The document summarizes a preliminary design review for a search and rescue satellite constellation. Key points: 1) The constellation would consist of 48 nano-satellites in 6 orbital planes to locate persons in distress within 15 minutes with 1 km accuracy. 2) Each satellite would have a mass of 3.11 kg, use passive thermal control, active 3-axis attitude control, a warm gas propulsion system, and communicate via two dipole antennas. 3) The design was updated since the preliminary design review, with the total satellite mass reduced to 2.3 kg and the delta-V budget reduced from 10.31 m/s to 7.94 m/s.

1. Team Members:
Yuval Porat
Moshe Sedero
Hanan Amar
Noam Leshem
Noam Leiter
Oshrat Marfogel
Ittai Cohen
Nitsan Bavli
Supervisor: :
Jacob Herscovitz
Winter 2010-2011

2. Contents:
1. Background 8. Structure
2. Requirements 9. ADCS sub-system
3. Completed PDR Design 10. Propulsion sub-system
and Summary 11. Cost Estimation
4. System engineering 12. Risk Management
5. Orbits and Constellation 13. Reliability
6. Geo-location 14. Work Breakdown Structure (WBS)
7. Formation Flying 15. Summary and Acknowledgments
6 November 2011

3. 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.
6 November 2011

4. Customer Requirements
1. 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.
6 November 2011

5. 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.
6 November 2011

6. Completed PDR Design Work:
Electric Power System:
EPS + Matching Solar Array to Battery to Consumers Max
Mass
Battery Battery Efficiency Efficiency DoD
ClydeSpace 3U EPS +
Battery Pack
90% 90% 20% 170 g
Efficiency @ 5E14 e-
Solar Panel Qty. Efficiency (BOL) Cell Area Cell Weight
/cm2
Azure TJ 3G30C 26 29.1% 30.18 cm2 2.6 g 26.5%
Thermal Control:
•Passive Control
•Steady State mean temperature: -6⁰
6 November 2011

7. GPS
Communication: Other CubeSat
User Geolocation Satellite Telemetry Main CubeSat
Segment to Satellite & Control
Satellite Ground Station Satellite Ground
Station
Antenna Monopole Patch Parabolic 2 Dipole 2 Dipole Yagi
Dipole
Transmitter 2.4Ghz 2.4Ghz -- 450Mhz 400Mhz 400Mhz
Receiver -- 2.4Ghz 2.4Ghz 450Mhz 400Mhz 400Mhz
6 November 2011 User's Beacon
Ground Station Ground Station
for GeoLocation for Control and
Broadcasting: data receiving
and handling
Telemetry
6 November 2011

8. 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
Payload
Inclination Mass capability Est.
Altitude space Accuracy Reliability
[deg] [kg] Cost
[m]
Any above D1.55 x i = 0.1 [deg]
LEO 800 to 700[km] Med $10.9M
9⁰ H1.7 Apogee = 15[km]
6 November 2011

9. PDR Summary - Mission:
Constellation : 700 km , i 45, e 0 , 48 satellites, 6 planes
Geolocation: TDOA algorithm, 97% location within 15 min,
3% location within 30 min
Formation: 2 satellites, In-plane formation, relative control,
distance = 200 ± 50 km
6 November 2011

10. 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=100
Available Average Power: 6.78 W
6 November 2011

11. System Engineering
(Updates)
6 November 2011

12. Mission Profile
ˆ
x
ˆ
y
ˆ
z
Launch and Initial Dispersion Mission De-orbiting
Deployment stabilization Operation
10 Mins 24 hours 14 days 2 years 1-2 years
6 November 2011

13. Budgets
ΔV Budget Mass Budget Power Budget
PDR CDR Power consumption [mW]
PDR CDR Sub System Consumers
Usage System Total System Total
ΔV[m/s] ΔV[m/s] Cruise Detection Maneuver
Mass [Kg] Mass [Kg]
Positioning OBDH 0.08 0.08 OBDH 200 600 600
Keeping ADCS 0.209 0.09 ADCS 430 630 630
Formation Propulsion 1.209 0.458 Propulsion 0 0 2000
Deorbiting 0
Thermal 0 0
Spare Thermal Control 0 0 0
0.94 Control
(20%) Communication 200 450 450
Communication 0.23 0.23
Total 10.31 Payload 0.105 0.105 Payload 200 450 450
GPS 0.003 0.003 GPS 200 200 200
Power 0.297 0.237 EPS 200 200 200
Structure 0.958 1.02 Structure 0 0 0
De-Orbit - 0.08 De-Orbit 0 0 0
Total 3.111 2.3 Total 1430 1880 4530
6 November 2011

14. Design Iteration:
Subsystem’s Mass: Satellite’s Mass:
Allocation for Mass [Kg] Comments
Sub-System
Sub-System Sub-System Dry Mass 2.3
Total Mass [Kg]
[Kg] 10% Margin 2.53 X+10%
Power 0.31 0.2376 Includes 10% Margin for
Fuel 0.031
ADCS 0.22 0.09 Fuel
Thermal Control 0 0 Includes:
Communication 0.265 0.23 Total 2.56 10% margin for Fuel and
Payload 0.11 0.105 10% margin for Dry Mass
GPS 0.004 0.003
OBDH 0.1 0.08
Propulsion 0.7 0.458
Structure 1.05 1.02
De-Orbit 0.2 0.08
Total 2.955 2.3
6 November 2011

15. System Hierarchy Diagram:
6 November 2011

16. Satellite Block Diagram 2 UHF
S-Band Antennas GPS
Antenna
S-Band Comm. OBDH UHF Comm. De-Orbit Device Antenna
S-Band Main UHF UHF “Nano
Receiver computer Transmitter Receiver Terminator”
GPS
Power
Attitude Determination & Control Propulsion
EPS
Pr. Tank
3x 3x
Magneto- Magneto-
Battery Torquers Meters Valves
Filter
Photo Voltaic Cells Regulator
2x Thruster
6 November 2011

17. Physical Hierarchy:
2
System interfaces ( N diagram):
6 November 2011

18. Mission Design
(Updates)
6 November 2011

19. Orbits and Constellation
PDR 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 v 13.35 sec m
6 November 2011

20. 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.
Constellation Revisit Time Vs Altitude
Revisit Time [min]
Altitude [km]
6 November 2011

21. De-Orbiting
m
• PDR calculation for de-orbiting from 700 km to 650 km: v 13.35
sec
• From 710 km to 650 km – even higher: v 16.01 m
sec
• 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.
6 November 2011

22. • NASA’s de-orbit mechanism increases satellite’s surface area, and
thus drag force, by 60%
• Device’s dimensions:
Material:
Aluminum plates
Germanium Film
28 cm
9.9 cm
9.9 cm Weight:
~200 gr (est.)
Device can be placed only on top or bottom panel
6 November 2011

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
6 November 2011

24. 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 Force
that opposes orbital motion
and produces electrodynamic drag.
 L  
F I B dl
0
6 November 2011

25. Device’s Performance:
• The extended tape’s surface area is about 152 cm², increases
spacecraft surface area by 50 %
• Deorbit Time Prediction with mechanism:
CAESAR
Satellites
6 November 2011

26. De-Orbit Mechanism Selection
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
0.2 24.4 yr 2 0.4 22.2 yr 3 0.6 <1 yr 5 1
Time
Compat-
0.3 1 0.3 3 0.9 5 1.5
ibility
Total 1.2 3 5
The nanoTerminator gives us the best deorbit time, for the lowest
additional mass, and is the easiest to integrate with the satellite.
6 November 2011

27. Geolocation
The TDOA location method
t21
1
s2 u
1
s1 u • The hyperbolic equation
c c can be transformed to a
2 2 2
si u Xi X Yi Y Zi Z quadratic form
c is the speed of light
M m u
uT Mu 2mT u m0 0 uT 1 0
mT m0 1
T
M 4 s1 s2 s1 s2 4d 2 I m 2
2 s2 s12 s1 s2 2d 2 s1 s2
2 2
m0 s2 s12 d2 2 s12 2
s2 d2 d c t21
6 November 2011

28. • 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.
Satellite Formation and Target on TDOA Hyperboloid Satellite Formation and Target on TDOA Hyperboloid
SAT1 SAT1
SAT2 SAT2
Target Target
TDOA Hyperboloid TDOA Hyperboloid
4
3
2 -4
1 -3
0
-2
Z
-1
-1
-2
0
-3
5 1
-4
2
4 5
0 3
2
4 Y
0 0
-2 -5 5
-4 -5 X
Y X
6 November 2011

29. • 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.
Target in the Intersection of a Sphere and 2 TDOA Hyperboloids
SAT11
SAT21
Target
SAT12
SAT22
8
6
4
Z
2
0
-2
5
5
0
0
-5
-5 -10
Y X
6 November 2011

30. • 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.
6 November 2011

31. 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.
Parameter Space Scale EchoLab Scale
Formation ~200 km ~500 mm
Target Range 700-3000 km 3000-4000 mm
V phase EM 300e3 km/sec Acoustic 340e3mm/sec
TDOA 0-300 microsecond 0-300 microsecond
SD time ~50 nanosecond ~50 microsecond
SD length ~0.015 km ~17 mm
6 November 2011

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.
Satellites
Ultrasound
Transmitter
6 November 2011

33. Satellite Formation Flight and Target Location on Table Plane
Nominal target range is 3.089[m]
Nominal distance in formation is 0.617[m]
0.2
0.1
0
-0.1
-0.2 Target
Y [m]
SAT1
-0.3 SAT2
-0.4
-0.5
-0.6
-0.7
-0.6 -0.4 -0.2 0 0.2 0.4 0.6
X [m]
6 November 2011

34. Evolution of Location Error
Time SD is 15[ sec] Sv's Location SD is 1[cm]
0.2
3 0
Z [m]
2.95
-0.2
0.1
0 -0.4
-0.1 -0.6
-0.2 X [m] Estimation
-0.8
Y [m] Initial
Target
6 November 2011

35. Final Estimation Error is 8.32[cm] with 3 = 20.9 [cm]
120
100
80
[cm]
60
40
20
0
2 4 6 8 10 12 14 16 18 20 22
[Estimation Steps]
6 November 2011

36. Formation Flying
PDR summary
• In the first semester we selected the following principals:
– 2 satellites per formation
– In-plane formation
– Relative control method
Nominal State Distance
Reaches Elliptic Verify Successful
+ Altitude
Boundary Maneuver Maneuver
Maintenance
• This means that in worst-case scenario, ∆V required to
m
maintain formation and altitude is V = 7.774
sec
6 November 2011

37. 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:
m
V = 2.52
sec
Elliptic
Distance Maneuver
Reaches Verify Successful
Nominal State
Boundary Hohmann Maneuver
Transfer
6 November 2011

38. 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 ( 0, 2.5 )
– 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
m
• Conclusion: ∆V needed to cover 99.99% of all cases is V = 0.18
sec
6 November 2011

39. Satellite Design
(Updates)
6 November 2011

40. Structure
PDR Summary
A 3U cubesat has been chosen for the satellite`s structure.
Inner Components Placement
Guidelines:
• 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
6 November 2011

41. CDR Updates
Two alternatives of the 3U cubesat were considered:
Pumpkin skeleton ISIS skeleton
6 November 2011

42. Structure Selection
Criteria Criteria Pumpkin structure Isis structure
weight
Compatibility 0.5 8 10
to ISIS ISIPOD
Flight heritage 0.1 8 1
Modularity 0.4 7 9
Total 7.6 8.7
The Isis structure was chosen.
6 November 2011

43. The satellite`s structure
6 November 2011

44. The Satellite`s Three Major Areas
Bottom - Electronics Middle – Propulsion System Top – Communications
6 November 2011

45. Exploded View
6 November 2011

46. Analysis
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
6 November 2011

47. Modal Analysis
The boundary conditions are fixed support on all eight legs of the
skeleton in order to simulate the satellite in the launch POD.
6 November 2011

48. Modal Analysis
1st mode 2nd mode 3rd mode
The first 6 modes are:
Mode Frequency [Hz]
1 695
2 708.11
3 755.18 4th mode 5th mode 6th mode
4 756.94
5 769.25
6 769.6
6 November 2011

49. Static Analysis
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.
16g
10 deg
6 November 2011

50. Static Analysis Results
Middle – Deformations Top – Stress Entire Satellite Deformation
6 November 2011

51. Attitude Determination & Control Subsystem
Requirements
1. Spacecraft shall be 3 axis stabilized
2. Spacecraft's long axis shall be Nadir Oriented
3. Maximum pointing error (per axis):
1. Cruise Mode: less than 5⁰
2. Engine Ignition: less than 10⁰
4. ADCS sub-system's mass shall be less than 190 grams
5. Maximum power consumption shall be less than 630 mWatt
6. Maximum time from deployment from launch pad until initial
stabilization shall be less than 24 hours
6 November 2011

52. PDR Review:
• Attitude control actuators: 3 magneto-torquers.
• Attitude Determination: Magnetometer + Analog Sun Sensors
(preliminary design)
• Hardware Selection
• Disturbance torque estimation
6 November 2011

53. Hardware Updates:
PDR CDR
Honeywell
Billingsley
HMC 5843
Magneto TFM65-VQS
(Integrated to OBC)
-meter
117 gr 50 milligram
3.51x3.23x8.26 [cm³] 4x4x1.3 mm
Satellite Services LTD Visio Torquer
Torquer rod (x3) PCB
Magneto
-Torquer m = 30 gr
m = 100 gr
L=7 cm
Size: 10 x 9 cm
D=0.9 cm
Dipole = 0.5 Am²
Dipole=0.2 Am²
6 November 2011

54. Analog Sun Sensor Design
Current to sun angle of attack relation: ˆ
y
I I max sin
ˆ
x
ˆ
z
I max is the current measured
when the sun shines directly
in the normal direction:

90
6 November 2011

55. The Current-Sun AOA Relations:
I1 I max cos 3 sin 1
I2 I max cos 3 sin 2 CAESAR
CAESAR
I3 I max sin 3
Top View Side View
I1
Finding the AOA angles using: sin 2 cos 1
and 1 arctan
I2
Sun’s vector in Body Frame is written as:
sin 1 cos 3 X 0 0 I1
VsB sin 2 cos 3  VsB 0 Y 0 I2 where, X , Y , Z 1
sin 3 0 0 Z I3
6 November 2011

56. Attitude Determination Algorithm
I
• Computing Sun Vector and Magnetic Vector in ECI - Vsun , Vmag
I
• Using sensor’s data to derive Sun Vector and Magnetic Vector in
B B
body frame: Vsun ,Vmag
• Finding a rotation matrix from Body Frame to ECI:
I I I I I B B B B 1
C B V sun V mag V sun V
mag V sun V mag V
sun V mag
• Finally, Finding rotation matrix from Body Frame to VVLH:
VVLH
CB CIVVLH CB
I
• From the rotation matrix it’s easy to derive Euler angles by:
C2,3 C1,3 C1,2
arctan , arctan , arctan
C3,3 2
C
2,3 C 2
3,3
C1,1
6 November 2011

57. 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:
 t t3 t 2 t3 t1 2t3 t1 t2
3 t1 t2 t3
t1 t2 t1 t3 t2 t1 t2 t3 t3 t1 t3 t2
 t t3 t 2 t3 t1 2t3 t1 t2
3 t1 t2 t3
t1 t2 t1 t3 t2 t1 t2 t3 t3 t1 t3 t2
t3 t 2 t3 t1 2t3 t1 t2
 t3 t1 t2 t3
t1 t2 t1 t3 t2 t1 t2 t3 t3 t1 t3 t2
6 November 2011

58. Simulation Results
Simulation time: 2 days. Disturbance Forces: STK Default
6 November 2011

59. Control Design
the control algorithm needs to deal with the following disturbances:
• Gravity • Engine Torque
• Solar Pressure
• Atmospheric Drag yˆ
• Magnetic Field ˆ
x
ˆ
z
g
6 November 2011

60. Control Design – State-Space
Our state-space equations will be:


      
P Q R
T
A

I m b ng nd 0
0 0 0 0 0 0
 0 0 0 1 0 0 0 0 0 0 0 0
 0 0 0 0 1 0 0 0 0 0 0 0
m1
 0 0 0 0 0 1 b3 b2 1 
0 m2 0 0 nd
 2 Ix Ix Ix
P 4 0 1 0 0 0 0 0 1 1 P
b3 b1 m3 1

Q 0 3 2
0 0 0 0 Q 0 0 0
0 2 Iy Iy Iy

R 0 0 2 2
1 0 0 R
0 3 0 3
b2 b1 0 0 1
0 Iz
Iz Iz
While: ng – The gravity gradient disturbance moment
nd – The remain disturbances moments
m – The control dipole moment
b – The magnetic field
6 November 2011

61. Control Design – Control System Topography
euler
distubances
moments
nd
nd+ngg -K-
Y=eye*X
T
u=-Kx r2d angles
T_ctrl X
1 err u u
In1 -K-
StateSpace
PID+
Anti WindUp r2d1 rates
-K-
Anti WindUp
Ks
PID -K-
1
 s
Integrator1
b
in x Ki Saturation1
P 1
err
-K- 1
u
Kp
Q 2 -K-
rates
R Kd
Scope3
6 November 2011

62. Control Design – Results
Steady State Eclipse Response
3 100
2
Angle [deg]
50
Angle [deg]
1
0 0
-1
0 1000 2000 3000 4000 5000 6000 7000 8000
time [sec] -50
0 1000 2000 3000 4000 5000 6000 7000 8000
time [sec]
180o command:
Controller
ST=4652.8566sec
250
200
150
Angle [deg]
100
50
0
-50
0 1000 2000 3000 4000 5000 6000 7000 8000
time [sec]
6 November 2011

63. 6 November 2011

64. ΔV Budget
ΔV[m/s]
Usage
PDR CDR
Positioning
Keeping
Formation
Deorbiting
Spare (10%)
Total
6 November 2011

65. 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.
6 November 2011

66. 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.
6 November 2011

67. The Propulsion System's Block Diagram
Block Diagram And Detailed Components
Pressure Straight
Main
Transducer Pipe
Connector
Pressure
Connector
Pressure Solenoid
Regulator Valve
Fill Pressure
Valve Regulator
Straight High
Latch
Connector Pressure
Valve
Tank
Curve Solenoid
Pipe Pressure Valve
Transducer Pressure
Solenoid Regulator
Valve
Thruster
Fill Valve House
Latch Thruster
Valve
Gas Tank
6 November 2011

68. Strength Analysis And Optimization
• 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.
6 November 2011

69. Design Parameters Optimization
In order to choose the most suitable design parameters we made graphs and at
iterative way we gathered to the best solution.
Thrust Vs. Pc Thrust Vs. Area Ratio
0.8 tpulse Vs. Area Ratio 0.08
Thrust Vs. Pc
Thrust Vs. Throat Diameter
Thrust Vs. Area Ratio
0.6
0.08 0.8
420
1.5 0.075
F [N]
F [N]
0.4
0.6 0.07
0.2
t pulse [sec]
0.0751
400
F [N]
F [N]
0 0.065
0.4
F [N]
5 10 15 20 25 30 35 40 45 50 55 60 20 40 60 80 100 120 140 160 180
Pc [atm] Area Ratio Ae/At
t pulse Vs. Pc Isp Vs. Area Ratio
0.5
380 80
3000
0.07 0.2
75
tpulse [sec]
Isp [sec]
2000
0 0
1000 360
0.065 0.1 5
0.2 10
0.3 15
0.4 20
0.5 25 0.6 30 0.7 350.8 40 0.9 45 1 50 1.155
70 60
20
20 40
40 60
60 80 100[atm]120 140 160 180
Throat Diameter [mm] 140
80 100
Pc 120 160 180
Area Ratio Ae/At
tpulse Vs.tRatio Vs. Pc
60 Area Ae/At
0 65
pulse Diameter
Throat
5 10 15 20 25 30 35 40 45 50 55 20 40 60 80 100 120 140 160 180
Pc [atm]
mIsp Vs. Area Ratio
prop
Vs. Area Ratio Area Ratio Ae/At
3000
3000
80
44 Thrust Vs. Throat Diameter tpulse Vs. Area Ratio
1.5
420
tpulse [sec]
2000
[sec]
1 2000
42
75 t pulse [sec]
mtprop [gr]
400
Isp [sec]
F [N]
pulse
0.5
40 1000
380
1000
0
0.1 0.2 0.3
70 0.5
0.4 0.6 0.7 0.8 0.9 1 1.1 360
20 40 60 80 100 120 140 160 180
38 Throat Diameter [mm]
Area Ratio Ae/At
0 0
tpulse Vs. Throat Diameter
mprop Vs. Area Ratio
3000 5 10 15 20
0.5 25 0.6 30 0.7 350.8 40 0.9 45 1 50 1.155 60
36 65 0.1 0.2 0.3 0.4 44
20
20 40
40 60
60 80 100[atm]
Pc 120
Throat Diameter 120
80 100 140
[mm] 140 160
160 180
180
tpulse [sec]
2000 42
mprop [gr]
Area Ratio Ae/At
Area Ratio Ae/At
40
1000
38
6 November 2011 0
0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8
Throat Diameter [mm]
0.9 1 1.1 36
20 40 60 80 100 120 140 160 180
Area Ratio Ae/At

70. Cold Gas Thruster - Specifications
The Cold Gas Thruster parameters:
(T≅278K)
Parameter Value Parameter Value
Throat diameter 0.3 mm 10.31 m/sec
Exit diameter 3 mm Pulse time 370 sec
Propellant mass (N2) 37.6 gr
Thrust 75.2 mN
Tank Pressure 137 atm = 2,015 Psi
Isp 74.8 sec
Pc 6 atm
Pe ~0 atm
6 November 2011

71. Programmatic Design
6 November 2011

72. Cost Estimation - Propulsion example
Gas Pressure Pressure Latch Solenoid Control
Part Name Thruster Fill Valve Fasteners
Tank Regulator Transducer Valve valve Board
Part's Cost $ 303 66 1,500 885 900 233 400 452 724
Time [Min] 120
Rate
Assembly Work 100
[$/Hour]
Cost $ 200
Opacity Test - Time [Min] 120
Helium mass Rate
Work 70
spectrometer [$/Hour]
with bell jar 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
Pressing pattern $ 10,000
Non- Opacity Equipment $ 5,000
recurrent Environmental
30,000
testing $
Total non-recurrent cost $ 45,000
6 November 2011

73. Cost Estimation
Recurrent cost $
Components Cost for Entire Non-recurrent cost $
Cost per satellite
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
6 November 2011

74. Risk management
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 undesired
impacts.
Pf - The likelihood of the event
C - The potential consequence to the project
R - Risk factor R P C
f
6 November 2011

75. Risks analysis-The 7 major risks in the project
Likelihood
Not Likely Likely Very Likely Consequence
Low Risk Low Risk Low Risk Benign
Low Risk Medium Risk Medium Risk Medium
Low Risk Medium Risk High Risk Harsh
Risk Pf C R-Risk Factor
Propulsion system: Safety risk-the system contains 0.8 0.8
high pressure, chance of explosion.
0.64
Risk Mitigation: Performing experiments and tests on the system and particularly on the tank.
Propulsion system: Schedule risk- the launch 0.8 0.7 0.56
company will not agree to launch the satellite.
Risk mitigation: Experiments and higher safety factors.
Propulsion system: Technical risk- The amount of 0.7 0.7 0.49
gas might not be enough -sun storms increase drag.
Risk mitigation: Increasing the percentage of spare gas in the tank. This spare gas will be used in
unexpected weather in space.
6 November 2011

76. Risk Pf C R-Risk Factor
Propulsion system: Technical risk-Center of mass 0.7 0.6
wouldn't coincide with the engine`s nozzle. 0.42
Risk mitigation: Designing a new propulsion system or moving components in the satellite.
Launch: Finding time windows suitable for the launch of 0.5 0.9
24 pairs of satellites in 6 different launch dates.
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 0.4 0.9
with space debris
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 0.5 0.8
attitude control system.
0.4
Risk Mitigation: Testing the satellite in a laboratory and performing simulations.
6 November 2011

77. Summary: Number of risks
system
The propulsion system has the found
most risks in the project and Propulsion 8
the consequence of its risks is Attitude control 2
the most severe. This is Geolocation 1
understandable since the Structure 2
launch 2
propulsion system is new and
Formation 2
we don’t have previous Keeping
experience with such systems. Orbits and 2
This means we would have to constellation
Electrical Power 4
perform more experiments
Communication 2
and tests on the system and of /payload
the system with the satellite Thermo control 1
in order to mitigate the risks. Total
6 November 2011

78. System Reliability
Reliability: “the probability that a device will work without failure
over a specific time periods or amount of usage” *IEEE, 1984].
R e t
R - Success Probability, - Failure Rate , t -Time Period
Series Reliability:
A B C RS RA RB RC
Parallel/Redundant Reliability :
A
B RP 1 1 R A
1 RB 1 RC
C
6 November 2011

79. For our mission t=2 years and is taken as constant,
2 years
so reliability is computed as: R e dt t
0
Example: Propulsion Subsystem Reliability
R5=0.9801 R6=0.992 R7=0.99
Pressure Solenoid
Thruster
Regulator Valve
Pressure Propellant Latch
Fill Valve
Transducer Tank Valve
R1=0.988 R2=0.9999 R3=0.996 R4=0.9994
Pressure Solenoid
Thruster
Regulator Valve
R5=0.9801 R6=0.992 R7=0.99
2
Rpropulsion R1 R2 R3 R4 1 1 R5 R6 R7 0.9638546
6 November 2011

80. Mission Reliability
In order to calculate the mission reliability we calculated the
reliability of each phase of the mission:
Satellite
Initial Mission
RLLaunch
0.97 R 0.94
Stabilization
S
RPositioning
Ph 0.91 ROperation
M
0.902 RDDeorbit
0.97
(Phasing)
De-Orbit
ADCS ADCS Computer
ADCS Computer Computer EPS
EPS EPS Communication Communication
Communication Communication EPS
Propulsion Propulsion
Payload
Mechanism
RMission RL RS RPh RM RD 0.735
6 November 2011

81. Summary - Compliance to Requirements:
Mission
Requirement Result Compliance
Constellation Revisit Time < 15 min 14.74 min 
Geo-Location Location Radius < 1 km 97% < 1 km 
De-Orbiting within 25 years 1-2 years 
Orbits Global Coverage between Available Coverage:
latitudes +60⁰) and (-60⁰) +60⁰) and (-60⁰) 
~$87K per Sat
Cost Cost-efficient
~$4.2M Total 
System
Satellite’s Mass Each Satellite’s mass < 10 kg 2.56 kg 
6 November 2011

82. Acknowledgments
We’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 supervisor
Jacob Herscovitz
6 November 2011

83. 6 November 2011