UNP Mission Design Course: Extracted Notes

= Don’t forget to round-up the Number of Packets to a whole number Determined by Software Architecture

=== Slide 137: Data Budget Example, ShipSat (One Day)

Data Budget Example, ShipSat (One Day) Source Sample Data (Bytes) Sample Overhead (Bytes) Freq (Hz) Duration (s) Collections / day Sample Total (Bytes) Number of Packets (1024 B packets) Packet Overhead (Bytes) Downlink Overhead (Bytes) Total (Bytes) Payload 20,000,000 18 1 1 5 100,000,090 97657 33 3,222,681 103,222,771 Event GNC for image location correlation 200 18 10 10 5 109,000 107 3,531 112,531 SOH total from satellite 400 18 .05 86,400 1 1,805,760 1764 58,212 1,863,972 Total 101,914,850 99,528 3,284,424 105,199,274 Packet header/footer: 11 bytes per packet Radio header/footer: 22 bytes per packet = ( )

=== Slide 138: Data Budget Example, ShipSat (One Day)

Data Budget Example, ShipSat (One Day) Source Sample Data (Bytes) Sample Overhead (Bytes) Freq (Hz) Duration (s) Collections / day Sample Total (Bytes) Number of Packets (1024 B packets) Packet Overhead (Bytes) Downlink Overhead (Bytes) Total (Bytes) Payload 20,000,000 18 1 1 5 100,000,090 97657 33 3,222,681 103,222,771 Event GNC for image location correlation 200 18 10 10 5 109,000 107 3,531 112,531 SOH total from satellite 400 18 .05 86,400 1 1,805,760 1764 58,212 1,863,972 Total 101,914,850 99,528 3,284,424 105,199,274 =

+

=== Slide 139: Data Budget Example, ShipSat (One Day)

Data Budget Example, ShipSat (One Day) Source Sample Data (Bytes) Sample Overhead (Bytes) Freq (Hz) Duration (s) Collections / day Sample Total (Bytes) Number of Packets (1024 B packets) Packet Overhead (Bytes) Downlink Overhead (Bytes) Total (Bytes) Payload 20,000,000 18 1 1 5 100,000,090 97,657 33 3,222,681 103,222,771 Event GNC for image location correlation 4 18 10 10 5 20,000 20 33 645 20,644 SOH total from satellite 16 18 .05 86,400 4,320 2,172,960 2,122 33 70,027 2,242,987 Total 102,193,050 99,799 3,293,331 105,486,381

=== Slide 140: Downlink Data Budget Assessment

Downlink Data Budget Assessment 147 𝑝𝑎𝑠𝑠𝑒𝑠 = 87,905 𝑠𝑒𝑐 = 105,486,381 𝑏𝑦𝑡𝑒𝑠 ∗ 8 𝑏𝑖𝑡𝑠 𝑏𝑦𝑡𝑒 𝑏𝑖𝑡𝑠 𝑠𝑒𝑐 Knowns (per day): * Total Data: 105,486,381 bytes * Radio capability: 9600 baud (bits/s) * From orbit and ground stations: 600 sec per contact (or pass) Understand required time for downlink … Well, that won’t work… We probably won’t have 147 passes a day…

=== Slide 141: A Note on Number of Passes A Day

• On average, a LEO spacecraft will have < 8 passes in a day

• (Left) shows a 6 pass example

• 3 ascending passes

• About 90 minutes apart

• Elevation transitions from Low (1), High (2), Low (3)

• 3 descending passes

• About 10 hours after the first 3

• Elevation transitions from Low (4), High (5), Low (6) A Note on Number of Passes A Day

=== Slide 142: What do I do when data budget doesn’t close?

147 𝑝𝑎𝑠𝑠𝑒𝑠 = 87,905 𝑠𝑒𝑐 = 105,486,381 𝑏𝑦𝑡𝑒𝑠 ∗ 8 𝑏𝑖𝑡𝑠 𝑏𝑦𝑡𝑒 𝑏𝑖𝑡𝑠 𝑠𝑒𝑐 What do I do when data budget doesn’t close? How to fix: * Alter experiment timeline/CONOPS * Take more days to downlink payload data * Will also have more SOH data * Will reduce experiment opportunities * Truncate/compress data * Buy better radio * Get more ground station passes Microsoft PowerPoint stock image For ShipSat, the data volume can’t necessarily be compressed enough. Further, there is too much data to add ground stations CONOPS is not an option here due to stakeholder requirements. That leaves us with using a higher data rate radio.

=== Slide 143: Rerun our Downlink Data Budget

• Total amount of data remains the same

• We want to select a higher rate downlink radio (that doesn’t necessarily require a higher uplink radio)

• This directly impacts the link budget.

• Radio Trade

• UHF (assumed initial solution) = 9600 bps

• S-band = ~1 Mbps

• X-band = ~100 Mbps Rerun our Downlink Data Budget 1-2 pass = 843.9 𝑠𝑒𝑐 = 105,486,381 𝑏𝑦𝑡𝑒𝑠∗8 𝑏𝑖𝑡𝑠 𝑏𝑦𝑡𝑒 1,000,000 𝑏𝑖𝑡𝑠 𝑠𝑒𝑐 8.439 𝑠𝑒𝑐 = 105,486,381 𝑏𝑦𝑡𝑒𝑠∗8 𝑏𝑖𝑡𝑠 𝑏𝑦𝑡𝑒 100,000,000 𝑏𝑖𝑡𝑠 𝑠𝑒𝑐 S-band Transmitter X-band Transmitter The mission could be achieved in either S-band or X-band. There are more S-band radio options, more S-band ground stations, typically less SWAP, and S-band options with both uplink and downlink.

=== Slide 144: Radio selection impacts

• Now that we meet our data budget, we should assess the rest of the system…

• After a radio trade of available options, you would want to select the radio that increases the data rate, still meets notional size, weight, and power (SWAP) needs. Radio selection impacts Product Max Data Rate Modulation Transmit Power Frequency Size Power draw S-band Transceiver 1.5Mbps QPSK, OQPSK 2W 2200 – 2290 MHz (Tx) 2025 – 2110 MHz (Rx) 0.5 U 6W SWAP OK? Impacts link budget, check there Does it license? (assume yes, here) Fits! Assess in power budget

=== Slide 145: Storage assessment

• While we now believe we can downlink an entire day’s data in a single pass, we still need to assess storage.

• You don’t need to keep the entire mission’s data on the satellite forever.

• Discuss in the context of CONOPS

• Assess how many days of data remain relevant to download

• Estimate total storage volume for flight software, etc.

• Know how much storage you need and if it’s sufficient with selected hardware.

• For this mission, image data volume vastly outweighs flight software or other telemetry storage volume.

• A few days of stored images are probably sufficient. Storage assessment Data capacity Data in Data out

=== Slide 146: Storage over Time

UHF Radio (9600 bps) S-Band Radio (~1 Mbps) Storage over Time Compares data storage in one day with UHF radio and S-Band radio

=== Slide 147: Data Budget for Mission Design Course

• Similar to ShipSat example

• Data/telemetry collection and storage

• Amounts

• Rates

• Reasonable overhead assumptions

• Downlink impacts/feasibility Data Budget for Mission Design Course Ensure data doesn’t overflow

=== Slide 148: Pointing Budget

Pointing Budget

=== Slide 149: What is a Pointing Budget?

What is a Pointing Budget? Mission objectives and requirements * Slew over a ground station and keep antenna pointed * Image an object in space * Point along a magnetic field line ADCS capability GPS capability Mission geometry System error Satellite properties Orbital mechanics Thermal effects Assembly alignment Can the system meet mission objectives? Two methods: * Determine requirements and allocate error to each subsystem * Combine errors to determine capability

=== Slide 150: ShipSat Objective

ShipSat Objective Objective: Count ships in pier, understand occupancy patterns Microsoft PowerPoint stock image Microsoft PowerPoint stock image

=== Slide 151: ShipSat Example

ShipSat Example Goal: Image as many ports/piers as possible; 20 second dwell provides needed data product Detroit is ~450 km out of orbit track. 600 km SSO meets 45° max viewing angle Note: This example makes some assumptions which, for a real mission, must come from analysis (such as the 45° viewing angle). All numbers used are approximate but lack the accuracy/precision required for a real mission. 0.9° Full field of view 45° Max viewing angle to avoid obstructions Zenith reference Orbit track

=== Slide 152: Geometry

Geometry Maximum viewing angle (45°) Minimum viewing angle (0°) 600 km altitude Don’t forget the earth’s curvature. The angle in pink is what we care about, not the angle in blue 45° km Earth radius (6371 km) Not 45°

=== Slide 153: Geometry

Geometry 45° (x2) km Earth radius (6371 km) 135° 4.7° 40.3° 94.7° 522.6 km 523.8 km Working the geometry: 1. Find satellite relative look angle using the law of sines: 2. Complete triangles: 3. Find satellite line-of-sight distance using the law of sines 4. Find surface distance to satellite nadir point: 40.3° = sin−1 6371𝑘𝑚 ∗ sin 135° 6371𝑘𝑚 + 600𝑘𝑚 94.7° = 180° − 40.3° − 45° 4.7° = 180° − 40.3° − 135° 807.78𝑘𝑚 = (6371𝑘𝑚 + 600𝑘𝑚) ∗ sin 4.7° sin 135° 523.8𝑘𝑚 = 6371𝑘𝑚 ∗ tan 4.7° 522.6𝑘𝑚 = 4.7° ∗ 𝜋 180° ∗ 6371𝑘𝑚

=== Slide 154: Required Error

Required Error Maximum viewing angle (45°) Minimum viewing angle (0°) 0.9° 0.9° To correlate the ships in the images with known ports, we need to know the position of the image within some margin. The allowable position knowledge error must come from mission objectives and analysis, however, for this presentation we will assume 50m Port size: 100m Allowable boresight error: 50m in the plane of the ground 600 km altitude Image credits: iconsdb.com

=== Slide 155: Source of Error

• Sensor error

• Accuracy of hardware such as a star tracker or GPS

• Provided in data sheet

• Estimator error

• The software combining the sensor measurements will not provide perfect results

• Statistics-based, most common algorithms follow a normal distribution

• Controller error

• The control laws and actuators will not be perfect

• Typically, about 10 times greater than estimator error

• Mechanical system alignment

• The CAD may be perfect, but the real satellite will not have everything perfectly aligned

• This value may be measurable and fixed

• Thermal gradients

• Sun facing side of the satellite swells, altering the mechanical alignment Source of Error Most error sources are modeled as a normal distribution. Errors are typically reported to 1σ or 3σ values. This means the system will perform within the stated error either 68.27% or 99.73% of the time respectively. The level of confidence required depends on the mission

=== Slide 156: Case 1: Overhead

Case 1: Overhead Port size: 100m Allowable boresight error: 50m 600 km altitude View width: 19 km 0.9° Error Source Error Type Error Value Resulting Boresight Ground Error GPS Position 10m 10.0m Mechanical alignment Angle 1 arcsec 2.96m ADCS Angle 8 arcsec 23.66m Thermal Angle 3 arcsec 8.87m Total 45.49m Assuming our values are accurate, the overhead case will work! Minimum viewing angle (0°)

=== Slide 157: Case 2: Side Angle

Case 2: Side Angle Error Source Error Type Error Value Resulting Boresight Error in Normal Plane Resulting Boresight Ground Error GPS Position 10m 7.1 10.0 Mechanical alignment Angle 1 arcsec 3.9m 5.5 ADCS Angle 8 arcsec 31.6m 44.7 Thermal Angle 3 arcsec 11.8m 16.7 Total 76.9 Side angle is not acceptable Port Size: 100m 600 km altitude View width: 36 km 0.9° Maximum viewing angle (45°) Allowable boresight error on ground: 50m 45° 807 km range

=== Slide 158: Additional Considerations

Additional Considerations Wider view will increase ground sample distance, optics system requirements must flow from worst case Options to help meet requirements: * Improved ADCS and GPS * Possibly only need better knowledge, not control * Active thermal control and/or altering CONOPS to only take pictures when the satellite temperature gradients are minimized * Reduce requirements: must be permitted by stakeholders * What quality of data is actually needed? * What quality of data can be achieved with reduced pointing requirements? * Change/improve image processing * Possibly feature match images to reduce pointing knowledge requirements * Work geometry backwards to determine maximum look angle and alter CONOPS Error Source Error Type Error Value Resulting Boresight Error in Normal Plane Resulting Boresight Ground Error GPS Position 10m 7.1 10.0 Mechanical alignment Angle 1 arcsec 3.9m 5.5 ADCS Angle 8 arcsec 31.6m 44.7 Thermal Angle 3 arcsec 11.8m 16.7 Total 76.9

=== Slide 159: Momentum Budget

Momentum Budget We determined the quasi-static pointing capabilities, what about the dynamics? 7.5 km/s ground speed 2200 km from start of first imaging opportunity to end of last: 290 seconds Slews and tracks Norfolk for 20 seconds Viewing Detroit Slews to prepare for Norfolk Slews and tracks Charleston for 20 seconds 870 km or 116 seconds

=== Slide 160: Satellite Properties

Satellite Properties Spacecraft Dynamics Spacecraft max slew rate: Spacecraft max angular acceleration: 8.7° 𝑠 = 0.15𝑟𝑎𝑑 𝑠 = 0.047𝑁𝑚𝑠 0.31𝑘𝑔𝑚2 15.0° 𝑠 = 0.26𝑟𝑎𝑑 𝑠 = 0.047𝑁𝑚𝑠 0.18𝑘𝑔𝑚2 2.2° 𝑠2 = 0.039𝑟𝑎𝑑 𝑠2 = 0.007𝑁𝑚 0.18𝑘𝑔𝑚2 1.3° 𝑠2 = 0.023𝑟𝑎𝑑 𝑠2 = 0.007𝑁𝑚 0.31𝑘𝑔𝑚2 0.18 kg-m2 0.31 kg-m2 0.31 kg-m2 Momentum storage: 0.047 N-m-s Torque: 0.007 N-m Max Speed: 4500 RPM Moment of Inertia: 0.0001 kg-m2 Note: Lens is on small face of satellite, so most slewing will be about high moment of inertia (MOI) axes www.Isispace.nl www.bluecanyontech.com From data sheet

=== Slide 161: Tracking a City

Tracking a City 49.7° 90.0° Case 1: Start (or end) of pass, minimum angular velocity Case 2: Middle of pass, maximum angular velocity km 49.7° Nadir Reference 40.3° 40.3° 7.5 km/s Velocity perpendicular to instantaneous center Look angle from nadir

=== Slide 162: Tracking a City

Tracking a City -60 -40 -20 -600 -500 -400 -300 -200 -100 0 100 200 300 400 500 600 Look angle from nadir (deg) 100x Angular velocity (deg/s) Range (km) Ground Distance From Target (km) Range Look Angle From Nadir Satellite Angular Velocity *100 Nadir Velocity Angular Velocity

=== Slide 163: Tracking a City

Tracking a City Using Instantaneous Centers: 1. Find component of velocity perpendicular to instantaneous center 2. Find angular velocity using instantaneous center method: 49.7° Nadir Reference 40.3° 40.3° 7.5 km/s 5.7𝑘𝑚 𝑠 = 7.5𝑘𝑚 𝑠 ∗ cos 40.3° Max Angle Case: Overhead Case: 7.5𝑘𝑚 𝑠 0.4° 𝑠 = 0.007𝑟𝑎𝑑 𝑠 = ൗ 5.7𝑘𝑚 𝑠 807.8𝑘𝑚 0.72° 𝑠 = 0.0125𝑟𝑎𝑑 𝑠 = ൗ 7.5𝑘𝑚 𝑠 600𝑘𝑚 90.0° 7.5 km/s km Max Slew Rate High MOI axes 8.7°/s 0.15 rad/s Low MOI axis 15.0°/s 0.26 rad/s Satellite Capabilities

=== Slide 164: Worst Case Slew Between Cities

Worst Case Slew Between Cities Detroit Boston km 77.9° Boston Detroit Max ground distance from pointing budget geometry Slew from Boston to Detroit is largest angle change in shortest amount of time 150 km of orbit track is needed for each city: (20 second dwell at 7.5 km/s) Top view Side view

=== Slide 165: Worst Case Slew Between Cities

Worst Case Slew Between Cities Max Slew Rate Max Angular Accel. High MOI axes 8.7°/s 0.15 rad/s 1.3°/s 0.023 rad/s2 Low MOI axis 15.0°/s 0.26 rad/s 2.2°/s 0.039 rad/s2 Satellite Capabilities Spacecraft Dynamics 1. Time to slew 77.9° (static to static) using constant acceleration equation and bang-bang maneuver: 2. Checking the maximum slew rate reached via this bang-bang method shows that the max slew rate has been exceeded: 3. Using a bang-coast-bang method. First determine how long it takes to reach maximum angular velocity and the angle slewed in this time: 4. Determine the angle slewed while coasting and the time period of the coast 5. Total slew time: 15.4𝑠 = 2 ∗ 2 ∗ 77.9° 2 ∗ 𝜋𝑟𝑎𝑑 180° ൗ 0.023𝑟𝑎𝑑 𝑠2 10.3° 𝑠 = 0.18𝑟𝑎𝑑 𝑠 = 15.4𝑠 ∗ 0.023𝑟𝑎𝑑 𝑠 𝜃 = 𝜔0𝑡
𝛼𝑡2 6.5𝑠 = ൗ 0.15𝑟𝑎𝑑 𝑠 ൗ 0.023𝑟𝑎𝑑 𝑠2 0.49𝑟𝑎𝑑 = ∗ 0.023𝑟𝑎𝑑 𝑠 ∗ 6.5𝑠 2 0.38𝑟𝑎𝑑 = 77.9° ∗ 𝜋𝑟𝑎𝑑 180° − 2 ∗ 0.49𝑟𝑎𝑑 2.5𝑠 = 0.38𝑟𝑎𝑑 ൗ 0.15𝑟𝑎𝑑 𝑠 15.5𝑠 = 2.5𝑠 + 2 ∗ 6.5𝑠

=== Slide 166: Putting it All Together

Putting it All Together Maneuver Time Duration (s) Track Length (km) Image Boston 20 150 Slew to Detroit 15.5 116.25 Image Detroit 20 150 Slew to Norfolk 15 112.5 Image Norfolk 20 150 Slew to Charleston 6 45 Image Charleston 20 150 Slew to Pensacola 10 75 Image Pensacola 20 150 Slew to Miami 13 97.5 Image Miami 20 150 Total 179.5 1346.25 Available 290 2200 7.5 km/s ground speed Mission sequence is possible based on expected maneuver spacing Imaging: Maneuver:

=== Slide 167: Additional Considerations

• Saturation

• When a reaction wheel is spinning at its maximum speed it is considered “saturated” It cannot store any additional momentum until it is desaturated via magnetorquers or thrusters

• CONOPS must consider saturation for successful operations

• Wheel biasing

• When the reaction wheels switch from spinning one direction to the other, the zero crossing causes jitter. This jitter can cause blurry images or other unwanted effects. One option is to bias the reaction wheels, so they do not rest at zero but at some higher RPM. This allows them to dump momentum into the satellite body during a maneuver without changing wheel direction. Typically, biasing is performed in the motor controllers, not the main control algorithm

• Space environment impacts (drag, radiation pressure, Earth oblateness, etc.)

• There will always be small disturbance torques on the satellite due to the shape of the satellite and the attitude. Over time, the satellite will begin to tumble if not actively controlled. If controlled, the reaction wheels will begin to saturate

• Propulsion

• Firing a propulsion system will likely create disturbance torque on the satellite. Utilization of propulsion must consider wheel saturation and other ADCS/CONOPS dependencies Additional Considerations

=== Slide 168: Real Life

Real Life Euler’s Equations Recommended Texts: * Space Mission Analysis and Design (Wertz, Larson, etc.) * An Introduction to the Mathematics and Methods of Astrodynamics (Battin) * Analytical Mechanics of Space Systems (Schaub, Junkins) * Fundamentals of Spacecraft Attitude Determination and Control (Markley, Crassidis) * Fundamentals of Astrodynamics and Applications (Vallado) The geometry, significant figures, and dynamics in this presentation were simplified for ease of understanding. For a real mission, the calculations described would provide only a “napkin math” feasibility study. Euler’s equations describe the coupled, 3-axis rotation seen in real applications: 1. Equation for angular momentum: 2. Evaluating the time rate of change of the body-frame moment of inertia matrix in an inertial frame is difficult, instead, apply the transport theorem to the second term: 3. Equation (2) then becomes: 4. Adding external torques gives Euler’s Equations in vector-matrix form: 𝐻 = 𝐼 𝜔 ሶ 𝐻 = 𝐼 ሶ 𝜔 Τ 𝐵 𝐼 + 𝜔 Τ 𝐵 𝐼 × 𝐼 𝜔 Τ 𝐵 𝐼 𝐼 𝑑 𝑑𝑡 = 𝐵 𝑑 𝑑𝑡 + 𝜔 Τ 𝐵 𝐼 × ሶ 𝐻 = 𝑇 = 𝐼 ሶ 𝜔 Τ 𝐵 𝐼 + 𝜔 Τ 𝐵 𝐼 × 𝐼 𝜔 Τ 𝐵 𝐼

=== Slide 169: Pointing Budget for Mission Design Course

• 1-D quasi-static pointing and slew analysis

• Worst-case conditions Pointing Budget for Mission Design Course 49.7° Nadir Reference 40.3° 40.3° 7.5 km/s Velocity perpendicular to instantaneous center 49.7° 90.0° Ensure necessary pointing and slewing can be achieved

=== Slide 170: Link Budget

Link Budget

=== Slide 171: Communications System Refresher

• Communications system responsible for connection between satellite and ground

• TT&C (telemetry, tracking, and command/control) critical to satellite

• System includes

• Satellite radio

• Satellite antenna(s)

• Ground radio

• Ground antenna(s)

• Other hardware necessary for transmit/receive functionality

• Amplifiers

• Filters

• RF cables

• Link budget helps ensure radio link is robust Communications System Refresher

=== Slide 172: Common Frequencies, Data Rates, and Traits

IEEE Frequency Band Frequency Range Rough Data Rate Range VHF 30 – 300 MHz 9.6 – ~30 kbps UHF 300 – 1000 MHz 9.6 – 38.4 kbps L 1 – 2 GHz 0.1 – 10 Mbps S 2 – 4 GHz 0.1 – 20 Mbps C 4 – 8 GHz 10 – 100 Mbps X 8 – 12 GHz 10 – 150 Mbps Ku 12 – 18 GHz 25 – 1000 Mbps K 18 – 27 GHz 25 – 1000 Mbps Ka 27 – 40 GHz 25 – 1000 Mbps Common Frequencies, Data Rates, and Traits * Lower data throughput * Narrower spectrum bands * Larger antenna size/wider beamwidth * Lower susceptibility to rain * Cheaper, simpler hardware * More power efficient hardware * Higher data throughput * Wider spectrum bands * Smaller antenna size/narrower beamwidth * Higher susceptibility to rain * More expensive, more complex hardware * Less power efficient hardware

=== Slide 173: Communication System Analogy

Communication System Analogy Bulb/LED transmits light, lens directs the energy into free space. Some power lost as heat Power emitted is not isotropic creating a main lobe, side lobes, and weak/zero points (nulls), as well as uneven power density throughout lobes. Note, the lobes and nulls displayed are realistic for RF systems but excessive for optics. Power density decreases with distance. A farther receiver must be more sensitive to see signal Close and directly in front of maximum output. Strong signal Signal weakens as function of distance squared Beams usually aren’t perfect. While diagram is drastic, there may be off-center spots that are dimmer/brighter. In RF systems these would be side lobes Less energy emitted in some directions. In RF systems these would be angles of no energy transmission known as nulls

=== Slide 174: General Satellite Communications Path

General Satellite Communications Path Does enough signal reach the receiver? Transmitter Creates RF signal from digital input Losses Lines, connectors, reflections, mismatches, etc. Power Amplifier Boosts RF signal Satellite Antenna Sends signal into free space Downlink shown, uplink similar. General configuration shown, many possible implementations Receiver Creates RF signal from digital input Low Noise Amplifier Boosts weak received signal Losses Lines, connectors, reflections, mismatches, etc. Antenna Collects signal from free space Losses Free space, atmosphere, rain, reflections, etc. Ground Station Radio Losses Lines, connectors, reflections, mismatches, etc. Losses Lines, connectors, reflections, mismatches, etc.

=== Slide 175: Link Budget Considerations

• Radio typically includes:

• Transmitter (Tx): Creates a weak radio-frequency (RF) signal

• Usually from a digital input

• Power output and other parameters found in data sheet

• Receiver (Rx): Collects a weak radio-frequency (RF) signal

• Usually provides digital output

• Sensitivity and other parameters found in data sheet

• Amplifier: Increases signal strength

• Data sheet likely provides radio inputs/outputs, rather than individual components

• Losses reduce signal strength. Values come from:

• Physics (free space path loss)

• Hardware data sheets (estimated line loss)

• Implementation (many considerations)

• Antennas vary greatly. Gains/losses come from:

• Data sheet (expected gain pattern)

• Implementation (real gain pattern, polarization loss)

• CONOPS (pointing and cross polarization) Link Budget Considerations

=== Slide 176: Link Budget Considerations: Antenna Gain

Link Budget Considerations: Antenna Gain Larger dish or array Maximum gain -3 dB -6 dB Antenna beamwidth (angle) measured 3 dB down from peak General trend: gain depends on multiple factors. Relative sizes and angles not to scale Higher directionality and gain Lower directionality and gain Smaller dish or array Yagi Monopole/dipole Patch * For a given frequency, a larger aperture provides higher directionality (gain) * For a given aperture, a higher frequency provides higher directionality (gain) * Physical size of antenna elements is tuned to the wavelength (not directly applicable to reflector elements such as dishes, but is applicable to the feed for the dish) * Common trends: * S-band and higher: patch or patch array on satellite, dish on ground * UHF and lower: monopole/dipole on satellite, yagi on ground

=== Slide 177: Link Budget Considerations: Antenna Gain

Link Budget Considerations: Antenna Gain All images from: oscarliang.com/how-antenna-gain-affects-range Impossible, perfectly isotropic: Spherical pattern Monopole/dipole: Donut-shaped pattern Patch: One roughly hemispherical main lobe Image source: dg7ybn.de/432MHz/GTV70_17m.htm * Antenna types and gains vary widely. A theoretical isotropic antenna is commonly used as the gain reference, hence gain units of dBi (dB relative to isotropic). * Gain and directivity directly related. An antenna with high gain will be highly directional and require greater pointing control * Gain patterns are affected by the structure to which they are mounted * Transmit and receive gain pattern are the same * Antenna patterns are commonly shown in two planes * Most antennas have at least 1 axis of near-symmetry. * Imagine cuts through a 3D pattern, resulting in figures to the right

=== Slide 178: Link Budget Considerations: Path Loss

Link Budget Considerations: Path Loss * Largest system loss * Caused by power spread, not atmosphere or other attenuators * Distance and power related by inverse square law * Applies to far-field * Power density may be inconsistent in near field * far field definition somewhat subjective: large distance relative to system scale Power transmitted by antenna Same amount of power must cover greater area as wave travels. Power density decreases by square of distance Receiving antenna collects power equivalent to power density multiplied by effective area

=== Slide 179: Link Budget Considerations: Noise

Link Budget Considerations: Noise * To communicate, power must generally be greater than noise * All systems have thermal noise * Robustness depends on signal to noise ratio * Noise varies with: * Frequency * Location * Temperature * Antenna pointing Pi-plates.com/spectrum-analyzer Noise floor at -120 dB Frequency (MHz) Amplitude (dB) Desired signal at 1000 MHz Other signals

=== Slide 180: Link Budget Considerations: Units

Link Budget Considerations: Units 𝑅𝑒𝑐𝑒𝑖𝑣𝑒𝑑 𝑃𝑜𝑤𝑒𝑟 = 𝑇𝑟𝑎𝑛𝑠𝑚𝑖𝑡𝑡𝑒𝑑 𝑃𝑜𝑤𝑒𝑟 × 𝐺𝑎𝑖𝑛𝑠 𝐿𝑜𝑠𝑠𝑒𝑠 * Values span many orders of magnitude * Equations with many multiplied/divided terms are easy to confuse * Standard is to use dB instead of decimal values * dB values are ratios, not absolutes * Variations of dB are absolutes due to a specified reference * dBW: relative to 1 W * dBm: relative to 1 mW * dBi: relative to isotropic antenna * Recall log rules: * Multiplication becomes addition * Division becomes subtraction 𝑑𝐵 = 10𝑙𝑜𝑔10 𝑂𝑢𝑡𝑝𝑢𝑡 𝐼𝑛𝑝𝑢𝑡 𝑅𝑒𝑐𝑒𝑖𝑣𝑒𝑑 𝑃𝑜𝑤𝑒𝑟 = 𝑇𝑟𝑎𝑛𝑠𝑚𝑖𝑡𝑡𝑒𝑑 𝑃𝑜𝑤𝑒𝑟 + 𝐺𝑎𝑖𝑛𝑠 − 𝐿𝑜𝑠𝑠𝑒𝑠

=== Slide 181: Bringing it Together

• Collect known values into spreadsheet and calculate others

• Link budgets vary greatly in complexity

• Use provided template for Mission Design Course

• Highly simplified, not thorough enough for real mission

• Ignores many terms and mission-specific factors

• Output results in common format

• C/N (carrier/signal to noise ratio)

• General power comparison, but requires understanding of assumptions

• Eb/No (Energy per bit-to-noise density ratio)

• Considers data rate

• Add margin

• UNP requires 6 dB link margin at 10 deg elevation mask

• Does link close?

• Usually yes: COTS products are designed for this Bringing it Together www.nasa.gov/smallsat-institute/sst-soa/communications

=== Slide 182: For This Course

For This Course Transmit (Tx) Radio * Data sheet provides power, frequency, data rate, etc. * Includes transmitter, amplifier, some losses Antennas * Include both transmit and receive * Data sheet provides gain, make reasonable assumptions about gain pattern if unknown Losses * Include free space path loss * Consider if known/applicable * Pointing losses * Polarization losses * Line losses * Circuit losses Receive (Rx) Radio * Data sheet provides sensitivity, frequency, data rate, etc. General * NASA State-of-the-Art Small Spacecraft Technology has a few helpful tables for initial radio and antenna trades. Section 9.7

=== Slide 183: Downlink Budget: Orbit Information

Downlink Budget: Orbit Information Key Derived value Value to Enter Orbit Information Parameter Value Unit Elevation Angle 10 ° Altitude of Satellite 600 km Radius of Earth 6378 km Speed of Light 299,792,458 m/s Boltzmann Constant 1.38E-23 W/K/Hz Distance 1932 km UNP Requirement Constants Solving using law of cosines Input

=== Slide 184: Downlink Budget: Transmitter Information

Downlink Budget: Transmitter Information Key Derived value Value to Enter Transmitter Information Parameter Value Unit Frequency 2.23 GHz Wavelength 0.134 m Utilized Bandwidth 0.0125 MHz Distance 1932 km Bit rate 1 Mbps Transmit Power 1.0 W Transmit Power 30 dBm Transmit antenna gain 6.5 dBi Transmit system losses -4 dB EIRP 32.5 dBm Path loss -163.33 dB S-Band Frequency, refer to datasheet Solving using law of cosines S-Band data rate, refer to datasheet S-Band radio transmit power, refer to datasheet 𝜆 = 𝑆𝑝𝑒𝑒𝑑 𝑜𝑓 𝐿𝑖𝑔ℎ𝑡 𝐹𝑟𝑒𝑞𝑢𝑒𝑛𝑐𝑦 Bandwidth radio uses/can be licensed for

=== Slide 185: Downlink Budget: Transmitter Information

Downlink Budget: Transmitter Information Key Derived value Value to Enter Converts Watts to dBm (decibels with reference to one milliwatt, mW) 𝑇𝑟𝑎𝑛𝑠𝑚𝑖𝑡 𝑃𝑜𝑤𝑒𝑟(𝑑𝐵𝑚) = 10𝑙𝑜𝑔10 𝑇𝑟𝑎𝑛𝑠𝑚𝑖𝑡 𝑃𝑜𝑤𝑒𝑟 (𝑊) ∗ 1000 30 𝑑𝐵𝑚 = 10𝑙𝑜𝑔10 1 ∗ 1000 (Spreadsheet includes Return Loss, next slide) Input, refer to data sheet Transmitter Information Parameter Value Unit Frequency 2.23 GHz Wavelength 0.134 m Distance 1570 km Bit rate 1 Mbps Transmit Power 1.0 W Transmit Power 30 dBm Transmit antenna gain 6.5 dBi Transmit system losses -4 dB EIRP 32.5 dBm Path loss -163.33 dB

=== Slide 186: Downlink Budget: Transmitter Information

Downlink Budget: Transmitter Information Transmitter Information Parameter Value Unit Frequency 2.23 GHz Wavelength 0.134 m Distance 1570 km Bit rate 1 Mbps Transmit Power 1 W Transmit Power 30 dBm Transmit antenna gain 6.5 dBi Transmit system losses -4 dB EIRP 32.5 dBm Path loss -163.33 dB Key Derived value Value to Enter Input, refer to datasheet Assumes losses: * Circuit Loss: -1 dB * Return Loss: -0.5 dB (VSWR of 2:1) * Pointing Loss: -2.5 dB Effective Isotropic Radiated Power (EIRP) 𝐸𝐼𝑅𝑃 = 𝑇𝑟𝑎𝑛𝑠𝑚𝑖𝑡 𝑃𝑜𝑤𝑒𝑟 + 𝐴𝑛𝑡𝑒𝑛𝑛𝑎 𝐺𝑎𝑖𝑛 + 𝐿𝑜𝑠𝑠𝑒𝑠 32.5 𝑑𝐵𝑚 = 30 + 6.5 − 4

=== Slide 187: Downlink Budget: Transmitter Information

Downlink Budget: Transmitter Information Key Derived value Value to Enter Make sure: * Distance and Wavelength are in meters 𝑃𝑎𝑡ℎ 𝐿𝑜𝑠𝑠 = −10𝑙𝑜𝑔10 4𝜋 ∗ 𝐷𝑖𝑠𝑡𝑎𝑛𝑐𝑒 𝜆 −165.14 𝑑𝐵 = −10𝑙𝑜𝑔10 4𝜋 ∗ 1932 ∗ 103 0.134 Transmitter Information Parameter Value Unit Frequency 2.23 GHz Wavelength 0.134 m Distance 1932 km Bit rate 1 Mbps Transmit Power 1 W Transmit Power 30 dBm Transmit antenna gain 6.5 dBi Transmit system losses -4 dB EIRP 32.5 dBm Path loss -165.14 dB

=== Slide 188: Downlink Budget: Transmitter Information

Downlink Budget: Transmitter Information Key Derived value Value to Enter Full Transmitter Information Transmitter Information Parameter Value Unit Frequency 2.23 GHz Wavelength 0.134 m Utilized Bandwidth 0.0125 MHz Distance 1932 km Bit rate 1.0 Mbps Transmit Power 1.0 W Transmit Power 30 dBm Transmit antenna gain 6.5 dBi Transmit system losses -4.0 dB EIRP 32.5 dBm Path loss -165.14 dB

=== Slide 189: Downlink Budget: Receiver Information

Downlink Budget: Receiver Information Key Derived value Value to Enter Receiver Information Parameter Value Unit Receive antenna diameter 5 m Antenna Efficiency 60 % Receive antenna gain 39.13 dBi Noise Temperature 250 K Receive system noise figure -174.62 dBm/Hz Total Received Power -87.70 dBm Receiver system losses -4 dB Cross polarization loss 0 dB Reference from datasheet or “worst case”

=== Slide 190: Downlink Budget: Receiver Information

Downlink Budget: Receiver Information Key Derived value Value to Enter Receiver Information Parameter Value Unit Receive antenna diameter 5 m Antenna Efficiency 60 % Receive antenna gain 39.13 dBi Noise Temperature 250 K Receive system noise figure -174.62 dBm/Hz Total Received Power -87.70 dBm Receiver system losses -4 dB Cross polarization loss 0 dB 𝑅𝑒𝑐𝑒𝑖𝑣𝑒 𝐴𝑛𝑡𝑒𝑛𝑛𝑎 𝐺𝑎𝑖𝑛 = 10𝑙𝑜𝑔10 𝐸𝑓𝑓𝑖𝑐𝑖𝑒𝑛𝑐𝑦 ∗ 4𝜋𝐴𝑟 𝜆 𝑅𝑒𝑐𝑒𝑖𝑣𝑒 𝐴𝑛𝑡𝑒𝑛𝑛𝑎 𝐺𝑎𝑖𝑛 = 10𝑙𝑜𝑔10 𝐸𝑓𝑓𝑖𝑐𝑖𝑒𝑛𝑐𝑦 ∗ 4 𝜋 ∗ 𝑅𝑒𝑐𝑒𝑖𝑣𝑒 𝑑𝑖𝑎𝑚𝑒𝑡𝑒𝑟 𝜆 39.13 𝑑𝐵𝑖 = 10𝑙𝑜𝑔10 0.6 ∗ 𝜋 ∗ 5 0.134 Make sure: * Diameter and Wavelength are in meters Transmitter Information Parameter Value Unit Wavelength 0.134 m Transmit Power 30 dBm Transmit antenna gain 6.5 dBi Path loss -163.33 dB

=== Slide 191: Downlink Budget: Receiver Information

Downlink Budget: Receiver Information Key Derived value Value to Enter Receiver Information Parameter Value Unit Receive antenna diameter 5 m Antenna Efficiency 60 % Receive antenna gain 39.13 dBi Noise Temperature 307.0 K Receive system noise figure -173.73 dBm/Hz Total Received Power -87.70 dBm Receiver system losses -4 dB Cross polarization loss 0 dB Generic estimate 𝑆𝑦𝑠𝑡𝑒𝑚 𝑁𝑜𝑖𝑠𝑒 𝐹𝑖𝑔𝑢𝑟𝑒 = 10𝑙𝑜𝑔10 𝐵𝑜𝑙𝑡𝑧𝑚𝑎𝑛𝑛 𝐶𝑜𝑛𝑠𝑡𝑎𝑛𝑡 ∗ 1000 ∗ 𝑁𝑜𝑖𝑠𝑒 𝑇𝑒𝑚𝑝 −173.73 𝑑𝐵𝑚 𝐻𝑧 = 10𝑙𝑜𝑔10 1.38 ∗ 10−23 ∗ 1000 ∗ 307

=== Slide 192: Downlink Budget: Receiver Information

Downlink Budget: Receiver Information Key Derived value Value to Enter Receiver Information Parameter Value Unit Receive antenna diameter 5 m Antenna Efficiency 60 % Receive antenna gain 39.13 dBi Noise Temperature 250 K Receive system noise figure -174.62 dBm/Hz Receiver system loss -3.09 dB Total Received Power -96.60 dBm Cross polarization loss -3 dB Transmitter Information Parameter Value Unit Wavelength 0.134 m Transmit Power 30 dBm Transmit antenna gain 6.5 dBi Path loss -165.14 dB Total Received Power = 𝐸𝐼𝑅𝑃 + 𝑆𝑦𝑠𝑡𝑒𝑚 𝐿𝑜𝑠𝑠 + 𝑃𝑎𝑡ℎ 𝐿𝑜𝑠𝑠 + 𝑅𝑒𝑐𝑒𝑖𝑣𝑒 𝐺𝑎𝑖𝑛 −96.60 𝑑𝐵𝑚 = 32.5 − 3.09 − 165.14 + 39.13 Assumes: * Antenna pointing loss - ground: -1 dB * Radome Loss: 0 dB * Circuit Loss: -1 dB * Atmospheric/rain loss: -1.088 dB If ground antenna vs. flight antennas have same-hand circular or linear (0dB), linear to circular (-3dB) or LHCP to RHCP is often about -20dB or -30dB loss

=== Slide 193: Downlink Budget: Receiver Information

Downlink Budget: Receiver Information Key Derived value Value to Enter Receiver Information Parameter Value Unit Receive antenna diameter 5 m Antenna Efficiency 60 % Receive antenna gain 39.13 dBi Noise Temperature 307.0 K Receive system noise figure -173.73 dBm/Hz Receiver system losses -3.09 dB Total Received Power -96.60 dBm Cross polarization loss -3 dB Full Receiver Information

=== Slide 194: Downlink Budget: Link Margin

Downlink Budget: Link Margin Key Derived value Value to Enter Receiver Information Parameter Value Unit Receive system noise figure -173.73 dBm/Hz Total Received Power -96.60 dBm Receiver system losses -4 dB Cross polarization loss 0 dB Transmitter Information Parameter Value Unit Bit Rate 1.00 Mbps Link Margin Computation Received Eb/No 14.13 dB Required Eb/No 3 dB Link margin 19.92 dB Required link margin 6 dB Bit rate (dB) = 10𝑙𝑜𝑔10 𝐵𝑖𝑡 𝑅𝑎𝑡𝑒(𝑀𝑏𝑝𝑠) ∗ 106 𝐸𝑏/𝑁0 = 14.13 𝑑𝐵 = −96.6 − −173.73 − 60 + (−3) 𝐸𝑏/𝑁0 = 𝑅𝑒𝑐𝑒𝑖𝑣𝑒 𝑃𝑜𝑤𝑒𝑟 − 𝑅𝑒𝑐𝑒𝑖𝑣𝑒 𝑆𝑦𝑠𝑡𝑒𝑚 𝑁𝑜𝑖𝑠𝑒 − 𝐵𝑖𝑡 𝑅𝑎𝑡𝑒 𝑑𝐵 + 𝑃𝑜𝑙𝑎𝑟𝑖𝑧𝑎𝑡𝑖𝑜𝑛 𝐿𝑜𝑠𝑠 𝑑𝐵

=== Slide 195: Downlink Budget: Link Margin

Downlink Budget: Link Margin Key Derived value Value to Enter Receiver Information Parameter Value Unit Receiver system losses -3 dB Cross polarization loss -3 dB Receive system implementation loss -2 dB Based upon various modulation schemes for a desired bit error rate. 𝐿𝑖𝑛𝑘 𝑀𝑎𝑟𝑔𝑖𝑛 = 𝑅𝑒𝑐𝑒𝑖𝑣𝑒 𝐸𝑏/𝑁0 − 𝑅𝑒𝑞𝑢𝑖𝑟𝑒𝑑 𝐸𝑏 𝑁 0 − 𝐼𝑚𝑝𝑙𝑒𝑚𝑒𝑛𝑡𝑎𝑡𝑖𝑜𝑛 𝐿𝑜𝑠𝑠 𝐿𝑖𝑛𝑘 𝑀𝑎𝑟𝑔𝑖𝑛 = 1.12 𝑑𝐵 = 14.12 − 11.0 − 2.0 UNP Requirement Link Margin Computation Received Eb/No 14.12 dB Required Eb/No 11.0 dB Link margin 1.12 dB Required link margin 6 dB

=== Slide 196: Downlink (Space to Ground)

Transmitter Information Frequency 2.23 GHz Wavelength 0.134 m Utilized Bandwidth 0.0125 MHz Distance 1932 km Bit rate 1.000 Mbps Transmit Power 1.0 W Transmit Power 30 dBm Transmit antenna gain 6.5 dBi Transmit system losses -4 dB EIRP 32.5 dBm Path loss -165.14 dB Downlink (Space to Ground) Key Derived value Value to Enter Preliminary link completed for S-band system with un-acceptable margin: may need redesign Orbit Information Elevation Angle 10 ° Altitude of Satellite 600 km Radius of Earth 6378 km Speed of Light 299,792,458 m/s Boltzmann Constant 1.38E-23 W/K/Hz Distance 1932 km Receiver Information Receive antenna diameter 5 m Antenna Efficiency 60 % Receive antenna gain 39.13 dBi Noise Temperature 307.0 K Receive system noise figure -173.73 dBm/Hz Receiver System losses -3 dB Total Received Power -98.51 dBm Cross polarization loss -3 dB Link Margin Computation Received Eb/No 12.2 dB Required Eb/No 11.0 dB Receive System Implementation Loss -2 dB Link margin 1.12 dB Required link margin 6 dB Values to enter may come from data sheets, specific implementation choices, CONOPS, and more Link budget taken from Space Mission Engineering - Wertz

=== Slide 197: Link Budget for Mission Design Course

• Choose an antenna and transceiver and use the data sheet values to feed data budget

• Note: Not expected to complete link budget During this course

• Ground station(s) are up to you

• Ask questions if you are stuck or need help

• NASA State-of-the-Art Small Spacecraft Technology has a few helpful tables for initial radio and antenna trades. Section 9.7 Link Budget for Mission Design Course www.nasa.gov/smallsat-institute/sst-soa/communications Ensure link closes

=== Slide 198: Power/Energy Budget

Power/Energy Budget

=== Slide 199: Power + Energy

These budgets answer… Power + Energy How should the system be sized? * How much solar power do I need/is it enough? * How much battery capacity do I need/is it enough? Am I power-positive in any given mode? * What is the power draw of any given mode? What is the impact(s) to CONOPS? * How long can my payload be on?

=== Slide 200: Power System Elements

Power System Elements Generation * Solar Cells * Peak Power Tracker Storage * Batteries * Circuit protection Distribution * Power Draw from subsystems * Regulation * Switches Power system design can get complex. Here, we’ll look at the amount of power we need and the time cadence. This assessment relies on the three shown elements.

=== Slide 201: Power Budget Elements

Power Budget Elements Do I create more power than I use? Do I have enough power in storage? Part 1: Power Consumption & Generation Part 2: Energy Balance

=== Slide 202: Assessing Power Consumption

Assessing Power Consumption Identify all components that consume power Step 1 * Identify how much power is consumed * Identify if there are different power states for a component (and split them out) * Example- a radio likely has different power draws for transmit and receive states Identify all CONOPS States/Modes Step 2 * UNP recommends a table with step 1 for the rows and step 2 for columns * Identify what systems are powered on in each mode * Identify the duty cycle of the component for each mode Identify power draw for components per mode Step 3 * Calculate based on duty cycle and component power draw for each mode * Multiply duty cycle by component power draw Sum the power consumption for each mode state Step 4 * Note the highest power draw mode, use for future analysis (conservative case) Component Draw Mode 1 Mode 2 Duty Cycle Power Draw Duty Cycle Power Draw A 1W 100% 1W 100% 1W B 2W 50% 1W 100% 2W Mode Average Power 2W 3W

=== Slide 203: Assessing Power Consumption

Assessing Power Consumption Identify all components that consume power Step 1 * Identify how much power is consumed * Identify if there are different power states for a component (and split them out) * Example- a radio likely has different power draws for transmit and receive states Identify all CONOPS States/Modes Step 2 * UNP recommends a table with step 1 for the rows and step 2 for columns * Identify what systems are powered on in each mode * Identify the duty cycle of the component for each mode Identify power draw for components per mode Step 3 * Calculate based on duty cycle and component power draw for each mode * Multiply duty cycle by component power draw Sum the power consumption for each mode state Step 4 * Note the highest power draw mode, use for future analysis (conservative case) Component Draw Mode 1 Mode 2 Duty Cycle Power Draw Duty Cycle Power Draw A 1W 100% 1W 100% 1W B 2W 50% 1W 100% 2W Mode Average Power 2W 3W

=== Slide 204: Assessing Power Consumption

Assessing Power Consumption Identify all components that consume power Step 1 * Identify how much power is consumed * Identify if there are different power states for a component (and split them out) * Example- a radio likely has different power draws for transmit and receive states Identify all CONOPS States/Modes Step 2 * UNP recommends a table with step 1 for the rows and step 2 for columns * Identify what systems are powered on in each mode * Identify the duty cycle of the component for each mode Identify power draw for components per mode Step 3 * Calculate based on duty cycle and component power draw for each mode * Multiply duty cycle by component power draw Sum the power consumption for each mode state Step 4 * Note the highest power draw mode, use for future analysis (conservative case) Component Draw Mode 1 Mode 2 Duty Cycle Power Draw Duty Cycle Power Draw A 1W 100% 1W 100% 1W B 2W 50% 1W 100% 2W Mode Average Power 2W 3W

=== Slide 205: Assessing Power Consumption

Assessing Power Consumption Identify all components that consume power Step 1 * Identify how much power is consumed * Identify if there are different power states for a component (and split them out) * Example- a radio likely has different power draws for transmit and receive states Identify all CONOPS States/Modes Step 2 * UNP recommends a table with step 1 for the rows and step 2 for columns * Identify what systems are powered on in each mode * Identify the duty cycle of the component for each mode Identify power draw for components per mode Step 3 * Calculate based on duty cycle and component power draw for each mode * Multiply duty cycle by component power draw Sum the power consumption for each mode state Step 4 * Note the highest power draw mode, use for future analysis (conservative case) Component Draw Mode 1 Mode 2 Duty Cycle Power Draw Duty Cycle Power Draw A 1W 100% 1W 100% 1W B 2W 50% 1W 100% 2W Mode Average Power 2W 3W

=== Slide 206: ShipSat Power Consumption

ShipSat Power Consumption Component Component Draw Safe Mode Standby Mode Experiment Mode (W) Duty Cycle Power Draw (W) Duty Cycle Power Draw (W) Duty Cycle Power Draw (W) CDH 1 50% 0.5 100% 1 100% 1 Radio Tx 6 5.56% 0.33 5.56% 0.33 5.56% 0.33 Radio Rx 2 100% 2 100% 2 100% 2 GPS 1 0% 0 100% 1 100% 1 ADCS 2 0% 0 100% 2 100% 2 Power System 0.5 100% 0.5 100% 0.5 100% 0.5 Heater(s) 2 50% 1 50% 1 50% 1 Payload Imager 5 0% 0 0% 0 60% 3 Power Draw Per Mode 4.33 W 7.83 W 10.83 W

=== Slide 207: Power Generation Assessment

Step 1 * Identify array count Step 2 * Identify cell power specifications Step 3 * Multiply cell power by number of cells or strings on the arrays Step 4 * Identify array power & correlate SV body axis Power Generation Assessment Note you may want to consider an efficiency loss in the peak power tracking

=== Slide 208: Solar Panel Configuration

Solar Panel Configuration 1 array per large face. 4 arrays total String 1 String 2 6 series cells per string 2 parallel strings per array

=== Slide 209: ShipSat Generation Example

ShipSat Generation Example Generated Power Array Series Cells/String # of Strings Cell Voltage (V) Cell Current (A) Array Voltage (V) Array Current (A) Array Power (W) Axis VOC VMP ISC IMP VOC VMP ISC IMP PMP Array 1 6 2 2.6 2.5 0.5 0.4 15.6 15 1 0.8 12 X Determined by mission Subscripts * OC: Open circuit * SC: Short circuit * MP: Max power

=== Slide 210: ShipSat Generation Example

ShipSat Generation Example Generated Power Array Series Cells/String # of Strings Cell Voltage (V) Cell Current (A) Array Voltage (V) Array Current (A) Array Power (W) Axis VOC VMP ISC IMP VOC VMP ISC IMP PMP Array 1 6 2 2.6 2.5 0.5 0.4 15.6 15 1 0.8 12 X Obtained from data sheet Subscripts * OC: Open circuit * SC: Short circuit * MP: Max power

=== Slide 211: ShipSat Generation Example

ShipSat Generation Example Generated Power Array Series Cells/String # of Strings Cell Voltage (V) Cell Current (A) Array Voltage (V) Array Current (A) Array Power (W) Axis VOC VMP ISC IMP VOC VMP ISC IMP PMP Array 1 6 2 2.6 2.5 0.5 0.4 15.6 15 1 0.8 12 X Obtained from data sheet ISC = JSC * Area of solar cell IMP = JMP* Area of solar cell Subscripts * OC: Open circuit * SC: Short circuit * MP: Max power

=== Slide 212: ShipSat Generation Example

ShipSat Generation Example Generated Power Array Series Cells/String # of Strings Cell Voltage (V) Cell Current (A) Array Voltage (V) Array Current (A) Array Power (W) Axis VOC VMP ISC IMP VOC VMP ISC IMP PMP Array 1 6 2 2.6 2.5 0.5 0.4 15.6 15 1 0.8 12 X = Subscripts * OC: Open circuit * SC: Short circuit * MP: Max power

=== Slide 213: ShipSat Generation Example

ShipSat Generation Example Generated Power Array Series Cells/String # of Strings Cell Voltage (V) Cell Current (A) Array Voltage (V) Array Current (A) Array Power (W) Axis VOC VMP ISC IMP VOC VMP ISC IMP PMP Array 1 6 2 2.6 2.5 0.5 0.4 15.6 15 1 0.8 12 X = Subscripts * OC: Open circuit * SC: Short circuit * MP: Max power

=== Slide 214: ShipSat Generation Example

ShipSat Generation Example Generated Power Array Series Cells/String # of Strings Cell Voltage (V) Cell Current (A) Array Voltage (V) Array Current (A) Array Power (W) Axis VOC VMP ISC IMP VOC VMP ISC IMP PMP Array 1 6 2 2.6 2.5 0.5 0.4 15.6 15 1 0.8 12 X = Subscripts * OC: Open circuit * SC: Short circuit * MP: Max power

=== Slide 215: ShipSat Generation Example

ShipSat Generation Example Generated Power Array Series Cells/String # of Strings Cell Voltage (V) Cell Current (A) Array Voltage (V) Array Current (A) Array Power (W) Axis VOC VMP ISC IMP VOC VMP ISC IMP PMP Array 1 6 2 2.6 2.5 0.5 0.4 15.6 15 1 0.8 12 X = Subscripts * OC: Open circuit * SC: Short circuit * MP: Max power

=== Slide 216: ShipSat Generation Example

ShipSat Generation Example Generated Power Array Series Cells/String # of Strings Cell Voltage (V) Cell Current (A) Array Voltage (V) Array Current (A) Array Power (W) Axis VOC VMP ISC IMP VOC VMP ISC IMP PMP Array 1 6 2 2.6 2.5 0.5 0.4 15.6 15 1 0.8 12 X = Subscripts * OC: Open circuit * SC: Short circuit * MP: Max power

=== Slide 217: ShipSat Generation Example

ShipSat Generation Example Generated Power Array Series Cells/String # of Strings Cell Voltage (V) Cell Current (A) Array Voltage (V) Array Current (A) Array Power (W) Axis VOC VMP ISC IMP VOC VMP ISC IMP PMP Array 1 6 2 2.6 2.5 0.5 0.4 15.6 15 1 0.8 12 X Array 2 6 2 2.6 2.5 0.5 0.4 15.6 15 1 0.8 12 -X Array 3 6 2 2.6 2.5 0.5 0.4 15.6 15 1 0.8 12 Y Array 4 6 2 2.6 2.5 0.5 0.4 15.6 15 1 0.8 12 -Y Totals 62.4 60 4 3.2 N/A Max generation with single face pointing at sun is 12W. Could consider 2 faces at angle in this configuration

=== Slide 218: Generation vs Consumption

• 10.83W > 8W which isn’t good. Consider:

• Power is not time-based. This is instantaneous power draw.

• Have we considered Orbit Average power?

• Do we need to generate more than we consume for every mode?

• For Safe mode, yes.

• For Standby, yes.

• For Experiment, maybe.

• Generation vs consumption doesn’t give us a full picture of the power system state in time. For that, we need to run an energy budget. Generation vs Consumption Safe Mode (W) Standby Mode (W) Experiment Mode (W) Consumption (From previous consumption table) 4.33 7.83 10.83 Generation 8 12 8 Assuming tumbling, average generation. (48W total generating surfaces divided by 6 faces) Assuming one panel sun facing Pointing imager at ground with secondary axis sun pointing. Assuming 2/3 of sun pointing generation as a rough estimate. This is not constant and may be a poor assumption

=== Slide 219: Power and Energy

Power and Energy Safe Standby Exp Explanation Mode Power (W) Draw 4.33 7.83 10.83 Mode/instantaneous power from earlier in the presentation Generation 8 12 8 Orbit Average Power (W) Draw 4.33 7.83 10.83 Orbit average power is a time weighted average of power generation and draw across an orbit. Generation only occurs in sunlight which is assumed to be 60% of an orbit for this analysis. So, the mode power generation values are multiplied by 0.6 (60%). Consumption is the same as mode power because the satellite consumes power equally in sunlight and eclipse Generation 4.8 7.2 4.8 Orbit Average Energy (Wh) Draw 6.5 11.75 16.25 Orbit average energy takes total time at a given power draw into account to determine the energy in and out of the battery. This is equal to orbit average power multiplied by the orbit period. In this analysis the period is assumed to be 1.5 hours. Generation 7.2 10.8 7.2

=== Slide 220: Battery Storage Assessment

• For battery storage, you may buy a pack that indicates pack power, or you may have to calculate it.

• Identify pack power for energy balance

• Identify min and max voltages

• Informs safe mode threshold

• Complete battery discharge curve should be measured/considered for real mission

• For ShipSat, we will assume a 71Wh battery pack. Battery Storage Assessment Storage (Battery) Cell Voltage (V) Cell Capacity (Ah) Series Cells/String # of Strings Pack Voltage (V) Pack Capacity (Ah) Pack Energy (Wh) Nominal 3.7 2.4 4 2 14.8 4.8 71.04 Max 4.2 16.8 Min 2.75 11.0

=== Slide 221: Energy Balance

• There are many ways to do this. Ultimately, you will likely want a more mature simulation or analysis during later phases of design. For early mission design, we can simplify this process. Energy Balance Identify Orbit Average Power Draw Step 1

• For both extremes (i.e., worst- and best-case power) Identify orbit average power generation Step 2

• This includes assessment of eclipse Compare / Identify battery discharge Step 3

• Per orbit, what is the depth of discharge in eclipse?

• What is the charge/discharge trend over long periods of time (many orbits/times of year)?

• If the trend is negative, the budget won’t close

=== Slide 222: Ultra-Simple Energy Balance

• Assumptions:

• Using mix of conservative and less conservative numbers for median analysis

• Generation

• Sunlit for 60% of the orbit (LEO)

• Panel facing sun in standby

• 2/3 of sun-facing generation in experiment mode

• Consumption: See power budget

• Increase fidelity by:

• Increasing duty cycle/timeline detail

• Increasing orbit/attitude detail Ultra-Simple Energy Balance Safe Standby Exp Orbit Average Energy (Wh) Draw 6.5 11.75 16.25 Generation 7.2 10.8 7.2 Battery Energy Difference Per Orbit 0.7 -0.95 -9.1

• Except for safe mode, orbit average consumption is greater than generation – not great

• In Imaging mode, we would lose ~10Wh per orbit, depleting our 70Wh pack quickly!

=== Slide 223: Including CONOPS

Including CONOPS Safe Standby Experiment * Semi-realistic mode sequence * 4 orbits safe * 4 orbits standby * 8 orbits experiment * Required behavior * Safe mode only for off-nominal cases * Standby mode indefinitely power positive * Strongly desired behavior – contributes greatly to meeting full success criteria for this mission * Experiment mode indefinitely power positive * Standby mode slightly power negative so configuration already unacceptable Mode Orbit Avg Power Consumed (W) Orbit Avg Power Generated (W) Safe 4.33 8 Standby 7.83 12 Experiment 10.83 8

=== Slide 224: Adding Realism and Safety

Adding Realism and Safety * Battery requirement: Battery voltage must stay above 12 V * At 12 V, satellite flips to safe mode * At 11 V, power system cuts power to protect batteries * Same mode sequence as last slide * 4 orbits safe * 4 orbits standby * ~1 orbit experiment until battery limit is reached * Remaining orbits in safe Mode Orbit Avg Power Consumed (W) Orbit Avg Power Generated (W) Safe 4.33 8 Standby 7.83 12 Experiment 10.83 8 Safe Standby Safe Exp Battery Voltage (V) Battery Voltage in One Day

=== Slide 225: Adjusting Energy Budget

• Payload will only be capturing images in sunlight - it can’t take visible spectrum pictures in eclipse

• We can point more solar panels towards zenith as the image points nadir – we’ll assume 4 panels total, but in various configurations

• We could also change CONOPS assumptions such as payload duty cycle Adjusting Energy Budget

=== Slide 226: Increasing Generation

Increasing Generation 2 Sun-Facing Panels 3 Sun-Facing Panels 4 Sun-Facing Panels Safe Standby Exp Safe Standby Exp Safe Standby Exp Orbit Average Power (W) Draw 4.33 7.83 10.83 4.33 7.83 10.83 4.33 7.83 10.83 Generation 4.8 14.4 9.6 4.8 21.6 14.4 4.8 28.8 19.2 Orbit Average Energy (Wh) Draw 6.5 11.75 16.25 6.5 11.75 16.25 6.5 11.75 16.25 Generation 7.2 21.6 14.4 7.2 32.4 21.6 7.2 43.2 28.8 Battery Discharge Meets hard requirements but will not allow indefinite period in imaging mode. Generation higher than consumption in all modes Generation higher than consumption in all modes. Likely excessive. * Experiment mode generation: If imager is nadir pointing, solar panels will face zenith. They will generally see sun, but with cosine loss. This analysis assumed imaging mode generation is 2/3 of sun pointing generation * Tumbling generation: Tumbling generation is highly simplified. Having solar panels on multiple faces may be beneficial to accommodate more tumble states, though average power won’t change * Battery state of charge (SOC): Assume 40% of orbit is eclipse (worst case). Calculate state of charge after eclipse to ensure battery does not reach limit

=== Slide 227: Mode Energy Balance: 2 Panels on Same Face

Mode Energy Balance: 2 Panels on Same Face * Assuming single mode for duration * Assuming 4 panels total for tumbling, but 2 are on other faces * Results * Safe mode power positive * Standby now power positive * Experiment still negative, but less so. May now get ~5 orbits before kicking into safe mode rather than ~1 Mode Orbit Avg Power Consumed (W) Orbit Avg Power Generated (W) Safe 4.33 4.8 Standby 7.83 14.4 Experiment 10.83 9.6

=== Slide 228: Mode Energy Balance: 3 Panels on Same Face

Mode Energy Balance: 3 Panels on Same Face * Assuming single mode for duration * Assuming 4 panels total for tumbling, but 1 is on another face * Results * Safe mode power positive * Standby power positive * experiment now power positive Mode Orbit Avg Power Consumed (W) Orbit Avg Power Generated (W) Safe 4.33 4.8 Standby 7.83 21.6 Experiment 10.83 14.4

=== Slide 229: Mode Energy Balance: 4 Panels on Same Face

Mode Energy Balance: 4 Panels on Same Face * Assuming single mode for duration * Assuming all 4 panels on same face * Results * Safe mode power positive * Standby power positive * Experiment power positive * Likely excessive * Further analysis with better assumptions may change results * Tumbling axes/rates may cause more drastic battery capacity swings with fewer faces of coverage * Imaging mode power generation assumption was very rough Mode Orbit Avg Power Consumed (W) Orbit Avg Power Generated (W) Safe 4.33 4.8 Standby 7.83 28.8 Experiment 10.83 19.2

=== Slide 230: Chosen Configuration: 3 Panels on Same Face

Chosen Configuration: 3 Panels on Same Face 3 arrays on large face 1 body-mounted, 2 deployable 4th array body mounted on opposite face * Much simpler than 4 panels on face * Provides nearly full sphere coverage of solar panels * Panels on opposite sides * Good for tumble * Avoids shading * Provides good safety if failure to deploy * Would still generate and would look like original configuration if two failures, or like the two panel on same face configuration if 1 failure * CONOPS would have to adapt but mission would still be successful * Earlier assumption stated the imager needs 3U length, so if it is pointing nadir, is the shown configuration valid for this analysis? 3 arrays on large face 1 body-mounted, 2 deployable Panels fold flat against sides when stowed so they are still exposed and functional if deployment fails Z X Y

=== Slide 231: Returning to CONOPS

Returning to CONOPS * Modeling with original mode sequence * 4 orbits safe * 4 orbits standby * 8 orbits experiment * Results * Power positive across all modes * Initial design closes, future work should critically evaluate assumptions Safe Standby Experiment Mode Orbit Avg Power Consumed (W) Orbit Avg Power Generated (W) Safe 4.33 4.8 Standby 7.83 21.6 Experiment 10.83 14.4

=== Slide 232: Refining Energy Balance

Improved CONOPS * Mission sequence * Mode and component switching * Environmental effects * Eclipse times * Solar panel temperatures * Orbit and attitude Consistency with other analyses * If the data budget only assumes 5 images per day, why is the imager on whenever it is in sun? * What are the requirements? Margins * Safe mode margin is small – somewhat concerning * Other modes have plenty of margin and can accept more risk Refining Energy Balance

=== Slide 233: Refining Energy Balance

Component duty cycles vs actual time-phasing of power draw * Power duty cycles are averaged over orbit in this analysis rather than instantaneous * Multi-orbit trends may be valid, but slopes within an orbit period are not * Imager on during day, off during night * Radio transmit on for a few minutes at high power draw, then off for a long time Battery and solar cell degradation * Generally not concerning for CubeSats due to short lifespans Power generation during imaging and tumble were very loose estimates Shading of panels by other panels Many more… Refining Energy Balance

=== Slide 234: Power Budget for Mission Design Course

• Preliminary numbers and duty cycles

• CONOPS mode definitions with power states

• Panel configuration

• Assume worst case Power Budget for Mission Design Course Achieve power balance, with margin, based upon system size and CONOPS choices Z X Y

=== Slide 235: Volume and Mass Budgets

Volume and Mass Budgets

=== Slide 236: Mission Design Expectations on Structural Design

• Estimate size of required spacecraft

• Estimate face occupancy (i.e., rough estimate of required placement for specific sensors)

• Does my payload have an aperture?

• What orientation does my vehicle need to be in?

• How can I get my solar panels maximally pointing?

• Where does the GPS antenna sit to get lock?

• Is my TT&C antenna angled well? Mission Design Expectations on Structural Design In order to refine the CONOPS and assess if the system is achievable, the platform and form-factor must be assumed. This will ultimately drive other elements too. Simplest method is to assume a size first (i.e. 3U, 6U, or 12U), then fill the box starting with the following: Continue through the exercise until the appropriate size and face occupancy is identified.

=== Slide 237: ShipSat Example

ShipSat Example * Baselining 12U due to current mission understanding – could change later * Lens is long and requires 3U length * Camera aperture points to ground * Solar panels face sun in highest power draw conditions * Patch antenna lobes should have good gain with respect to ground in most occupied state * Sun sensors aligned with solar generation faces for sun pointing * Star trackers point to sky and avoid keep out zones (i.e. sun) * GPS antenna points zenith as we are in LEO

=== Slide 238: Steps for Mass Budget Design

• Account for all components and structure

• Materials

• Coatings

• Make sure it is in the allowable mass range for CubeSats

• 3U ~ 6kg, 6U ~ 12 kg, 12U ~ 24 kg

• Have a 25% contingency in the design phase (this will be reduced over time) Steps for Mass Budget Design Component Estimated Mass Contingency Total Mass Component 1 100g 25% 125g Component 2 80g 25% 100g …. … … …

=== Slide 239: ShipSat Mass Budget

Component Identification Mass Contingency Total mass Solar Arrays 900g 25% 1125g Structure 10kg 25% 12500g Fasteners 189g 25% 236.25g Harness 924g 25% 1155g Power System 100g 25% 125g ADCS 1100g 25% 1375g Radio 450g 25% 563g Payload Imager 1600g 25% 2000g Heater(s) 50g 25% 63g CDH 250g 25% 313 g GPS 250g 25% 313g Max Allowed 24kg Total 19.766kg ShipSat Mass Budget

=== Slide 240: Mass/Volume Budget for Mission Design Course

• Similar to ShipSat example

• Simple mass budget with known components and contingency

• Simple volume budget to determine form factor

• Basic configuration diagrams/models

• Show faces/orientation of important external components

• CubeSat size Mass/Volume Budget for Mission Design Course Ensure form factor and configuration is feasible 3 solar arrays generally sun pointing, all deployable. 1 body- mounted on remaining face Star tracker orientation and location avoids earth and sun in field of view in nominal attitudes. Placed on this side with cutout in solar panel to ensure deployable doesn’t cover it Payload optic generally nadir pointing S-band antenna generally nadir pointing GPS antenna facing generally zenith on back face.

=== Slide 241: Budget Interdependencies

First round of budgeting showed: * Closing power budget requires CONOPS, TT&C, payload, and solar array adjustments * Closing data budget requires CONOPS, ground architecture, and TT&C adjustment Budget Interdependencies Power budget Link budget, communication design Data budget Experiment plan, CONOPS Mass and volume budgets Pointing budget

== ShipSat final presentation and guidance

=== Slide 242: ShipSat Final Presentation

ShipSat Final Presentation

=== Slide 243: ShipSat Mission Narrative

A city would like to build a dedicated space asset to track port traffic and occupancy on the city water bodies. Understanding vehicle traffic to and from the ports will help them identify improvement needs. The city intends to improve city infrastructure near high traffic areas in the next 5 years ShipSat Mission Narrative Microsoft PowerPoint stock images

=== Slide 244: ShipSat Mission Overview

Mission Statement * Count ships in specific areas from a space-based platform to determine consumer traffic trends in order to assess parking needs [and traffic flow]. Mission Objectives * MO-1: Identify ships within an image with 90% accuracy * MO-2: Image same location at least once daily * MO-3: Geolocate image to within 50 meters Mission Success Criteria * MSC: Obtain ship quantity for 1 location at the same time daily for 7 days * FSC: Obtain ship quantity for 5 locations at the same time daily for 49 days ShipSat Mission Overview

=== Slide 245: CONOPS - Example Experiment A

Choose image target Determine collect window Schedule collect on spacecraft Spacecraft is configured for collect Spacecraft executes collect Spacecraft store data Downlink data Data Analysis CONOPS - Example Experiment A

=== Slide 246: ShipSat Mode Details

ShipSat Mode Details Safe Mode * Used for fault recovery * Most subsystems off * Attitude: Tumbling Standby Mode * Nominal mode, used when experiments are not occurring * Most subsystems on, except for payload * Attitude: largest solar panel area pointing at sun Experiment Mode * Used for running experiments (taking images) * Most subsystems on * Attitude: Imager nadir pointing Certain faults (low battery) may autonomously demote straight to safe mode Certain faults (unresponsive imager) may autonomously demote to standby Promotion (commanded) Demotion to safer state (autonomous or commanded)

=== Slide 247: ShipSat Spacecraft Overview

• COTS camera and lens (constraint) drives resolution, requires 12U

• CubeSat size attitude system

• Solar arrays likely deployable (>40W)

• TT&C system with 1.5 Mbps downlink enables payload downlink

• CONOPS cannot support image collect on every pass ShipSat Spacecraft Overview 3 solar arrays generally sun pointing, all deployable. 1 body- mounted on remaining face Star tracker orientation and location avoids earth and sun in field of view in nominal attitudes. Placed on this side with cutout in solar panel to ensure deployable doesn’t cover it Payload optic generally nadir pointing S-band antenna generally nadir pointing GPS antenna facing generally zenith on back face.

=== Slide 248: Data Budget

UHF Radio (9600 bps) S-Band Radio (~1 Mbps) Data Budget Compares data storage in one day with UHF radio and S-Band radio

=== Slide 249: Case 1: Overhead

Case 1: Overhead Port size: 100m Allowable boresight error: 50m 600 km altitude View width: 19 km 0.9° Assuming our values are accurate, the overhead case will work! Minimum viewing angle (0°) Error Source Error Type Error Value Resulting Boresight Ground Error GPS Position 10m 10.0m Mechanical alignment Angle 1 arcsec 2.96m ADCS Angle 8 arcsec 23.66m Thermal Angle 3 arcsec 8.87m Total 45.49m

=== Slide 250: Case 2: Side Angle

Case 2: Side Angle Error Source Error Type Error Value Resulting Boresight Error in Normal Plane Resulting Boresight Ground Error GPS Position 10m 7.1 10.0 Mechanical alignment Angle 1 arcsec 3.9m 5.5 ADCS Angle 8 arcsec 31.6m 44.7 Thermal Angle 3 arcsec 11.8m 16.7 Total 76.9 Side angle is not acceptable Port Size: 100m 600 km altitude View width: 36 km 0.9° Maximum viewing angle (45°) Allowable boresight error on ground: 50m 45° 807 km range

=== Slide 251: Pointing Budget

Pointing Budget Maneuver Time Duration (s) Track Length (km) Image Boston 20 150 Slew to Detroit 15.5 116.25 Image Detroit 20 150 Slew to Norfolk 15 112.5 Image Norfolk 20 150 Slew to Charleston 6 45 Image Charleston 20 150 Slew to Pensacola 10 75 Image Pensacola 20 150 Slew to Miami 13 97.5 Image Miami 20 150 Total 179.5 1346.25 Available 290 2200 7.5 km/s ground speed Mission sequence is possible based on expected maneuver spacing Imaging: Maneuver:

=== Slide 252: Downlink (Space to Ground)

Transmitter Information Frequency 2.23 GHz Utilized Bandwidth 0.0125 MHz Wavelength 0.134 m Distance 1932 km Bit rate 1.000 Mbps Transmit Power 1.0 W Transmit Power 30 dBm Transmit antenna gain 6.5 dBi Transmit system losses -4 dB EIRP 32.5 dBm Path loss -165.14 dB Downlink (Space to Ground) Key Derived value Value to Enter Preliminary link completed for S-band system with unacceptable margin Orbit Information Elevation Angle 10 ° Altitude of Satellite 600 km Radius of Earth 6378 km Speed of Light 299,792,458 m/s Boltzmann Constant 1.38E-23 W/K/Hz Distance 1932 km Receiver Information Receive antenna diameter 5 m Antenna Efficiency 60 % Receive antenna gain 39.13 dBi Noise Temperature 307.0 K Receive system noise figure -173.73 dBm/Hz Receiver System losses -5 dB Total Received Power -96.6 dBm Cross polarization loss -3 dB Link Margin Computation Received Eb/No 12.2 dB Required Eb/No 11.0 dB Link margin 1.12 dB Required link margin 6 dB Values to enter may come from data sheets, specific implementation choices, CONOPS, and more Link budget taken from Space Mission Engineering - Wertz

=== Slide 253: Energy Balance

Energy Balance * Chosen Configuration: 3 Panels on Same Face * Modeling with original mode sequence * 4 orbits safe * 4 orbits standby * 8 orbits experiment * Results * Power positive across all modes * Initial design closes, future work should critically evaluate assumptions Safe Standby Experiment Mode Orbit Avg Power Consumed (W) Orbit Avg Power Generated (W) Safe 4.33 4.8 Standby 7.83 21.6 Experiment 10.83 14.4

=== Slide 254: ShipSat Mass Budget

ShipSat Mass Budget Component Identification Mass Contingency Total mass Solar Arrays 900g 25% 1125g Structure 10kg 25% 12500g Fasteners 189g 25% 236.25g Harness 924g 25% 1155g Power System 100g 25% 125g ADCS 1100g 25% 1375g Radio 450g 25% 563g Payload Imager 1600g 25% 2000g Heater(s) 50g 25% 63g CDH 250g 25% 313 g GPS 250g 25% 313g Max Allowed 24kg Total 19.766kg

=== Slide 255: Further Guidance

• Template provided, but make it your own

• The goal is to develop as far as you can

• No baseline expectation

• Good approach is more important than level of development

• Have fun with it too! Further Guidance Credit XKCD: xkcd.com/1992