cause momentary disconnects.

EPS MECH SLIDES



Mission Operations & Analysis

Iuliu Ardelean


Overview of Mission Sequence of

Events

ROLES & RESPONSIBILITIES





FINAL INTEGRATION AND TESTING:

• Between 0800 and 1200.

• Full team involved, except NZ.

• Multiple CanSat I&T procedures will be done before the competition to

ensure everything runs smoothly. Performed by CanSat Crew.

• Antenna and GCS setup will be performed by GCS Crew. Simple plug-and-

play philosophy stands behind the design of the GCS and Antenna systems.

• A detailed I&T Plan will be created to aid this process – see Mission

Operations Manual.

• Telemetry Data File handed to Competition Officials after termination of GCS

Operation.



Overview of Mission Sequence of

Events

Arrival – 0800 –Full Team

Final Integration and Testing –CanSat Crew

Official Inspection –

1200 – NZ

Collect CanSat –NZ

CanSatintegration with

rocket – AS

Check Communications

- IA

GCS & Rocket with CanSat

transportation to Launchpad – NZ,

AS, IA

Rocket Installation –

Officials

GCS operational – GCS Crew

Launch Procedures

Execution – NZ

Flight + GCS Operational –

GCS Crew

All CanSatLaunched

Recovery –Recovery Team

Handout CanSatfor Final Judging

– NZ

Terminate GCS Operation – IA

Submit USB with collected and

received data –NZ

Begin PFR work



Field Safety Rules Compliance

The Mission Operations Manual will be compiled at individual- and team-level to ensure suitable accuracy and

detail of tests and procedures. The Mission Operation Manual will contain instructions to the following:

1. CanSat Integration and Testing Responsible: LAF, RS, IA, NZ

1.1. Integration Procedure (can be skipped)

1.2. Testing Procedure

1.3. Operational Checks

2. GCS Setup and Operation Responsible: IA, NZ, LAF

2.1. Setup Procedure

2.2. Operational Checks

3. CanSat-Rocket Integration Responsible: LAF, AS

4. Launch Responsible: Competition Staff

4.1. Preparation Procedure

4.2. Launch Procedure

5. Other Procedures Responsible: Full team

The Mission Operations Manual work will begin after the 22nd of April Test Launch.

The Mission Operations Manual will be printed (and suitably bound) in multiple copies and distributed across team

members.

One copy will be handed to the Launch day Flight Coordinator.



CanSat Location and Recovery

BUZZER

Continuous beeping on Probe

COLOR

Bright Orange

GPS LOCATION

Using Acquired Telemetry data

TEAM MEMBERSWill track down the

CanSat as it descends

The following measures will ensure that the CanSat including Probe and Heatshield will be recovered.

Moreover, in case the CanSat is not recovered, both the Probe and the Heatshield will be labeled with the

Manchester CanSat Project’s address (including email) and all other relevant contact details.

Manchester CanSat Project, University of Manchester

Team 5002, [email protected]

George Begg Bulding, University of Manchester, M1 7DN

Manchester, United Kingdom

mailto:[email protected]

Mission Rehearsal Activities


Activities

When it will be rehearsed

Weekly Lab

Session

Full System

Test Launch

11th of March

Full System

Test Launch

22nd of April

Ground system radio link check procedures X X X

Powering on/off the CanSat X X X

Launch configuration preparations (e.g., final

assembly and stowing appendages)X X X

Loading the CanSat in the launch vehicle X X

Telemetry processing, archiving, and analysis X X X

Recovery X X

The team already had a full system test. Some results were unsatisfactory. We identified the issues were related

to stripboard short-circuiting. The team plans to order a PCB to solve this.

Mission Rehearsal Activities


The Mission Operations Manual will be created to guide the following parties during mission day.

The Mission Operations Manual work will begin after the 22nd of April Full System Test.





The Chief Engineer (IA) will ensure the coordination of all parties.


Requirements Compliance

Iuliu Ardelean


Overview


Current design complies with all requirements, except Bonus 2, which is not being attempted.

Design shall be tested to ensure Requirement Compliance, following the procedure explained in the

Integration and Test section of this document.

The Design has been altered since the PDR, because the PDR weight estimates were too high. The most

important constraint on the design is the 500g weight requirement. The current design has been built and it

DOES satisfy this requirement, along with all other ones.

At the moment, the team’s main focus is building and manufacturing an all round sturdy CanSat.

The following 3 slides trace and demonstrate compliance with all Requirements. Comments have been

added where necessary.

The legend gives color coding to indicate if a Requirement is met.

Comply

Partial

No Comply



(multiple slides, as needed)


RE# Description Compliance

Reference

Slides CommentsRE1 Total mass of the CanSat (probe) shall be 500 grams +/- 10 grams. 68

RE2

The aero-braking heat shield shall be used to protect the probe while in the rocket only and when deployed


rocket. The rear end of the probe can be open 19, 32RE3 The heat shield must not have any openings. 32, 39 – 45 RE4 The probe must maintain its heat shield orientation in the direction of descent. 45RE5 The probe shall not tumble during any portion of descent. Tumbling is rotating end-over-end. 45

RE6

The probe with the aero-braking heat shield shall fit in a cylindrical envelope of 125 mm diameter x 310 mm

length. Tolerances are to be included to facilitate container deployment from the rocket fairing. 19RE7 The probe shall hold a large hen's egg and protect it from damage from launch until landing. 61 – 62

RE8

The probe shall accommodate a large hen’s egg with a mass ranging from 54 grams to 68 grams and a

diameter of up to 50mm and length up to 70mm. 61 – 62

RE9

The aero-braking heat shield shall not have any sharp edges to cause it to get stuck in the rocket payload

section which is made of cardboard. 19RE10 The aero-braking heat shield shall be a florescent color; pink or orange. 44, 135RE11 The rocket airframe shall not be used to restrain any deployable parts of the CanSat. 19RE12 The rocket airframe shall not be used as part of the CanSat operations. 11 – 13 19

RE13 The CanSat, probe with heat shield attached shall deploy from the rocket payload section. 11 – 13RE14 The aero-braking heat shield shall be released from the probe at 300 meters. 24, 96RE15 The probe shall release a parachute at 300 meters. 24, 96

RE16

All descent control device attachment components (aero-braking heat shield and parachute) shall survive 30

Gs of shock. 66 – 67RE17 All descent control devices (aero-braking heat shield and parachute) shall survive 30 Gs of shock. 66 – 67

RE18

All electronic components shall be enclosed and shielded from the environment with the exception of

sensors. 66 – 67 RE19 All structures shall be built to survive 15 Gs of launch acceleration. 66 – 67RE20 All structures shall be built to survive 30 Gs of shock 66 – 67

RE21

All electronics shall be hard mounted using proper mounts such as standoffs, screws, or high performance

adhesives. 66 – 67 RE22 All mechanisms shall be capable of maintaining their configuration or states under all forces 51 – 73 RE23 Mechanisms shall not use pyrotechnics or chemicals. 51 – 73




RE# Description Compliance Reference Slides Comments

RE24

Mechanisms that use heat (e.g., nichrome wire) shall not be exposed to the outside environment to reduce

potential risk of setting vegetation on fire. 51 – 73

RE25

During descent, the probe shall collect air pressure, outside air temperature, GPS position and battery

voltage once per second and time tag the data with mission time. 21 – 30, 79, 96

RE26

During descent, the probe shall transmit all telemetry. Telemetry can be transmitted continuously or in

bursts. 80 – 83, 96

RE27

Telemetry shall include mission time with one second or better resolution. Mission time shall be maintained

in the event of a processor reset during the launch and mission. 79, 96

RE28

XBEE radios shall be used for telemetry. 2.4 GHz Series 1 and 2 radios are allowed. 900 MHz XBEE Pro

radios are also allowed. 80 – 83

RE29 XBEE radios shall have their NETID/PANID set to their team number. 80 – 83

RE30 XBEE radios shall not use broadcast mode. 80 – 83

RE31 Cost of the CanSat shall be under $1000. Ground support and analysis tools are not included in the cost. 145 – 149

RE32 Each team shall develop their own ground station. 107 – 109

RE33 All telemetry shall be displayed in real time during descent. 107 – 109

RE34 All telemetry shall be displayed in engineering units (meters, meters/sec, Celsius, etc.) 107 – 109

RE35 Teams shall plot each telemetry data field in real time during flight 107 – 109

RE36

The ground station shall include one laptop computer with a minimum of two hours of battery operation,

XBEE radio and a hand held antenna. 104 – 106

RE37

The ground station must be portable so the team can be positioned at the ground station operation site

along the flight line. AC power will not be available at the ground station operation site. 104 – 106

RE38 Both the heat shield and probe shall be labeled with team contact information including email address. 135

RE39

The flight software shall maintain a count of packets transmitted, which shall increment with each packet

transmission throughout the mission. The value shall be maintained through processor resets. 96

RE40 No lasers allowed. 21

RE41 The probe must include an easily accessible power switch. 85 – 86, 89

RE42 The probe must include a power indicator such as an LED or sound generating device. 85 – 86, 89, 96




RE# Description Compliance

Reference

Slides Comments

RE43 The descent rate of the probe with the heat shield deployed shall be between 10 and 30 meters/second. 47 – 49

RE44

The descent rate of the probe with the heat shield released and parachute deployed shall be 5

meters/second. 47 – 49

RE45 An audio beacon is required for the probe. It may be powered after landing or operate continuously.

85 – 86, 89, 96

RE46

Battery source may be alkaline, Ni-Cad, Ni-MH or Lithium. Lithium polymer batteries are not allowed.

Lithium cells must be manufactured with a metal package similar to 18650 cells. 85 – 86, 90

RE47

An easily accessible battery compartment must be included allowing batteries to be installed or removed in

less than a minute and not require a total disassembly of the CanSat. 58 – 60

RE48

Spring contacts shall not be used for making electrical connections to batteries. Shock forces can cause

momentary disconnects. 89

RE49

A tilt sensor shall be used to verify the stability of the probe during descent with the heat shield deployed

and be part of the telemetry. 28

Bonus

1

Camera: Add a color video camera to capture the release of the heat shield and the ground during the last

300 meters of descent. The camera must have a resolution of at least 640x480 and a frame rate of at least

30 frames/sec. The camera must be activated at 300 meters. 29

Bonus

2

Wind Sensor: A radio transmitter shall be added to transmit the wind speed by changing its 10 frequency.

The frequency change shall be 1 Hz per 0.1 meter/sec. The transmitter must be custom designed and built.

It cannot be a commercial product. The frequency must be in the 433 MHz ISM band or if a team member

has an amateur radio license, an amateur radio band can be used. The transmitter must be able to be set to

8 different frequencies in the 433 MHz ISM band with 25 KHz separation. The transmitter must turn off after

the probe lands to minimize interference. The team can use a commercial receiver. 30, 110

Not an Issue, as this bonus is not

being attempted.


Management

Iuliu Ardelean


Status of Procurements

•

•

Status of Procurement

Absolutely all components necessary

have arrived.

We have enough spares

for building two CanSats now.

We intend to purchase two more sets of components – to support testing and

manufacturing.

Philosophy No. 1 – “practice makes perfect”

Philosophy No. 2 – “throw in money until it starts working”

Procurement is not exactly a concern for the team, as most components tend to arrive

within one-two working day.

A full list of components can be found in the following slides.


CanSat Budget – Hardware

•

•Subsystem Estimated Cost

Structures ₤72.83

Electronics ₤174.06

Tools ₤0

Total ₤246.89

The following table shows the estimated budget for hardware in subsystems of the CanSat:

Legend

Estimated XX Actual XX



Electronic components

Part Name Function Reuse Quantity Total Cost (₤) Total Cost ($)

Adafruit 10-DOF IMU Temp., Press., Alt, Tilt No 1 21.11** 29.95**

Adafruit Ultimate GPS

Breakout

GPS No 1 40* 56.75*

Modified SQ11 Pawaca

Camera

Camera No 1 14.99* 20.90*

Teensy 3.2 USB

Microcontroller

Microcontroller No 1 19.80* 28.08*

Breakout for SD Card On board data storage No 1 4.20* 5.86*

16 GB SD Card SD Card No 1 6.80* 9.48*

DS1338 RTC No 1 3.08* 2.84*

XBee Pro S2C Transceiver No 2 52.42* 74.37*

Energizer Lithium Battery No 1 7.05* 9.83*

Servo Mechanisms No 1 4* 5.67*

Switch On/Off Switch No 1 0.61* 0.85*

Total 174.06 242.69

Legend


*Current Market Value

**Market Value of Discontinued Item



Legend


3D printed components

Equipment Part Name/Specifications Reuse Quantity Total Cost (₤) Total Cost ($)

Holder Egg Containment

No

1kg

(including

failures/pro

totyping)

@22 per kg of

spool = 22

31.22

Cover Egg Containment

Nose Cone HS

Deployment Bay HS

HS Attachment -

Plates (Floors) HS Release Mechanism Bay,

Camera Bay, Parachute Bay

Camera Bay -

Electronics Cover -

Parachute Bay -

Total 22.00 31.22



Off the shelf components

Equipment Part Name/Specifications Reuse Quantity Total Cost (₤) Total Cost ($)

Sponge Egg Containment No - 1 1.42

Nuts and Bolts M3 and M1.6 No 20 11 15.61

Carbon Fiber Rods HS structure No 4 7 9.93

Springs HS deployment No 10 13 18.45

Nylon HS material No - 4.99 7.08

Bearings HS deployment No 2 1.96 2.78

Wire HS and Parachute No - 1.95 2.77

Carbon FiberSpacers 120 mm and 30 mm No 6 2.95 4.19

Horn Parachute Release Mechanism No 1 2.69 3.82

Rod Parachute Release Mechanism No 1 3.25 4.61

Hinge Parachute Release Mechanism No 1 1.03 1.46

Total 50.83 72.14


CanSat Budget – Other Costs

Source Amount (₤) Additional Information

School of MACE 5,000 Possibility of increasing to 10,000

School of Physics 3,000 -

BAE Systems 2,000 -

Aerospace Research Institute 500 -

Fund IT Students Union 500 -

Airbus 5,000 -

Income

Total Income Confirmed (₤) 16,000

Legend



CanSat Budget – Other Costs

Detail Description Unit Cost Quantity Total Cost

Travel,

Accommodation

and Sustenance

Costs

Travel Flights, rental car, train ₤780 10 People ₤7800

Visas Student/Tourist Visa ₤113.12 6 People ₤678.72

Housing Based on a stay from

07/06/2018 to 10/06/2018

₤100 10 People ₤1000

Food Assuming ₤15/person/day ₤60 10 People ₤600

GCS Hardware

Cost

Display Laptop (provided by team

member)

N/A 1 N/A

Emergency Can

Sat

All Can Sat

parts

- £303.49 1 £303.49

Competition

Entry Fee

- - ₤70.72 1 ₤70.72

UK CanSat

Competition

Costs

Organization

and

Participation

- ₤3000 1 ₤3000

Legend


151

Program Schedule

These Gantt Charts were developed using Microsoft Project.

1. Academic Gantt Chart is displayed above

2. Project Gantt Chart is split across the next 2 slides

a) Chart is made using summary tasks from the detailed task list

shown on slide after Gantt Chart

b) Chart uses linkages (i.e., FF, FS, SS, SF) and lag periods to

show dependence on other tasks

c) Deliverables used to set internal deadlines and milestones; seen

in more detail in task list

Gantt Chart Colour Coding Legend

Time Period: Normal

Time Period: Potential Hindrance to work done

Time Period: No work done

Tasks

Deadline

Milestone

Microsoft Project File

https://1drv.ms/u/s!AnWXOhepIwD_hJ4gomFC9U2PbAIVjw

Presenter: Iuliu Ardelean CanSat 2018 CDR: Team 5002

152

Program Schedule

GANTT Chart Overview

Includes: Competition Milestones, Major Development Activities, Component/Hardware deliveries AND Major

I&T activities and milestones.


153

Program Schedule

GANTT Chart Overview (continued)

Includes: Competition Milestones, Major Development Activities, Component/Hardware deliveries AND Major

I&T activities and milestones.


154

Program Schedule

Project Task List


155

Program Schedule

Changes since PDR:

• Updated to include PDR Results (Invitation

to Final) on 13/03/18 and associated

marksheet analysis task

• Updated to include dates on past and future

test launches – for environmental testing

• Test Launch 1: 11/03/18 (occurred)

• Test Launch 2: 22/04/18

• Test Launch 3: 13/05/18

Project Task List (continued)


156

Program Schedule

CRITERIA

PERCENTAGE

DONE

Project Analysis 100%

Mission Analysis 100%

System Concept 100%

Subsystem Design 100%

Procurement + Manufacturing 100%

Subsystem Testing 92%

System Integration 60%

Integrated Functional Level Testing 75%

PDR 100%


Test launches 33%

Drop, vibe, therm, dim 76%

Design Iterations 95%

CDR 100%

Requirements compliance (49/49) 100%

Prototype cloning (aim to have 4 clones) 25%

Destructive environmental testing 10%

OVERALL 79%


Shipping and Transportation


The team plans to build two fully functional CanSats, which will be shipped to Stephenville, Texas using an

express courier service that can guarantee the delivery before the arrival of the full team in Stephenville.

The team will liaise with the competition officials to sort this matter out.

The team also plans to carry the exact same equipment for another two CanSat with them to ensure enough

spares are available in case of a contingency.

From an initial analysis, sending a 5kg package worth 2000 GBP, the cost is 100-150 GBP.


Conclusions


Major accomplishments

The design is now well within the weight limit.

Coding is 100% finalized.

A very generous budget has been achieved.

We are now more than two month ahead of our real schedule from last year.

We have actually built and almost completed testing of our design.

We are confident to say we know what can fail and what will not.

We have inspired a handful of UK Universities to join the UK Competition.

Major unfinished work

Some testing (please see below)

Travel Arrangements including plane tickets and VISAs.

Operations Manual

Testing to complete

Major Packages associated to

Radio Communication

Parachute Mechanical Strength

Egg Protection Structure

Bonus 2 Video Camera

Flight software status

100% Finalized


Documents

CanSat 2018 Critical Design Review (CDR) Outline Version 1cansatcompetition.com/docs/teams/Cansat2018_5002_CDR_v12.pdfCanSat 2018 CDR: Team 5002 4 Acronyms HS Heat shield CDH Communications