2011 06 17

Cubesat Array for Emergency Search And Rescue

CAESARקיסאר ראשוני ואיתור לסיוע ייעודית קונסטלציה

Team Members:

Yuval PoratMoshe SederoHanan AmarNoam LeshemNoam LeiterOshrat MarfogelIttai CohenNitsan Bavli

Supervisor:Jacob Herscovitz

Winter 2010-2011

הצוות חברי

פורת יובלסדרו משה

עמר חנןלשם נעםלייטר נעם

מרפוגל אושרתכהן איתיבבלי ניצן

: הפרויקט מנחההרשקוביץ יעקב

Nano-Satellite Student ProjectCritical Design Review

קונסטלציה CAESAR קיסארראשוני ואיתור לסיוע Cubesat Array for Emergency Search And Rescue ייעודית

April 13, 2023 2

Contents:1. Background2. Requirements3. Completed PDR Design

and Summary 4. System engineering 5. Orbits and Constellation 6. Geo-location7. Formation Flying

8. Structure 9. ADCS sub-system10. Propulsion sub-system11. Cost Estimation12. Risk Management13. Reliability14. Work Breakdown Structure (WBS)15. Summary and Acknowledgments


Background• The oceanic surrounding is hazardous and present risks of

drowning , hypothermia, shark attacks and more…• Due to the nature and size of the oceanic surrounding, the

process of receiving distress signals and locating people in distress accurately is somewhat problematic.

• Between hundreds to thousands of sea-related accidents occur every year.

April 13, 2023 3


Customer Requirements1. The system shall locate and a person in distress in any watery surrounding around

the world (oceans, seas, rivers…)2. The user shall wear an emergency beacon that will transmit a distress signal when

activated.3. The time interval from distress signal transmission to notification in one of the

ground stations shall not exceed 15 minutes.4. The computed location of the person in distress shall be no more than 1 km of his

true location.5. As an option, the system shall allow enhanced capability for future applications

such as search and rescue services for “land incidents”, given the appropriate modifications.

6. The system shall be based on space and satellites technology.7. The space segment should be implemented using Nano-satellites ("Cube-Sat").8. Each satellite's mission life-time shall be at least 2 years.

April 13, 2023 4


Top Level Mission Requirements• A user in distress shall be detected in less than 15 minutes, from signal transmission to ground

station notification.• A user in distress shall be geo-located with an accuracy < 1 km• The distress signal shall be relayed to a ground station • The system's services shall be affordable to the common end user.• The system shall be capable to identify its users in distress, as valid subscribers. • Earth coverage range shall be at least between latitudes )+60⁰) and (-60⁰)• International space-related standards and regulations should be met, as much as possible

Top-level System Requirements:• "Cubesat" satellite platforms shall be considered. • Each satellite's mass shall be less than 10 kg • Each satellite life time shall be at least 2 years.• Satellite bus shall be designed using space-proven COTS sub-systems and components, as much as

possible.• Satellite's sub-systems shall withstand launch load and space environment.• Geo-location shall be performed using DTOA technique, using 2 or 3 reception satellites.

April 13, 2023 5


Completed PDR Design Work:Electric Power System:

Thermal Control:•Passive Control•Steady State mean temperature: -6⁰

April 13, 2023 6

EPS + Matching Battery

Solar Array to Battery Efficiency

Battery to Consumers Efficiency

Max DoD Mass

ClydeSpace 3U EPS + Battery Pack 90% 90% 20% 170 g

Efficiency @ 5E14 e-/cm2 Cell Weight Cell Area Efficiency (BOL) Qty . Solar Panel

26.5% 2.6 g 30.18 cm2

29.1% 26 Azure TJ 3G30C


Communication:

Broadcasting:

April 13, 2023 7

April 13, 2023

Telemetry &Control

Satellite to Satellite

Geolocation User Segment

Ground Station

Satellite Ground Station Satellite

Yagi 2 Dipole 2 Dipole Parabolic Patch Monopole Dipole

Antenna

400Mhz 400Mhz 450Mhz -- 2.4Ghz 2.4Ghz Transmitter

400Mhz 400Mhz 450Mhz 2.4Ghz 2.4Ghz -- Receiver

User's Beacon

Ground Station for GeoLocation data receiving and handling

Ground Station for Control and

Telemetry

GPS

Main CubeSat

Other CubeSat


Launch Segment:1. Poly-PicoSatellite Orbital Deployer “P-Pod MKIII”:• mass 1.5 kg• can carry 3 (1U) cubesats or 1 (3U) cubesats• number of deployers can be mounted together on a L.V

2. Launch Vehicle: SpaceX - Falcon 1e

April 13, 2023 8

Inclination[deg] Altitude Mass capability

[kg]Payload

space[m]

Accuracy Reliability Est. Cost

Any above 9⁰ LEO 800 to 700[km] D1.55 x

H1.7i = 0.1 [deg]

Apogee = 15[km] Med $10.9M


PDR Summary - Mission:

Constellation : 700 km , , , 48 satellites, 6 planes Geolocation: TDOA algorithm, 97% location within 15 min,

3% location within 30 minFormation: 2 satellites, In-plane formation, relative control,

distance = 200 ± 50 km

April 13, 2023 9

0e45i


PDR Summary - System:

Mass: 3.11 kg Communication: 2 dipole, receiver, transmitter

Thermal Ctrl: Passive Payload: Patch antenna, transceiver

Attitude Ctrl: Active, 3-axis Propulsion: Warm gas, Isp=100Available Average Power: 6.78 WApril 13, 2023 10


System Engineering(Updates)

April 13, 2023 11


April 13, 2023 12

Dispersion Mission Operation

14 days 2 years

Initial stabilization

24 hours

Launch and Deployment

De-orbiting

1-2 years10 Mins

x

y

z

Mission Profile


Budgets

April 13, 2023 13

ΔV Budget Mass Budget Power Budget

Sub SystemPDR

System Total Mass [Kg]

CDR System Total

Mass [Kg]

OBDH 0.08 0.08

ADCS 0.209 0.09

Propulsion 1.209 0.458

Thermal Control

0 0

Communication 0.23 0.23

Payload 0.105 0.105

GPS 0.003 0.003

Power 0.297 0.237

Structure 0.958 1.02

De-Orbit - 0.08

Total 3.111 2.3

Usage PDR ΔV[m/s]

CDR ΔV[m/s]

Positioning 9.19 9.19

Keeping Formation 9.32 0.18

Deorbiting 13.35 0

Spare (20%) 6.37 0.94

Total 38.23 10.31

ConsumersPower consumption [mW]

Cruise Detection Maneuver

OBDH 200 600 600

ADCS 430 630 630

Propulsion 0 0 2000

Thermal Control 0 0 0

Communication 200 450 450

Payload 200 450 450

GPS 200 200 200

EPS 200 200 200

Structure 0 0 0

De-Orbit 0 0 0

Total 1430 1880 4530


Design Iteration:Subsystem’s Mass: Satellite’s Mass:

April 13, 2023 14

Sub-SystemAllocation for Sub-System

[Kg]Sub-System

Total Mass [Kg]Power 0.31 0.2376ADCS 0.22 0.09

Thermal Control 0 0Communication 0.265 0.23

Payload 0.11 0.105GPS 0.004 0.003

OBDH 0.1 0.08Propulsion 0.7 0.458Structure 1.05 1.02De-Orbit 0.2 0.08Total 2.955 2.3

Mass [Kg] CommentsDry Mass 2.3

10% Margin 2.53 X+10%

Fuel 0.031 Includes 10% Margin for Fuel

Total 2.56Includes:

10% margin for Fuel and 10% margin for Dry Mass


System Hierarchy Diagram:

April 13, 2023 15


April 13, 2023 16

OBDH

Main

computer

S-Band Comm.

S-Band

Receiver

S-Band Antenna

2 UHF Antennas

UHF Comm.

UHF

Transmitter

Satellite Block Diagram

GPS

GPS

Antenna

Power

EPS

Battery

Propulsion

2x Thruster

Pr. Tank

Valves

Filter

Regulator

UHF

Receiver

De-Orbit Device

“Nano

Terminator”

Attitude Determination & Control

3x Magneto-Torquers

3x Magneto-Meters

Photo Voltaic Cells


Physical Hierarchy:System interfaces ( diagram):

April 13, 2023 17

2N


Mission Design(Updates)

April 13, 2023 18


Orbits and ConstellationPDR Summary:• A Walker constellation 45:24/6/1

• Constellation altitude - 700 km

• Constellation inclination of 45⁰

• Total of 48 satellites.

• Total of 24 formations

• 2 satellites per formation with nominal distance of 200 km between satellites.

• 6 orbital planes, each orbital plane consisting of 8 satellites (4 formations)

• Satellite de-orbitization at EOL using propulsion to lower the satellite from 700 km to 650 km, requiring

April 13, 2023 19

sec13.35 mv


Design Updates Since PDR: • Altitude had been changed from 700 km to 710 km.• At 710 km ionization dose is about 6 krad for 0.6 mm shielding thickness.

still well within the 10 krad restriction of the sensitive EPS system.• In 2 years (mission life time) satellites decline approximately 10 km.

at EOL, altitude is around 700 - higher than the minimum of 697 km • No altitude correction maneuvers are required throughout the entire mission.

April 13, 2023 20

'710' '709' '708' '707' '706' '705' '704' '703' '702' '701' '700' '699' '698' '697' '696' '695' '694' '693'13

13.514

14.515

15.516

16.517

17.518

Constellation Revisit Time Vs Altitude

Altitude [km]

Revi

sit

Tim

e [m

in]


De-Orbiting• PDR calculation for de-orbiting from 700 km to 650 km: • From 710 km to 650 km – even higher: • 2 Alternatives for De-Orbiting were considered:

Alternative #1: “Jack in the Box”• De-Orbit mechanism designed and manufactured by NASA for the O/OREOS

mission.

April 13, 2023 21

13.35sec

mv

16.01sec

mv


• NASA’s de-orbit mechanism increases satellite’s surface area, and thus drag force, by 60%

• Device’s dimensions:

April 13, 2023 22

Material: Aluminum plates Germanium Film Weight: ~200 gr (est.)

Device can be placed only on top or bottom panel

28 cm9.9 cm9.9 cm


April 13, 2023 23

Alternative #2: Tether Unlimited © nanoTerminator• Designed and manufactured be Tether Unlimited ©

Specifications: • Mechanism consists of a 30-m long, 0.8-mm thick conductive tape.• Mechanism can be mounted on every panel.• Photo-voltaic sells can be integrated onto it • Mechanism mass is ~ 80 grams


Method of Operation:• The conductive tape produces

current up the tape upon interaction with ionspheric plasma

• Charged tape interacts back with earth’s magnetic field to produce Lorentz Forcethat opposes orbital motion and produces electrodynamic drag.

April 13, 2023 24

0

L

F I B dl


Device’s Performance:• The extended tape’s surface area is about 152 cm², increases

spacecraft surface area by 50 %• Deorbit Time Prediction with mechanism:

April 13, 2023 25

CAESAR Satellites


De-Orbit Mechanism Selection

The nanoTerminator gives us the best deorbit time, for the lowest additional mass, and is the easiest to integrate with the satellite.

April 13, 2023 26

Criterion Criterion Weight Propulsion Based "Jack-in-the-Box" nanoTerminator

Value Score Total Value Score Total Value Score Total

Mass 0.5 400 gr 1 0.5 200 gr 3 1.5 80 gr 5 2.5

Deorbit Time 0.2 24.4 yr 2 0.4 22.2 yr 3 0.6 <1 yr 5 1

Compat-ibility 0.3 1 0.3 3 0.9 5 1.5

Total 1.2 3 5


GeolocationThe TDOA location method

April 13, 2023 27

21 2 1

2 2 2

1 1

i i i i

t s u s uc c

s u X X Y Y Z Z

c is the speed of light

• The hyperbolic equation can be transformed to a quadratic form

00

2 2 2 21 2 1 2 2 1 1 2 1 2

22 2 2 2 2 20 2 1 1 2 21

2 0 1 01

4 4 2 2

2

T T TT

T

M m uu Mu m u m u

m m

M s s s s d I m s s s s d s s

m s s d s s d d c t


• If no measurement errors exist the target must lie on the hyperboloid defined by this quadratic form where the 2 satellites in the formation are the focal of the hyperboloid. In this case 3 TDOA measurements can define the 3 unknown target coordinates.

April 13, 2023 28

-5

0

5

-4

-3

-2

-1

0

1

2

3

4

5

Satellite Formation and Target on TDOA Hyperboloid

X

Y

SAT1

SAT2

Target

TDOA Hyperboloid

-5

0

5

-4

-2

0

2

4

-4

-3

-2

-1

0

1

2

3

4

Z

Satellite Formation and Target on TDOA Hyperboloid

XY

SAT1

SAT2Target

TDOA Hyperboloid


• If the targets location is known to be constrained on the surface of the Earth only 2 more TDOA’s are needed to find the location.

• Based on the analytical solution shown by Ho and Chan for a 3 satellite formation and a single TDOA measurement, we have derived an iterative algebraic method for a 2 satellite formation using 2 TDOA measurements.

April 13, 2023 29-10

-5

0

5

-5

0

5

-2

0

2

4

6

8

Z

Target in the Intersection of a Sphere and 2 TDOA Hyperboloids

Y X

SAT11

SAT21

TargetSAT12

SAT22


• In the presence of measurement errors the initial location can be far from the true location of the target. In order to improve the initial location error an Extended Kalman Filter starting with the initial location is used with all of the TDOA data. The estimated target location then drifts from the initial location closer to the true location.

April 13, 2023 30


Experiment Scale Down• In order to improve the reliability of the geolocation algorithms and

examine them in a more realistic environment we have conducted an experiment at the Distributed Space Systems Laboratory (DSSL) in the Asher Space Research Institute.

April 13, 2023 31

EchoLab Scale Space Scale Parameter

~500 mm ~200 km Formation

3000-4000 mm 700-3000 km Target Range

Acoustic 340e3mm/sec EM 300e3 km/sec V phase

0-300 microsecond 0-300 microsecond TDOA

~50 microsecond ~50 nanosecond SD time

~17 mm ~0.015 km SD length

http://www.technion.ac.il/~dssl/


April 13, 2023 32

Acoustic TDOA Experiment:The satellite formation is hovering on a 4 on 4 meters air table. The target is mounted 3 meters above the table and transmits 40 KHz acoustic pulses.

UltrasoundTransmitter

Satellites


April 13, 2023 33

-0.6 -0.4 -0.2 0 0.2 0.4 0.6

-0.7

-0.6

-0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

Satellite Formation Flight and Target Location on Table PlaneNominal target range is 3.089[m]

Nominal distance in formation is 0.617[m]

X [m]

Y [m

]

Target

SAT1

SAT2


April 13, 2023 34

-0.8

-0.6

-0.4

-0.2

0

0.2

-0.2-0.1

00.1

2.953

X [m]

Evolution of Location ErrorTime SD is 15[sec] Sv's Location SD is 1[cm]

Y [m]

Z [m

]

Estimation

Initial

Target


April 13, 2023 35

2 4 6 8 10 12 14 16 18 20 220

20

40

60

80

100

120Final Estimation Error is 8.32[cm] with 3= 20.9 [cm]

[cm

]

[Estimation Steps]


April 13, 2023 36

Formation Flying PDR summary

• In the first semester we selected the following principals:– 2 satellites per formation– In-plane formation– Relative control method

• This means that in worst-case scenario, ∆V required to maintain formation and altitude is

Distance Reaches Boundary

Elliptic Maneuver

Verify Successful Maneuver

Nominal State + Altitude Maintenance

V = 7.774m

sec


CDR Revisions:1. No Altitude Maintenance

– Satellites are allowed to lose altitude– Once correction is needed, the maneuvering satellite also changes its

altitude to that of its partner’s (Hohmann Transfer)– In worst-case scenario, ∆V required to maintain formation is:

April 13, 2023 37

Distance Reaches Boundary

Elliptic Maneuver

Verify Successful ManeuverNominal State Hohmann

Transfer

V = 2.52m

sec


2. Statistical Analysis• In an attempt to reduce ∆V required, we performed a statistical

analysis of the actual scenarios that may occur.– Euler angles of satellite are normally distributed ( )– 800 simulations performed

• Results:– 69% of cases –

No correction required– 31% of cases – 1 correction required– In none of the cases were 2 corrections needed

• Conclusion: ∆V needed to cover 99.99% of all cases is

April 13, 2023 38

0 , 2.5

V = 0.18m

sec


Satellite Design(Updates)

April 13, 2023 39


Structure PDR SummaryA 3U cubesat has been chosen for the satellite`s structure.Inner Components PlacementGuidelines:• Maximum distance between magnetometer and magnetic field generators

(magnetorquers, electrical components)• Center of mass should be as close to geometric center as possible • Thrust vectors should pass as close to the center of mass as possible• Patch antenna facing Nadir direction• GPS Antenna facing Zenith direction

April 13, 2023 40


CDR UpdatesTwo alternatives of the 3U cubesat were considered: Pumpkin skeleton ISIS skeleton

April 13, 2023 41


Isis structure Pumpkin structure Criteria weight

Criteria

10 8 0.5 Compatibility to ISIS ISIPOD

1 8 0.1 Flight heritage 9 7 0.4 Modularity

8.7 7.6 Total

April 13, 2023 42

Structure Selection

The Isis structure was chosen .


The satellite`s structure

April 13, 2023 43


The Satellite`s Three Major Areas

April 13, 2023 44

Middle – Propulsion SystemBottom - Electronics Top – Communications


April 13, 2023 45

Exploded View


April 13, 2023 46

A Finite Elements method is required – In order to reduce the complexity of the geometric model a simplified model was suggested:

• The inner components are referred as “Point Mass”• Three mass points simulate the three major parts• Each point mass is connected through 8 points to the satellite’s

skeleton in order to simulate the real assembly

Analysis


April 13, 2023 47

The boundary conditions are fixed support on all eight legs of the skeleton in order to simulate the satellite in the launch POD.

Modal Analysis


April 13, 2023 48

The first 6 modes are:

Modal Analysis

Frequency [Hz] Mode

695 1

708.11 2

755.18 3

756.94 4

769.25 5

769.6 6

3rd mode 2nd mode 1st mode

6th mode 5th mode 4th mode


April 13, 2023 49

A 16g load was set in the longitudinal direction and a 2.75g was set in the lateral direction. The results show the satellite will endure the launch loads even with a 10 degree misalignment with its long axis.

Static Analysis

16g10 deg


April 13, 2023 50

Static Analysis ResultsMiddle – Deformations Entire Satellite DeformationTop – Stress


April 13, 2023 51

Attitude Determination & Control SubsystemRequirements

1. Spacecraft shall be 3 axis stabilized 2. Spacecraft's long axis shall be Nadir Oriented3. Maximum pointing error (per axis):

1. Cruise Mode: less than 5⁰2. Engine Ignition: less than 10⁰

4. ADCS sub-system's mass shall be less than 190 grams5. Maximum power consumption shall be less than 630 mWatt6. Maximum time from deployment from launch pad until initial

stabilization shall be less than 24 hours


PDR Review:• Attitude control actuators: 3 magneto-torquers.• Attitude Determination: Magnetometer + Analog Sun Sensors

(preliminary design)• Hardware Selection• Disturbance torque estimation

April 13, 2023 52


Hardware Updates:

April 13, 2023 53

CDR PDR

HoneywellHMC 5843(Integrated to OBC)

BillingsleyTFM65-VQS Magneto

-meter50 milligram4x4x1.3 mm

117 gr3.51x3.23x8.26 [cm³]

Visio TorquerPCB

Satellite Services LTDTorquer rod (x3)

Magneto-Torquer

m = 100 grSize: 10 x 9 cm

Dipole = 0.5 Am²

m = 30 grL=7 cm

D=0.9 cmDipole=0.2 Am²


Analog Sun Sensor DesignCurrent to sun angle of attack relation:

is the current measured when the sun shines

directly in the normal direction:

April 13, 2023 54

z

x

y

max sinI I

maxI

90


April 13, 2023 55

CAESARCAESAR

Top View Side View

1 max 3 1

2 max 3 2

3 max 3

cos sin

cos sin

sin

I I

I I

I I

2 1sin cos

The Current-Sun AOA Relations:

Finding the AOA angles using: and Sun’s vector in Body Frame is written as:

where,

11

2

arctanI

I

1 3

2 3

3

sin cos

sin cos

sin

BsV

1

2

3

0 0

0 0

0 0

Bs

X I

V Y I

Z I

, , 1X Y Z


Attitude Determination Algorithm• Computing Sun Vector and Magnetic Vector in ECI - ,• Using sensor’s data to derive Sun Vector and Magnetic Vector in

body frame: • Finding a rotation matrix from Body Frame to ECI:

• Finally, Finding rotation matrix from Body Frame to VVLH:

• From the rotation matrix it’s easy to derive Euler angles by:

April 13, 2023 56

IsunV I

magV

,B Bsun magV V

1I I I I I B B B BB sun mag sun mag sun mag sun magC V V V V V V V V

VVLH VVLH IB I BC C C

2,3 1,3 1,2

2 23,3 1,12,3 3,3

arctan , arctan , arctanC C C

C CC C


Problem: During Eclipse sun’s Vector in body frame is unattainable.Consequence: Attitude determination of the satellite during eclipse

is unattainable.Solution: Rotational rate estimation from 3 attitude measurements,

using Lagrange interpolation formula:

April 13, 2023 57

3 2 3 1 3 1 23 1 2 3

1 2 1 3 2 1 2 3 3 1 3 2

3 2 3 1 3 1 23 1 2 3

1 2 1 3 2 1 2 3 3 1 3 2

3 2 3 1 3 1 23 1 2 3

1 2 1 3 2 1 2 3 3 1 3 2

2

2

2

t t t t t t tt t t t

t t t t t t t t t t t t






Simulation ResultsSimulation time: 2 days. Disturbance Forces: STK Default

April 13, 2023 58


April 13, 2023 59

Control Designthe control algorithm needs to deal with the following disturbances:• Gravity• Solar Pressure• Atmospheric Drag• Magnetic Field

z

x

y

g

• Engine Torque

z

x

y


April 13, 2023 60

Control Design – State-SpaceOur state-space equations will be:

While: ng – The gravity gradient disturbance moment

nd – The remain disturbances moments

m – The control dipole momentb – The magnetic field

g dI m b n n

3 2

20 1 0 1

2 3 10 2

2 20 3 0 3

2 1

0 0 0

0 0 00 0 0 1 0 0

0 0 00 0 0 0 1 0

0 0 0 0 0 1 04 0 0 0 0 1

00 3 0 0 0 0

0 0 1 0 00

x x

y y

z z

b bI I

PPb b

QQ I IRR b b

I I

1

2

3

0 0 0

0 0 0

0 0 0

1 0 0

10 0

10 0

dx

y

z

m

m nIm

I

I

0

TP Q R A


April 13, 2023 61

Control Design – Control System Topography

nd+ngg

T_ctrlu=-Kx

Y=eye*X

rates

-K-

r2d1

-K-

r2d

eule

rn

d

distubancesmoments

anglesT

uX

StateSpace

err u

PID+Anti WindUp

1

In1

in xP

Q

R

b

1

u

Scope3

Saturation1

-K-

Ks

-K-

Kp

-K-

Ki

-K-

Kd

1s

Integrator1

2

rates

1

err

Anti WindUp

PID


April 13, 2023 62

Control Design – Results

0 1000 2000 3000 4000 5000 6000 7000 8000-1

0

1

2

3

Ang

le [d

eg]

time [sec]

Steady State

0 1000 2000 3000 4000 5000 6000 7000 8000-50

0

50

100

150

200

250

ControllerST=4652.8566sec

Ang

le [

deg]

time [sec]

180o command:

0 1000 2000 3000 4000 5000 6000 7000 8000-50

0

50

100

Ang

le [

deg]

time [sec]

Eclipse Response


Propulsion SystemRequirements1. The Propulsion system shall provide orbit maintenance during satellite's

lifetime in orbit.2. The Propulsion system shall provide.3. The system shall provide thrust for keeping the satellite’s formation and

altitude. 4. The propulsion system total mass shall be < 1400 gr.5. "Green fuel" that isn’t toxic should be considered.6. The propulsion system's volume shall be less than 10x10x10.7. The propulsion system shall function normally within temperature range of

April 13, 2023 63


April 13, 2023 64

ΔV Budget ΔV[m/s]

UsageCDR PDR

9.19 9.19 Positioning

0.18 9.32 Keeping Formation

0.00 13.35 Deorbiting

0.94 6.37 Spare (10%)

10.31 38.23 Total


PDR Summary• There were 3 missions for the Propulsion System:

1. Positioning.

2. Keeping formation.

3. Deorbiting.

• We selected a warm gas propulsion system of “MicroSpace”.

• We designed an external high pressure gas tank for the propulsion system.

• The total propulsion system mass was 436 g.• The cost of the propulsion system without the

external gas tank was € 81,000.

April 13, 2023 65


Design updates since PDR• There are 2 missions for the Propulsion System:

1. Positioning.

2. Keeping formation.

• We noticed that “MicroSpace’s” propulsion system is too heavy, complicated and expensive. So, we designed a new Cold Gas Propulsion System that meet our specific requirements.

• The total propulsion system mass is 429 g.• The cost of the propulsion system is $ 7,321.

April 13, 2023 66


April 13, 2023 67

Block Diagram And Detailed Components

Solenoid Valve

Fill Valve

PressureTransducer

LatchValve

High Pressure

Tank

Solenoid Valve

Pressure Regulator

Pressure Regulator

The Propulsion System's Block Diagram

Pressure Regulator

Pressure Connector

Straight Connector

CurvePipe

Solenoid Valve

Thruster House

Thruster

Straight Pipe

Pressure Transducer

Main Connector

Latch Valve

Gas Tank

Fill Valve


Strength Analysis And Optimization

April 13, 2023 68

• In order to design the most optimal components, we made analysis with “SimulationXpress”.

• At iterative work, we fit the wall thickness to the applied pressure(at extreme conditions of 50°C), so we get the optimal weight.


5 10 15 20 25 30 35 40 45 50 55 600

0.2

0.4

0.6

0.8Thrust Vs. Pc

F [

N]

Pc [atm]

5 10 15 20 25 30 35 40 45 50 55 600

1000

2000

3000

tpulse

Vs. Pc

t puls

e [se

c]

Pc [atm]

5 10 15 20 25 30 35 40 45 50 55 600

0.2

0.4

0.6

0.8Thrust Vs. Pc

F [

N]

Pc [atm]

5 10 15 20 25 30 35 40 45 50 55 600

1000

2000

3000

tpulse

Vs. Pc

t puls

e [se

c]

Pc [atm]

April 13, 2023 69

Design Parameters OptimizationIn order to choose the most suitable design parameters we made graphs and at iterative way we gathered to the best solution.

20 40 60 80 100 120 140 160 1800.065

0.07

0.075

0.08Thrust Vs. Area Ratio

F [

N]

Area Ratio Ae/At

20 40 60 80 100 120 140 160 18065

70

75

80Isp Vs. Area Ratio

Isp

[sec

]

Area Ratio Ae/At

20 40 60 80 100 120 140 160 180360

380

400

420

tpulse Vs. Area Ratio

t puls

e [se

c]

Area Ratio Ae/At

20 40 60 80 100 120 140 160 18036

38

40

42

44

mprop Vs. Area Ratio

mpr

op [

gr]

Area Ratio Ae/At

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.10

0.5

1

1.5Thrust Vs. Throat Diameter

F [

N]

Throat Diameter [mm]

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.10

1000

2000

3000

tpulse

Vs. Throat Diameter

t puls

e [se

c]


0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.10

0.5

1

1.5Thrust Vs. Throat Diameter

F [

N]


0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.10

1000

2000

3000

tpulse

Vs. Throat Diameter

t puls

e [se

c]


20 40 60 80 100 120 140 160 1800.065

0.07

0.075

0.08Thrust Vs. Area Ratio

F [

N]

Area Ratio Ae/At

20 40 60 80 100 120 140 160 18065

70

75

80Isp Vs. Area Ratio

Isp

[sec

]

Area Ratio Ae/At

20 40 60 80 100 120 140 160 180360

380

400

420

tpulse

Vs. Area Ratio

t puls

e [se

c]

Area Ratio Ae/At

20 40 60 80 100 120 140 160 18036

38

40

42

44

mprop

Vs. Area Ratio

mpr

op [

gr]

Area Ratio Ae/At


Cold Gas Thruster - Specifications

The Cold Gas Thruster parameters:(T 278K)≅

Parameter Value

Throat diameter 0.3 mm

Exit diameter 3 mm

Thrust 75.2 mN

Isp 74.8 sec

Pc 6 atm

Pe ~0 atm

April 13, 2023 70

Parameter Value10.31 m/sec

Pulse time 370 secPropellant mass (N2) 37.6 grTank Pressure 137 atm = 2,015 Psi

The system's parameters according to our mission: ()


Programmatic Design

April 13, 2023 71


April 13, 2023 72

Cost Estimation - Propulsion example

Part NameGas

TankThruster

Pressure Regulator

Pressure Transducer

Latch Valve

Solenoid valve

Fill Valve FastenersControl Board

Part's Cost$ 303 66 1,500 885 900 233 400 452 724

Assembly Work

Time [Min] 120

Rate [$/Hour] 100

Cost$ 200

Opacity Test - Helium mass spectrometer with bell jar

Work

Time [Min] 120

Rate [$/Hour] 70

Cost$ 140

Quantity for the constellation 48 96 96 48 48 96 48 48 48

Total Cost per Satellite$ 7,268

Total Cost for Entire Constellation$ 348,856

Non-recurrent

Pressing pattern$ 10,000

Opacity Equipment$ 5,000

Environmental testing$ 30,000

Total non-recurrent cost$ 45,000


April 13, 2023 73

Cost Estimation

ComponentsRecurrent cost$

Non-recurrent cost $Cost per satellite

Cost for Entire Constellation

Satellite Structure 6,585 316,113

Propulsion System 7,268 348,856 45,000

ADCS 19,604 941,008

Payload System 13,355 641,072

Communication System 23,149 1,111,192

Power 17,368 833,664

Formation 16,600

Geolocation 35,480

Total cost 87,329 4,191,905 97,080

Constellation cost 4,288,985

Launching the entire constellation (6 launches) costs : ~ $ 67,714,464


Every project has risks –uncertainties that weren't anticipated earlier Risk Management -identifying, analyzing and responding to project risk. project risks are uncertainties that may result in schedule delays, cost overruns,performance problems, adverse environmental impacts or other undesiredimpacts.

- The likelihood of the event - The potential consequence to the project - Risk factor

April 13, 2023 74

Risk management

fP

fR P C CR


Risk PfC R-Risk Factor

Propulsion system: Safety risk-the system contains high pressure, chance of explosion.

0.8 0.8 0.64

Risk Mitigation: Performing experiments and tests on the system and particularly on the tank .

Propulsion system: Schedule risk- the launch company will not agree to launch the satellite.

0.8 0.7 0.56

Risk mitigation: Experiments and higher safety factors.

Propulsion system: Technical risk- The amount of gas might not be enough -sun storms increase drag.

0.7 0.7 0.49

Risk mitigation: Increasing the percentage of spare gas in the tank. This spare gas will be used in unexpected weather in space.

April 13, 2023 75

Risks analysis-The 7 major risks in the project

Low Risk Low Risk Low Risk BenignLow Risk Medium Risk Medium Risk MediumLow Risk Medium Risk High Risk Harsh

ConsequenceLikelihood

Very LikelyLikelyNot Likely


Risk PfC R-Risk Factor

Propulsion system: Technical risk-Center of mass wouldn't coincide with the engine`s nozzle.

0.7 0.6 0.42

Risk mitigation: Designing a new propulsion system or moving components in the satellite .Launch: Finding time windows suitable for the launch of

24 pairs of satellites in 6 different launch dates. 0.5 0.9 0.45

Risk mitigation: Communicating with launch provider in advance as possible in order to decrease the probability of such failure .

Orbits and constellation: Technical risk-Satellite collision with space debris

0.4 0.9 0.36

Risk mitigation: Running Debris assessment simulations using NASA's Debris Assessment Tool, and STK. Launching redundant (extra) satellites to account for damaged satellites.

Attitude control: Stabilization of the satellite by the attitude control system.

0.5 0.8 0.4

Risk Mitigation: Testing the satellite in a laboratory and performing simulations.

April 13, 2023 76


Number of risksfound system

8 Propulsion

2 Attitude control

1 Geolocation

2 Structure

2 launch

2 Formation Keeping

2 Orbits and constellation

4 Electrical Power

2 Communication /payload

1 Thermo control

26 Total

April 13, 2023 77

Summary:The propulsion system has the most risks in the project and the consequence of its risks is the most severe. This is understandable since the propulsion system is new and we don’t have previous experience with such systems. This means we would have to perform more experiments and tests on the system and of the system with the satellite in order to mitigate the risks.


System ReliabilityReliability: “the probability that a device will work without failure over a specific time periods or amount of usage” [IEEE, 1984].

- Success Probability, - Failure Rate , -Time PeriodSeries Reliability:

Parallel/Redundant Reliability :

April 13, 2023 78

tR e R t

B CAS A B CR R R R

B

C

A

1 1 1 1P B CAR R R R


For our mission t=2 years and is taken as constant, so reliability is computed as:Example: Propulsion Subsystem Reliability

April 13, 2023 79

2

0

yearstR e dt

Propellant Tank

Latch Valve

Pressure Regulator

Solenoid Valve

Fill Valve

Pressure Regulator

Solenoid Valve

Thruster

Thruster

Pressure Transducer

R1=0.988 R2=0.9999 R3=0.996 R4=0.9994

R5=0.9801 R6=0.992 R7=0.99

R5=0.9801 R6=0.992 R7=0.99

2

1 2 3 4 5 6 71 1 0.9638546propulsionR R R R R R R R


Mission ReliabilityIn order to calculate the mission reliability we calculated thereliability of each phase of the mission:

April 13, 2023 80

LaunchInitial

Stabilization

Satellite Positioning(Phasing)

MissionOperation

Deorbit

ADCS Computer EPS Communication PropulsionADCS Computer EPS CommunicationADCS Computer EPS Communication Propulsion PayloadDe-Orbit

MechanismEPSCommunication

0.735Mission L S Ph M DR R R R R R

0.97LR 0.94SR 0.91PhR 0.902MR 0.97DR


Summary - Compliance to Requirements:

April 13, 2023 81

MissionCompliance Result Requirement

14.74 min Revisit Time < 15 min Constellation

97% < 1 km Location Radius < 1 km Geo-Location

1-2 years De-Orbiting within 25 yearsOrbits

Available Coverage: )+60⁰) and (-60⁰)

Global Coverage between latitudes )+60⁰) and (-60⁰)

~$87K per Sat~$4.2M Total Cost-efficient Cost

System

2.56 kg Each Satellite’s mass < 10 kg Satellite’s Mass


AcknowledgmentsWe’d like to express our appreciation and gratitude to all Those who have helped us:

Prof. Pini Gurfil, Dr. David Mishne, Dr. Zvi Hominer, Dr. Avi Vershavski, Ofer Slama.

And special thanks to our supervisorJacob Herscovitz

April 13, 2023 82


April 13, 2023 83

Thank You For Listening!Any Questions?

2011 06 17

Technology

system engineeringupdates6

pdr summary system

propulsion subsystem

electric power system

satellites mass

distress signals

distress signal transmission

satellites subsystems