ECE 521E — Satellite Subsystems: Extensive Study Notes

Section 1

1. Overview of a Satellite System

1.1 The Two Halves of Every Satellite

Any satellite, whether a 1U CubeSat or a 6-tonne GEO communications spacecraft, divides into two functional halves:

The bus (platform): everything that keeps the satellite alive and correctly oriented — structure, power, thermal, attitude control, propulsion, TT&C, command and data handling. The bus is essentially identical in function across missions.
The payload: the mission equipment — transponders and antennas for a communications satellite, imaging radar or radiometers for an Earth-observation satellite, navigation signals for GNSS, or scientific instruments. The payload defines the mission; the bus supports it.

Design principle

The payload is designed first (mission requirements → payload sizing), and the bus is then sized to carry and support the payload. Bus mass and power typically scale with payload mass and power through simple fractions: on GEO comsats the bus is roughly 25–40% of launch mass; power subsystem mass ≈ 2.5–4 kg per 100 W of generated power.

1.2 Classification of Satellites

Class	Orbit altitude	Period	Typical missions
LEO (Low Earth Orbit)	300 – 1,500 km	~90 – 110 min	Earth observation, SAR, ISS, imaging, CubeSats, LEO broadband constellations
MEO (Medium Earth Orbit)	~8,000 – 20,000 km	~6 – 12 h	GPS/GNSS (20,200 km), navigation
GEO (Geostationary)	35,786 km (equatorial)	23 h 56 min (sideral)	Fixed comms, weather imaging, broadcast
HEO / elliptical	perigee ~500 km, apogee ~40,000 km	varies (e.g., 12 h Molniya)	High-latitude communications (Molniya), early-warning
Sun-synchronous LEO	~600 – 800 km, retrograde, i ≈ 98°	~100 min	Mapping, radar altimetry — constant local solar time (e.g., 10:30 AM descending node)

1.3 The Subsystem Map

Standard satellite engineering recognizes these bus subsystems:

Mechanical structure and mechanisms
Attitude Determination and Control Subsystem (ADCS)
Propulsion subsystem
Electrical Power Subsystem (EPS)
Thermal Control Subsystem (TCS)
Telemetry, Tracking and Command (TT&C)
Command and Data Handling (C&DH)
Payload(s) with their dedicated antennas/sensors

Sections 2–8 treat the bus subsystems in detail; Sections 9–10 treat communication and radar payloads; Section 11 covers reliability engineering; Section 12 the ground segment.

1.4 The Space Environment (Why Subsystems Exist)

Vacuum: no convection — heat leaves only by radiation; cold-welding of bare metals; outgassing of materials.
Radiation: trapped Van Allen belts (electrons/protons), solar particle events, cosmic rays — cause single-event upsets (SEU), total ionizing dose (TID), and sensor damage; drives shielding, parts selection, and error correction.
Thermal extremes: eclipse transitions of ~−150 °C to +150 °C every orbit for LEO; continuous solar input for GEO (with eclipse seasons near equinoxes).
Microgravity: fluids behave by surface tension; propellant must be settled by thrust or surface-tension devices (PMDs) before firing.
Atomic oxygen (LEO): erodes polymers and silver coatings.
Micrometeoroids and debris: shielding (Whipple shields) and collision-avoidance manoeuvres.

Section 2

2. Mechanical Structure and Mechanisms

2.1 Functions

Carry launch loads (quasi-static 3–6 g; random vibration 14–20 g rms at component level; acoustic pressure inside fairing).
Provide stiffness — first natural frequency must exceed the launcher requirement (often > 100 Hz longitudinal, > 50–70 Hz lateral) to avoid dynamic coupling.
Provide mounting, alignment and thermal conduction paths for all subsystems.
Support deployment: solar arrays, antennas, radiators, instrument booms, covers.

2.2 Construction

Materials: aluminium honeycomb panels with carbon-fibre face sheets (high stiffness-to-mass), titanium fittings, CFRP (carbon-fibre reinforced polymer) for stable optical benches — near-zero coefficient of thermal expansion (CTE) for imaging payloads.
Configuration types: box/module (most comsats), central cylinder with shear panels (spins or three-axis), frame-and-panel, and for constellations stacked flat-plate satellites.
Mechanisms: hold-down and release mechanisms (pyrotechnic bolts, non-pyrotechnic HDRMs, burn wires for CubeSats), deployment hinges with dampers, gimbals for steerable antennas/solar arrays, momentum wheels and reaction wheels, rotary joints for despun antenna platforms.

Mechanism design rule: mechanisms are among the highest-risk items (they must work exactly once, after months in space). Design practice demands redundancy of release initiators, generous margins, ground testing in vacuum/thermal-vacuum, and where possible "test as you fly" on the flight unit.

Section 3

3. Attitude Determination and Control Subsystem (ADCS)

3.1 Mission of the ADCS

The ADCS keeps the satellite pointed (attitude control) and knows where it is pointing (attitude determination). Typical pointing requirements:

Mission	Pointing accuracy	Stability
GEO comsat antenna beam	±0.05° – 0.1°	0.001–0.005 °/s
Earth-imaging camera	±0.01° – 0.1°	very low jitter
SAR antenna	±0.1° – 0.5°	attitude knowledge drives image registration
Space telescope (e.g., JWST)	milli-arcseconds	extreme stability
CubeSat (basic)	±1° – 5°	modest

3.2 Attitude Determination Sensors

Sensor	Principle	Accuracy	Notes
Earth horizon sensors (IR)	Earth IR limb at 14–16 µm	0.1° – 1°	LEO standard; fails beyond GEO distance
Sun sensors	Photodiode fine/course heads	0.01° – 1°	Cheap, robust; ambiguous axis along Sun line
Star trackers	Image star field, pattern match	arcseconds	Best accuracy; occultation by Sun/Earth/Moon; radiation upsets
Magnetometers	Measure local B field	0.5° – 2°	Needs accurate field model (IGRF); LEO only (strong field)
Gyroscopes (rate sensors)	Momentum / optical (FOG, RLGs)	rate: 0.001°/h – 1°/h drift	Propagate attitude between fixes; MEMS for CubeSats
GPS receiver	Carrier phase / code	position m-level; attitude with multi-antenna	LEO mainly; increasingly used for orbit determination

Standard practice: gyros propagate the attitude between absolute fixes (star tracker / Earth sensor / Sun sensor) — a complementary filter (Kalman filter) merges them, estimating gyro bias.

3.3 Attitude Control Actuators

Actuator	Principle	Typical torque	Use
Reaction wheels	Exchange momentum with body (wheel accelerates)	0.01 – 5 N·m	Precision 3-axis control; limited by wheel saturation
Momentum wheels	Spin at biased speed for stiffness (gyroscopic)	as reaction wheels	Dual-spin GEO comsats, pitch momentum bias
Control Moment Gyros (CMGs)	Gimballed spinning wheel redirects momentum	50 – 500+ N·m per unit	Agile large spacecraft (ISS, imaging sats)
Magnetic torquers (torque rods)	Coil interacts with Earth's B field: τ = m × B	0.001 – 1 N·m	LEO momentum dumping, desaturation; no propellant
Thrusters (cold gas / chemical)	Jet momentum	0.01 – 20 N	Large slews, orbit corrections, wheel desaturation
Gravity-gradient boom	Tidal torque aligns long axis with nadir	passive	Simple LEO sats; accuracy ~1–5°

3.4 Control Modes and Disturbance Torques

Typical control modes: detumble after separation → acquisition (Sun, Earth) → nominal pointing → manoeuvre → safe mode (Sun-pointing, minimal activity, awaiting ground command).

Disturbance torques to be rejected (LEO example magnitudes for a 1 m³, 100 kg satellite at 500 km):

Torque source	Order of magnitude	Character
Gravity gradient	10⁻⁵ – 10⁻³ N·m	Restoring; used passively
Aerodynamic drag (VLEO)	10⁻⁵ – 10⁻³ N·m	Peaks at low perigee; variable
Solar radiation pressure	10⁻⁷ – 10⁻⁵ N·m	Steady; large for high area/mass
Magnetic residual dipole	10⁻⁶ – 10⁻⁴ N·m	Interacts with B field
Micrometeorite impacts	impulsive, rare	~10⁻⁸ N·m average

3.5 Momentum Budget and Desaturation

Wheels absorb cyclic disturbance momentum (gravity-gradient, magnetic residuals) but secular disturbance (solar pressure on an asymmetric body) accumulates until wheels saturate. The momentum capacity needed over time T:

H_req ≥ T_max · T (worst-case disturbance torque × duration)

Desaturation ("unloading") options: magnetic torquers in LEO (τ = m × B with dipole m aligned to oppose wheel momentum), or thruster pairs. GEO spacecraft unload with thrusters every few days; LEO CubeSats use torque rods continuously.

Exam favourite: A reaction wheel can only exchange momentum — it cannot change the spacecraft's total angular momentum. External torques (thrusters, magnetic, gravity gradient) are mandatory for secular momentum management.

Section 4

4. Propulsion Subsystem

4.1 Functions of On-Board Propulsion

Orbit acquisition: apogee/perigee raising, plane change (LEO→GEO GTO apogee burns; inclination change is famously expensive: Δv ≈ v·Δi for small angles).
Orbit maintenance / station-keeping: GEO E-W (luni-solar ~1–2 m/s per year) and N-S (~40–55 m/s per year, i.e., inclination drift) corrections; LEO drag makeup (VLEO: up to 10+ m/s/day); constellation phasing.
Attitude control: reaction control thrusters, momentum unloading.
End-of-life disposal: GEO re-orbit to graveyard (+300 km, typically 3–11 m/s depending on area/mass); LEO passivation/de-orbit within 25 years (guideline).
Collision avoidance: conjunction screening manoeuvres.

4.2 Thruster Types

Type	Propellant / principle	Isp (s)	Thrust	Use
Cold gas	N₂, Ar, CO₂ stored pressurized	50 – 75	mN – N	Simple RCS, CubeSats
Solid (Vernier)	Once-only, e.g., apogee kick stage	280 – 300	kN	AKM for GEO insertion
Monopropellant hydrazine	N₂H₄ over catalyst bed	220 – 235	0.5 N – 22 N (0.9 lbf class)	Workhorse station-keeping/RCS
Bipropellant (MMH/NTO)	Hypergolic ignite on contact	300 – 325	10 N – 450 N	Apogee engines (490 N), larger manoeuvres
Electrothermal/resistojet	Heated propellant	300 – 350	mN	Small Δv, long life
Electric propulsion	Hall thrusters, ion (gridded), PPT	1,500 – 4,000+	mN – 0.5 N	High-Δv missions, orbit raising (all-electric GEO), station-keeping

4.3 The Rocket Equation (Tsiolkovsky)

Δv = g₀ · Isp · ln(m₀ / m_f) = g₀ Isp · ln(1 + m_p/m_f)g₀ = 9.81 m/s²; m₀ = initial mass; m_f = final (dry) mass; m_p = propellant mass

Solving for propellant fraction:

m_p/m₀ = 1 − exp(−Δv / (g₀ Isp))

Example: station-keeping a GEO comsat for 15 years needs Δv ≈ 550 m/s north-south + east-west. With Isp = 220 s (hydrazine): m_p/m₀ = 1 − exp(−550/2158) ≈ 0.226 — nearly a quarter of launch mass is propellant. With Isp = 3000 s (electric): m_p/m₀ = 1 − exp(−550/29,430) ≈ 0.019 — ~2% of mass. This single comparison explains the modern shift to electric orbit raising and station-keeping.

4.4 Propellant Storage and Feed

Surface-tension propellant management devices (PMDs — gallery arms, vanes, sponges) trap liquid at the tank outlet for bubble-free expulsion in weightlessness.
Pressurant (He) with regulators; diaphragm or bladder tanks for smaller systems; surface tension tanks (COTS for CubeSats).
Propellant gauging is notoriously hard: bookkeeping (PvT), ultrasonic/RF sensors, or thrust-time integration.

Section 5

5. Electrical Power Subsystem (EPS)

5.1 Architecture

The EPS generates, stores, regulates, and distributes electrical power:

Generation: photovoltaic solar arrays (essentially universal for Earth-orbit satellites).
Storage: rechargeable batteries for eclipse and peak loads.
Power conditioning: regulation to unregulated or regulated bus voltages (28 V, 50 V, 100 V class), MPPT or direct-energy transfer, battery charge control.
Distribution: switchgear, fusing, load shedding, power telemetry.

5.2 Solar Array Sizing

Solar cell efficiency (modern triple-junction GaInP/GaAs/Ge cells ≈ 29–32% BOL, degrading ~2–3% absolute over 15 years in GEO). Array power at beginning of life (BOL):

P_BOL = P_sun · η_cell · A · cos θ · K_tempP_sun ≈ 1361 W/m² solar constant; A = cell area; θ = Sun incidence angle; K_temp ≈ 0.9 (temperature derating)

End-of-life (EOL) power after radiation degradation and contamination: P_EOL = P_BOL · (1 − d)^N with yearly degradation d (0.5–1%/yr typical for triple-junction, more for Si). Arrays are sized for EOL conditions at worst Sun angle.

Array sizing example

A LEO satellite needs 500 W EOL at 600 km. Cells: η = 30%, degradation 1%/yr over 5 yr, worst-case Sun angle 23° off-normal (cos = 0.92), temp derate 0.9, packing factor 0.9.

P_EOL = 1361 × 0.30 × A × 0.92 × 0.9 × 0.9 × (0.99)⁵ = 1361 × 0.30 × A × 0.92 × 0.9 × 0.9 × 0.951
Required: A = 500 / (1361×0.30×0.92×0.9×0.9×0.951) = 500 / 289 ≈ 1.73 m² of cell area — roughly 2 m² with engineering margin.

5.3 Batteries and Eclipse Design

Chemistry: historically NiCd (robust, memory-free with proper cycling); NiH₂ (long GEO life, ~30–60% DoD operation); today predominantly Li-ion (high specific energy 100–150 Wh/kg; managed DoD 20–40% for cycle life, low-Earth-orbit sats endure ~35 min eclipse every 90–100 min ≈ 5,500 cycles/yr).
Capacity sizing: C = P_eclipse · t_eclipse / (V · DoD · η_disch).
LEO eclipse: worst-case ~35–36 min with 16 eclipses/day; GEO: up to 72 min during equinox seasons, none otherwise.
Battery heaters and charge regulators (CC/CV taper charging) essential; Li-ion needs cell balancing and strict thermal limits (0–20 °C charge range typical).

C&DH link: the EPS is managed by the power control unit under C&H software commands: array pointing (if articulated), battery charge modes, load shedding priorities (payload first to shed, survival heaters last).

Section 6

6. Thermal Control Subsystem (TCS)

6.1 The Governing Physics

In vacuum the only heat transfer modes are conduction (within structure) and radiation (to space). Steady-state balance for each node:

α·S·A_proj + ε·σ_SB·A·T_env⁴ + Q_int = ε·σ_SB·A·T^⁴ + Q_condα = solar absorptivity; ε = infrared emissivity; σ_SB = 5.67×10⁻⁸ W/m²K⁴; Q_int = internal dissipation

Design targets are typically −10 °C to +50 °C for electronics, ~0–40 °C for batteries, and cryogenic or tightly stable temperatures for some sensors. The ratio α/ε is the master thermal design variable: low α/ε (white paint, optical solar reflectors) keeps equipment cool; high α/ε (black paint) keeps it warm.

6.2 Passive Thermal Control

Surface coatings/films: optical solar reflectors (OSR, α/ε ≈ 0.1/0.8), white paint (α/ε ≈ 0.2/0.85), black paint (α/ε ≈ 0.95/0.9), silvered Teflon (α/ε ≈ 0.08/0.8).
Multi-layer insulation (MLI): 10–20 layers of aluminized Kapton/Mylar with netting; effective emittance ~0.01–0.03 — blankets propellant tanks, lines, and the anti-Sun faces.
Radiators: dedicated high-emittance panels (OSR, white paint) rejecting internal heat; sized from Q_int = εσA(T⁴ − T_space⁴) with T_space ≈ 4 K background (or ~250 K for Earth-facing views — never radiate to Earth if you want cold).
Thermal doublers, heat pipes, louvers: spread or throttle heat without power.

6.3 Active Thermal Control

Electric heaters (survival heaters on batteries, propellant lines, thruster catalyst beds) with thermostats or software PID.
Mechanically pumped fluid loops (large platforms, crewed vehicles; ammonia/water loops).
Heat pipes / loop heat pipes: two-phase capillary devices with effective conductivity 10,000× copper — standard in modern satellites.
Louvers: bimetallic blades modulating radiator view factor (varying ε_eff ~0.2–0.8).
Cryocoolers / radiative coolers: for IR instruments (Stirling, pulse-tube cryocoolers to 20–80 K).

Typical sizing calculation: a 300 W internal dissipation rejected from a 2 m² radiator at ε = 0.85 toward 3 K space: T = [300/(0.85×5.67e-8×2) + 3⁴]^¼ ≈ [3.11e9]^¼ ≈ 237 K (−36 °C) — comfortable margin above freezing for most electronics; add heater power for eclipse cold-soak cases.

Section 7

7. Telemetry, Tracking, and Command (TT&C)

7.1 The Three Functions

Telemetry (downlink): satellite health and status — voltages, currents, temperatures, pressures, attitudes, event counters — via beacon or the main comm link.
Command (uplink): discrete and serial commands from ground to configure, switch, point, and recover the spacecraft; command protection via authentication/encryption (e.g., CCSDS Space Data Link Security protocol).
Tracking: range (two-way tone/pseudonoise delay), range-rate (Doppler), and angle (monopulse at the ground antenna) — combined in orbit determination to produce ephemerides.

7.2 Frequencies and Link Budget

Band	Uplink (MHz)	Downlink (MHz)	Notes
VHF	148 – 150	137 – 138	CubeSats, beacons; congested
UHF	435 – 438	400 – 402	CubeSat command; AIS/ARGOS
L	1,631.5 – 1,626.5 (LEO)	1,525 – 1,559	Mobile sat, GNSS; GPS L1 1575.42
S	2,025 – 2,120	2,200 – 2,290	TT&C workhorse, NASA/ESA deep space (DSN S-band)
X	7,145 – 7,235	8,400 – 8,500	Deep-space uplink/downlink; SAR payloads
Ku	14.0 – 14.5	10.7 – 12.75	FSS/VSAT; DTH broadcast downlink
Ka	27.5 – 31	17.7 – 21.2	High-throughput beams, gateway links

7.3 Transponder (Bent-Pipe) Architecture

1RF front end: antenna (often a low-gain hemispherical/omni for TT&C plus medium-gain horn) → diplexer → low-noise amplifier (LNA) → down-converter (mixer + local oscillator).

2Channelization: input demultiplexer (IMUX) filter bank splits the wide receive band into transponder channels (36 MHz or 72 MHz typical for GEO comsats); each channel has automatic gain control and a linearized (TWTA or SSPA) high-power amplifier.

3Output: output multiplexer (OMUX) combines channels → high-power amplifier (HPA) → output filter (to suppress harmonics/out-of-band emissions) → transmit antenna via diplexer.

Regenerative (processing) transponders demodulate on board and re-modulate — enabling on-board switching (SS/TDMA), beam routing, and interference isolation at the cost of complexity.

7.4 Tracking Measurements

Two-way range: R = c·τ/2 with tone ranging accuracy ~few metres (phase ambiguity resolved by code).
Doppler range-rate: f_d = −(2f₀/c)·Ṙ (two-way) → m/s-level accuracy with 1 s count.
Angle tracking: monopulse (simultaneous lobing) at the ground station, ~0.01–0.05° (1σ) for 13-m antennas.
Modern orbit determination fuses these with GNSS carrier-phase on LEO missions, achieving cm–dm accuracy.

Section 8

8. Command and Data Handling (C&DH)

8.1 Role

The C&DH subsystem is the spacecraft's brain: it receives and validates commands, executes them autonomously, collects and formats telemetry, manages the onboard computer (OBC), mass memory, and the data buses, and implements autonomy (safe modes, fault detection, isolation and recovery — FDIR).

8.2 Typical Architecture

On-board computer (OBC): radiation-tolerant processors (e.g., LEON/SPARC V8, RAD750 PowerPC) with EDAC-protected memory, typically 100–500 MHz with tens of MB of RAM (plus SDRAM for payload buffering).
Data buses: CAN (CubeSats, ESA's CANopen; 1 Mbit/s), MIL-STD-1553B (legacy, robust, 1 Mbit/s, command/response), SpaceWire (high-rate point-to-point, up to 200+ Mbit/s for payloads), CCSDS SpaceWire/Low-voltage differential (LVDS) links for instruments.
Mass memory: NAND flash stores (tens of GB to several TB for EO constellations), with wear-leveling and bad-block management; downlink scheduled when in ground-station view.
Software: real-time operating systems (RTEMS, VxWorks, Linux for CubeSats), layered autonomy with timeouts on every command chain.

8.3 Telemetry and Command Standards

CCSDS (Consultative Committee for Space Data Systems) frame/packet standards: TM/TC Space Data Link Protocols, AOS, Space Packet Protocol — international interoperability.
Encoding: convolutional + Reed–Solomon concatenated codes, Turbo codes, LDPC (CCSDS); ranging and beacon formats per CCSDS.
Time standard: spacecraft elapsed time (MET) coordinated with UTC via ground.

FDIR hierarchy: Level 1 — hardware watchdog and component retries; Level 2 — autonomous reconfiguration to redundant units (cross-strapping); Level 3 — safe mode (Sun acquisition, spinning or 3-axis, minimal power, await ground). Every fault path must be testable on the ground.

Section 9

9. Communication Payloads in Detail

9.1 Payload Elements

Antennas: global-coverage horns, shaped reflectors (contoured beams), phased arrays (multiple steerable beams; HTS sats carry 50–200+ beams), feeds with low passive intermodulation (PIM) design.
High-power amplifiers: traveling-wave tube amplifiers (TWTAs, 100–250 W at Ku, ~50% efficiency) vs. solid-state power amplifiers (SSPAs, GaN-based, 30–60% efficiency, linearity better, mass lower at lower power).
Channel routing: static (channelized), switchable (IF switch matrix), or full on-board processing (OBP) with demod/remod (regenerative).

9.2 The Satellite Link Budget

The payload performance is expressed through EIRP (uplink: satellite receive G/T; downlink: satellite EIRP):

(C/N₀)_downlink = EIRP_sat − L_fs − L_atm − L_pointing + (G/T)_ground − k (all in dB)EIRP = P_t + G_t (dBW + dBi); L_fs = 92.45 + 20log f(GHz) + 20log R(km); k = −228.6 dBW/K/Hz

Overall link quality combines uplink, downlink, and intermodulation/interference:

(C/N₀)_total⁻¹ = (C/N₀)_up⁻¹ + (C/N₀)_down⁻¹ + (C/N₀)_IM⁻¹ + (C/I)⁻¹ (powers in linear units)

Required C/N₀ from BER: e.g., QPSK with BER 10⁻⁵ needs Eb/N₀ ≈ 9.6 dB; C/N₀ = Eb/N₀ + 10log R_b. Add coding gain (Turbo/LDPC ≈ 5–7 dB at BER 10⁻⁵) and implementation margin (1–2 dB).

9.3 Multiple Access

Technique	Principle	Example
FDMA	Channels by frequency	Classic C-band transponders
TDMA	Slots in time, burst carriers	Intelsat business service (120 Mbit/s)
CDMA	Spread codes	GPS (CDMA), some LEO constellations
SDMA	Spatial beams	HTS spot-beam frequency reuse (4-color reuse)

Frequency reuse multiplies capacity: spot beams with orthogonal polarization/frequency reuse factor 4–8 give total throughput = per-beam throughput × reuse factor (beam isolation > ~16–20 dB required to keep C/I acceptable).

Section 10

10. Radar and Earth-Observation Payloads

10.1 Why Radar Payloads

Microwave radar provides day/night, all-weather imaging (cloud-penetrating at L/C-band), precision ranging (altimetry), and wide-swath ocean/ice monitoring — complementing passive optical sensors. Spaceborne radar must live within tight mass, power, and data-rate budgets; hence the heavy use of SAR (synthetic aperture radar) to achieve fine azimuth resolution without a physically long antenna.

10.2 Synthetic Aperture Radar (SAR) Payload

Key relationships

Azimuth resolution: δ_az ≈ L/2 (independent of range — the SAR advantage)L = physical antenna length along track

Range resolution: δ_r = c / (2B·sin γ)B = chirp bandwidth; γ = incidence angle; pulse compression (chirp) achieves metres with μs pulses

Peak power need: P_t via the radar equation (two-way) with σ = σ⁰·δ_rδ_az·(resolution cell)σ⁰ = normalized backscatter coefficient

Payload subsystems

RF front end: chirp generator (digital, DDS), up-conversion to L/C/X/Ku band, TWTA or GaN SSPA (hundreds of watts to a few kW peak, duty cycle 5–20%).
Antenna: planar phased array (active TR modules — e.g., TerraSAR-X: 384 T/R modules, X-band, 300 km swath StripMap), or reflector-fed for altimetry.
Receiver/data: low-noise receiver, ADC at IF/baseband, real-time raw-data buffering (multi-Gbit/s), onboard range compression sometimes, solid-state recorder for 100–1,000 km scenes.
Duty/power: SAR is power-hungry (1–4 kW payload power) — scheduling and thermal design must average over the orbit duty cycle.

Modes: StripMap (fixed pointing), ScanSAR (steered elevation bursts for wide swath, coarser azimuth), Spotlight (steered to dwell on target — sub-metre to decimetre resolution, e.g., ~25 cm commercial products).

10.3 Other Radar Payloads

Payload	Principle	Application
Radar altimeter	Nadir pulse timing: h = cτ/2; 1–2 cm precision with tracking gates	Ocean topography (Jason/TOPEX), ice-sheet elevation
Scatterometer	Normalized σ⁰ vs incidence angle at multiple azimuths	Sea-surface wind speed/direction (ASCAT, QuikSCAT)
Cloud/rain radar	W/Ka-band profiling of reflectivity	CloudSat (94 GHz Cloud Profiling Radar), GPM DPR
GNSS reflectometry	Bistatic use of L1/L5 signals reflected off Earth's surface	Soil moisture, ocean winds (CYGNSS)

Link to ECE 521E radar equation material: the SAR/altimeter sizing problems are direct applications of the two-way radar equation — received power ∝ R⁻⁴ forces kW-class peak powers, pulse compression for average-power reduction, and large antennas (T/R module count) for gain — the same trade-offs studied in the radar-equation notes.

Section 11

11. Reliability, Redundancy, and Parts Engineering

11.1 Quantitative Reliability

R(t) = exp(−λt)constant failure rate λ; reliability function (fraction surviving)

P_s(t) = Σ over redundant chains; for two parallel identical units: P_s = 1 − (1 − R)²

Mission requirement drives design: a GEO comsat needing 0.9 probability of 15-year life with inherently ~0.7 reliability demands redundancy in every long-life chain (TWTAs ×2 with loop/backup switching, receivers, computers, gyros ×4 in 2×2).

11.2 Redundancy Strategies

Strategy	Description	Penalty
Cold (standby) spare	Spare unpowered until needed	Slower switchover; dormant-failure risk
Warm/hot redundancy	Spare powered, tracking, instant switch	Power/mass; cross-strapping complexity
Cross-strapping	Any unit to any function via switch matrix	Relays/switches add their own failures
N-of-M voting	Majority logic (e.g., 2-of-3 gyros)	Extra mass; voter must be ultra-reliable

11.3 Parts, Materials, and Processes (PM&P)

Radiation-hardened (rad-hard) or radiation-tolerant parts; SEE screening; derating (voltage, current, junction temperature 10–20 °C margins below rating).
TID budget: internal shielding 2–4 mm Al equivalent typical; parts qualified to mission dose (LEO 1–10 krad(Si); GEO/MDE 10–100 krad(Si)).
Outgassing limits (CVCM < 0.1%, TML < 1.0% per ASTM E595) for all polymers in vacuum.
Lot traceability, ESD control, and destructive physical analysis (DPA) for critical lots.

Section 12

12. Ground Segment and Operations

12.1 Elements

Ground stations: tracking antennas (5–13 m for LEO TT&C/TDRSS services; 13–70 m class for deep space — DSN 34 m/70 m), redundant chains, weather/all-weather operation.
TT&C network services: NASA TDRSS (GEO data relay over S/Ku/Ka), ESA ESTRACK, commercial networks (KSAT, SSC) — buy contact time, minimize own stations.
Mission control centre (MCC): flight dynamics (orbit determination & prediction), spacecraft controllers, procedures, simulation, 24/7 staffing for critical phases.
Payload data systems: reception, processing, archiving, and distribution for EO data; level 0 (raw) → level 1 (calibrated products) → level 2+ (geophysical).

12.2 Contact Geometry (LEO)

Pass duration and elevation determine link time: a 600 km orbit with a 5° minimum elevation gives ~8–10 min contacts and 2–4 passes/day from a single mid-latitude station; global coverage needs a relay constellation or multiple ground stations.

Max slant range at elevation E: R_max = √( (R_e+h)² − (R_ecos E)² ) − R_e sin ER_e = 6371 km; h = altitude; at E = 0°, R_max ≈ √(2R_eh) for small h

Section 13

13. Worked Design Examples

Example 1 — GEO comsat EIRP and G/T check

A Ku-band transponder: TWTA 120 W (20.8 dBW), antenna transmit gain 32 dBi, path 35,786 km at 11.95 GHz, ground station G/T = 30 dB/K. Downlink C/N₀ for C/T of −160 dBW/K/Hz?

EIRP = 20.8 + 32 = 52.8 dBW.
L_fs = 92.45 + 20log(11.95) + 20log(35786) = 92.45 + 21.55 + 91.07 = 205.1 dB.
C/N₀ = 52.8 − 205.1 − 0.5 (atmos) − 0.3 (pointing) + 30 − (−228.6) = 105.5 dBHz.
For a 36 MHz channel carrying 54 Mbit/s QPSK (27 MBaud), required C/N₀ = Eb/N₀ (9.6 + coding 5 ≈ 5 dB) + 10log(27×10⁶) = 10.6 + 74.3 = 84.9 dBHz → margin ≈ 20 dB. ✓ Comfortable.

Example 2 — Hydrazine vs electric station-keeping

15-year GEO life needs Δv = 550 m/s. Compare propellant mass fraction for Isp = 220 s vs 3000 s.

Hydrazine: 1 − exp(−550/(9.81×220)) = 1 − exp(−0.255) = 22.5%.
Electric: 1 − exp(−550/(9.81×3000)) = 1 − exp(−0.0187) = 1.85%.
A 3,000 kg GEO satellite saves ≈ 3,000 × (0.225 − 0.0185) ≈ 620 kg — traded against months of electric orbit-raising time.

Example 3 — Reaction wheel momentum sizing (LEO)

A 200 kg Earth-observation satellite suffers solar-pressure residual torque 2×10⁻⁴ N·m (LEO example exaggerated to include aero). Wheels must store one orbit of secular momentum (94 min) with 50% margin. Required wheel momentum?

H = τ·T = 2×10⁻⁴ × 5640 s = 1.13 N·m·s; with 50% margin: H_req ≈ 1.7 N·m·s.
A single 2 N·m·s wheel saturates before one orbit → use 4 wheels in pyramid (2 N·m·s each, sharing 1.4 N·m·s total) and unload with torque rods every orbit. This is exactly how 10–15 kg CubeSats are also designed (scaled down).

Example 4 — Battery sizing for LEO

Payload+bus draw 350 W during 35 min eclipse at 28 V bus; allowed DoD 30%; discharge efficiency 0.95. Capacity?

Energy needed: E = P·t = 350 W × 0.583 h = 204 Wh.
C = E / (V·DoD·η) = 204 / (28 × 0.30 × 0.95) = 204 / 7.98 ≈ 25.6 Ah → a single 30 Ah Li-ion string (with cell balancing), or two 30 Ah strings for redundancy; BOL capacity margin ~15% covers aging over the mission life.

Example 5 — SAR payload power sanity check

StripMap SAR at 700 km, X-band (9.65 GHz), 40 MHz chirp bandwidth, antenna 4.8 m × 0.7 m (TerraSAR-X-like), peak Tx 3.3 kW. Rough link margin logic: resolution cell σ⁰·δr·δaz with σ⁰ ≈ −10 dB, δr = 3 m, δaz = 3 m → σ = 0.1×9 ≈ −0.5 dBsm. Accepting the computed SNR ≈ 15–20 dB from the radar equation for this class, the design point is consistent with published SAR mission parameters — always verify against the radar equation as taught in the radar-equation sections.

Section 14

14. Formula and Design-Rule Sheet

#	Relation	Use
1	Δv = g₀·Isp·ln(m₀/m_f)	Propellant mass from manoeuvre budget
2	m_p/m₀ = 1 − exp(−Δv/g₀Isp)	Propellant fraction
3	P_BOL = 1361·η·A·cosθ·K_temp	Solar array sizing
4	C_bat = P·t_ecl/(V·DoD·η)	Battery capacity
5	Q = εσ_SBA(T⁴ − T_space⁴)	Radiator sizing
6	τ = m × B	Magnetic torque (LEO desaturation)
7	H_wheel ≥ τ_dist·T	Momentum storage requirement
8	L_fs = 92.45 + 20logf_GHz + 20logR_km	Free-space path loss
9	C/N₀ = EIRP − L + G/T − k	Link budget (k = −228.6 dBW/K/Hz)
10	(C/N₀)_tot⁻¹ = Σ(C/N₀)_i⁻¹	Combining uplink/downlink/IM/interference
11	R_slant,max = √( (R_e+h)² − (R_ecosE)² ) − R_esinE	Ground-contact geometry
12	δ_az = L/2; δ_r = c/(2B sinγ)	SAR resolutions
13	R(t) = e^−λt; two-parallel: P_s = 1 − (1−R)²	Reliability math
14	EIRP = P_t + G_t; G/T = G − 10log T_s	Payload figures of merit

Section 15

15. Glossary of Key Terms

Term	Definition
Bus / platform	The satellite structure and supporting subsystems that carry and service the payload
Payload	Mission equipment (transponders, sensors) that the satellite exists to operate
ADCS	Attitude Determination and Control Subsystem
BOL / EOL	Beginning / End of Life (array power, reliability conditions)
TTC / TT&C	Telemetry, Tracking and Command — the satellite's link to the ground
Bent-pipe transponder	Repeater that translates and amplifies signals without demodulation
EIRP	Effective isotropic radiated power, P_t·G_t (dBW + dBi)
G/T	Figure of merit: antenna gain over system noise temperature (dB/K)
FDIR	Fault Detection, Isolation and Recovery (autonomy)
MLI	Multi-layer insulation (thermal blanketing)
OSR	Optical solar reflector (low α/ε radiator surface)
PMD	Propellant management device (surface-tension liquid acquisition)
CCSDS	Consultative Committee for Space Data Systems (standards body)
SEU / TID	Single-event upset / total ionizing dose (radiation effects)
Safe mode	Minimal autonomous configuration awaiting ground recovery
Station-keeping	Small manoeuvres maintaining assigned orbital box

References

J. J. Pocha, An Introduction to Mission Design for Geostationary Satellites, Springer, 1987.
M. J. Sidi, Spacecraft Dynamics and Control: A Practical Engineering Approach, Cambridge, 1997.
W. J. Larson & J. R. Wertz (eds.), Space Mission Analysis and Design, 3rd ed., Microcosm, 1999.
G. Maral & M. Bousquet, Satellite Communications Systems, 5th ed., Wiley, 2009.
I. G. Ngoma (ed.), Satellite Systems for Personal Applications, Wiley, 2010.
A. K. Maini & V. Agrawal, Satellite Technology, 2nd ed., Wiley, 2010.
J. C. Anselmo et al., Synthetic Aperture Radar for Earth Observation, ESA reference materials.