This document outlines the design of a satellite constellation to monitor the US-Mexico border. The constellation will consist of 16 satellites in low Earth orbit with an orbital period of 15.93 revolutions per day. Each satellite will carry an imaging payload with 1m resolution to collect data on potential threats along the border. The total cost of developing, launching and operating the constellation is estimated to be $498.7 million, meeting the budget constraint of less than $500 million. The constellation is designed to provide continuous coverage of the border region with gaps of no more than 12 minutes.

1. Satellite Final
Team Rocket
Andy Bothun
Christian Knill
Cordarryl Solomon-
Williams
Perry Chengyu Zhang
Yijie Wang

2. Overview
• Mission Objective
• Requirements/Constraint
• Orbit
• Payload
• Sizing
• Performance
• EPS
• Battery
• Communication
• Reliability
• Control Systems
• Other Conditions
• Conclusion

3. The Mission
Main: To monitor the southern
border of the U.S and alert the
Department of Homeland Security
of a potential threat in near real
time.
Secondary: To collect data on
transiting entities in order to assess
the extent and nature of relevant
potential threat activities.

4. Requirements
Resolution: 1.0 m
Wavelength Range: 0.40-4.0 μm
Orbit Size Range: 14-16 revs/day
Constellation: min 16 Satellites
Earth Latitude Coverage: +33⁰ to -
33⁰
Swath Width: Min 100 km
Point Stability: 0.5 km
Mission Duration: 10 yrs.
Constraints
Total Fund <$500M (16 sats)
Mission Ops: $16M/yr (16 sats)
Facility Location: Schriever AFB
Initial Op Capability > 50 months
Downlink Freq.: 2.2 GHz, Uplink: 2.5 GHz
Ground Station Receiver Gain: 3.0
GS Receiver Band.: 200KHz, Temp.: 20⁰C
Max Pass Time: 12 min
Orbit Maintenance ΔV: 120 m/s/yr per sat
Solar Panel Absorb: 0.3
Solar Pane; Degrade: 3%/yr
ITAR Regulated

5. Orbit
16 Target Locations along US-
Mexico Border
Schriever AFB

6. Orbit
Mean Motion: 15.93 revs/day Semimajor Axis: 6,671 km

7. Payload
Wavelength: 2 μm
Ground Resolution: 1m
Aperture Dia.: 1.34 m
Earth Angular Radius: 73.48⁰, η=64.278⁰
Swath: 100 km
Pixel Dia.: 7E-6m
Pixel #: 2843
Image Place Radius: 0.02m, Focal Length: 0.11m
10% Margin: 0.022m
Angular Field of View: 20.643⁰
0.5 km Point Stability: 20.643±0.211⁰

8. Payload Sizing
Compared To DMC
R=7.864
Linear Dim.: 6.25
Volume: 2.44.42 m^3
Mass: 6808.6 kg
Peak Power: 4863.3 W
10% Duty Cycle
Average Power: 486.33 W

9. Performance
Ground Speed: 7.421 km/s
Swaths per Second: 7421
8 bit/pixel
Data: 22744 bits/swath
10% Duty Cycle
Data Generated: 9.11E11 bits/orbit
Memory: 106 Gbytes
Pass Time of 15 min.
Data Rate: 120.7 MB/s

10. EPS Design
From Power Budget
Total Power: 1080.73 W
Eclipse Time: 2183.97 s
Daylight Time: 3216.03 s
Assume DMC Design
Power Output Needed: 2574.11 W
Power Density @ Start: 294.724 W/m^2
Assume Sun-Tracking Solar Panel
Power Den. @ 10 yrs: 266.54 W/m^2
Solar Panel Area: 9.657 m^2

11. Battery Design
Charging Cycle: 58440 cycles
Depth of Discharge: 35%
Battery Capacity Cr: 2081.377 Whr
NiH Battery w/ 50 Whr/kg Energy
Density
Mass: 41.627kg

12. EPS
Data

13. EPS
Gain of Dish Antenna
Operating @ 200 MHz
Monopole Whip Antenna
Half-power beamwidth in
degrees with varying diameters
2.2 GHz downlink
2.5 GHz uplink

14. Communication
Uplink: The frequency at which
ground is communicating with the
satellite
Downlink: When the satellite
transponder convert the signal
and send it down to the second
earth station

15. Communication - Uplink
Frequency: 2.2 GHz
Antenna Diameter: 2 m
Antenna Aperture Eff.: 0.65
Transmit Gain: 31.399 dBi
Power at Feed: 2574.11 W
EIRP: 65.505 dBW
Range: 300 km
Path Loss dBW: 148.84 dB
PFD at Satellite: -55.04 dBW/m^2
Bandwidth: 200 MHz
Uplink C/N: 92.25 dB
C/interference: 28 dB
Satellite C/intermod: 21 dB

16. Downlink
Frequency: 2.5GHz
Antenna Diameter: 2m
Antenna Eff.: 0.65
Noise Temp.: 293 K
Antenna Gain: 32.51 dBi
Antenna G/T: 7.84 dB/K
Path Loss: 149.95 dB
Downlink C/N: 33.48 dB
C/interference: 28 dB
Total Link C/N:19.36 dB

17. Reliability
R=e^(-t/MTBF)
Payloa
d
Struct
ure
Therm
al
Power TT&C ADC
S
Propulsi
on
Comm
s
R(10 yr) .557 .916 .645 .803 .803 .606 .864 .803
MTBF(y
r)
17.1 114.1 22.8 45.6 45.6 20.0 68.4 45.6
t(yr) 10 10 10 10 10 10 10 10
Table 1: 10 Year Reliabilities of
Individual Components

18. Reliability
1 Set of Each System
No Redundancies
n Sets of Each System
n-1 Redundancies
Rsystem = [1-(1-RA)^n]*
[1-(1-RB)^n]*...

19. Reliability
Idealizing Redundancies
# of
Redundancie
s
Rsyste
m
(10 yr)
P(16/16)
(%)
16/R P(16/{16/R}
)
(%)
Nadj P(16/Nadj
)
(%)
0 .0892 1.6*10^-
15
180 N/A N/A N/A
1 .5136 2.3*10^-3 32 62.98 41 95.91
2 .7977 2.69 21 76.08 24 96.13
3 .9189 25.84 18 82.48 20 98.06
4 .9672 59.14 17 89.43 18 98.01
5 .9866 80.59 17 97.86 17 97.86

20. Control Systems
Main Body:
Uniform Cube of width: 6.25 m
Mass of 6808.6 kg
Solar Panels:
Four (1.1 m X 1.1 m X .02 m) panels on
each arm
Each connection spans 2.5 cm
Area Density of solar panels is .35
kg/m^2
1 0 0
0 1 0
0 0 1
I(Main Body)=
(M*W^2)/6* w^2
+
th^2
0 0
0 w^2 + L^2 +
12*(W/2 +
L/2)^2
0
0 0 th^2 + L^2 +
12*(W/2 +
L/2)^2
I(Solar Panel)=
Mspa/12*

21. Control Systems
*From ADACS slides:
Cp - Cg = .3 m
Residual Dipole = 1.0 A m^2
Cd = 2.5
Ro atm = 1.47 *10^-13 kg/m^3
Me = 7.96*10^15 T m^3
q = .6
Fs = 1376 W/m^2
I (Satellite)=
44327.2 kg
m^2
- -
- 44430.5 kg
m^2
-
- - 44430.2 kg m^2

22. Control Systems
Gravity Disturbance Torque:
Tg max = (1.5𝛍e/R^3)*(Imax - Imin)*(1)
=2.074*10^-4 N m
Magnetic Torque:
Tm max = [dipole]*Me*(1)/R^3
=2.673*10^-5 N m
Aerodynamic Drag Torque:
v = [𝛍e/R]^.5 -= 7.73 km/s
Ta max = .5*Roatm*Cd*SA *(Cp - Cg)*v^2
=1.318*10^-4 N m
Solar Pressure Torque:
As ≈ (15 m)*(6.5 m) = 100 m^2
Tsrp = Fs*As*(1 + q)*(1)*(Cp - Cg)/c
=2.202*10^-4 N m
System will accommodate
Total Disturbance Torque:
Td=∑(Tx^2)^.5
=3.310*10^-4 N m

23. Cost Estimation
The Total Cost= Launch cost + cost of satellites
Launch Cost:
With a satellite mass of 1191.8 kg, the Athena 3 would have the
capability of bringing 3 satellites to our designed orbit, which is
quite close to LEO, for $31M per launch. Therefore with 5
launches of Athena 3, we’d deliver 15 satellites. Another one
left we could use one Taurus for launching, which would be
$20M. Therefore, the total cost of launch vehicles would be
31*5+20=$175M

24. Cost of units

25. Total Cost
With Theoretical First Unit Cost, and the following formula(with S of 95%),
we were able to calculate the costs of every single unit of the satellite, and the
total amount is added up to 323.7 Million
Therefore, the total Cost is the total operation cost (Launch vehicle cost) plus
the Cost of 16 satellites, which is 353.7M+ 175M=498.7M. We just made it
under the budget.

26. Other Cond. - Launch
• This mission is constrained to
be in operation in less than 50
months.
• The launch window for each
satellite exists nearly every
day, setting aside the final 2 + 1
months for launching.
• The final phase before launch
requires up to 3 months of
testing and final satellite
integration.
• The team has 44 months to
build these satellites, launch
vehicles, and launch sites.

27. Conclusion
Total Number of Satellites: 16
15.93 rev/day @ 300 km altitude
Coverage (24 hrs)
Target
Total
Coverage
(sec)
Longest
Break (min) Target
Total
Coverage
(sec)
Longest
Break (min)
1 52875.04 12 9 53000.27 12
2 46451.36 12 10 52505.7 12
3 46379.27 12 11 51653.87 12
4 46721.79 12 12 49387.95 12
5 46766.29 12 13 49163.29 12
6 45103.36 12 14 50051.69 12
7 44853.98 12 15 49141.85 12
8 45215.36 12 16 48080.25 12