Learning Objectives
By the end of this study guide, the student should be able to:
- Identify the major subsystems of a communications satellite and explain the function of each.
- Distinguish between the bus (platform) subsystems and the payload, and explain how they interact.
- Describe the operating principles and hardware of the Attitude Determination and Control Subsystem (ADCS), power, thermal, propulsion, TT&C, and command & data handling subsystems.
- Perform preliminary sizing calculations for the power subsystem (solar array and battery) and analyse a simplified communication link budget.
- Compare attitude control actuators and thermal control techniques, selecting appropriate options for given mission constraints.
- Relate subsystem design trades to overall satellite mass, cost, reliability and mission lifetime.
Overview: The Satellite as a System
A communications satellite is conventionally divided into two functional blocks:
- The bus (platform): everything needed to keep the satellite alive and correctly oriented — structure, ADCS, power, thermal, propulsion, TT&C, and command & data handling. The bus typically consumes 30–40% of the spacecraft's mass and much of its cost.
- The payload: the mission equipment that earns revenue or produces science — for a comsat, the repeater (receiver, transponder channels, high-power amplifiers) and the antenna system. The payload determines the satellite's value; the bus merely enables it.
Fig. 1 — Functional decomposition of a communications satellite: bus, payload and ground segment.
Design driver: every subsystem is sized by a “worst-case” scenario — eclipse for power, solstice for thermal, station-keeping propellant for lifetime, and end-of-life (EOL) conditions (degraded arrays, aged batteries) for the link budget. Designs are verified against these corner cases, not nominal conditions.
SUBSYSTEM 1 — STRUCTURE & MECHANISMS
Structural Subsystem
The structure provides the mechanical frame that carries all other subsystems through launch loads and maintains their alignment in orbit. It must survive the most violent part of the mission — launch — where quasi-static loads of 6–10 g and random vibration of tens of grms are typical.
Key Functions
- Withstand quasi-static acceleration, random vibration, acoustic noise and shock during launch and separation.
- Maintain dimensional stability for antenna pointing (structure deflections translate directly into pointing error).
- Provide thermal conduction paths and grounding for electronics.
- Interfaces: launch vehicle adapter (separation ring), apogee motor (if used), solar array and antenna hinges.
Materials and Construction
- Primary structure: aluminium honeycomb sandwich panels or aluminium/composite monocoque with longerons — high stiffness-to-mass ratio.
- Materials: aluminium alloys (classic), carbon-fibre-reinforced polymer (CFRP) for high stiffness and near-zero thermal expansion (critical for antenna reflectors and optical benches), titanium fittings for high-stress joints.
- Mass fraction: structure is typically 15–25% of dry mass; every kg of structure is a kg of lost payload or propellant.
Mechanisms and Deployment
- Deployment devices: solar array hinges/dampers, antenna booms, reflector deployment — usually single-shot, spring or motor-driven, with hold-down/release mechanisms (HDRMs) such as pyrotechnic or non-pyrotechnic (paraffin, shape-memory alloy) release nuts.
- Momentum management: solar array drive mechanisms (SADMs) rotate arrays to track the Sun; antenna pointing mechanisms (APMs) steer beams.
- Reliability is paramount: deployment failures are a leading cause of mission loss, so mechanisms are designed with redundancy and extensive ground testing.
SUBSYSTEM 2 — ATTITUDE DETERMINATION & CONTROL (ADCS)
Attitude Determination and Control Subsystem
The ADCS keeps the satellite pointed in the right direction — antennas toward Earth (or coverage zone), solar arrays toward the Sun — and manages the angular momentum budget of the spacecraft. Typical pointing requirements range from ±0.05° (comsat antennas) to arc-minutes (imaging).
1. Attitude Determination (sensing)
| Sensor | Principle | Accuracy | Use |
|---|---|---|---|
| Earth horizon sensor (IR) | Detects Earth's IR horizon (15 μm CO₂ band) | ~0.1° | LEO/GEO roll & pitch |
| Sun sensor | Photodiode array measures Sun angle | 0.01–1° | Safe mode, array pointing |
| Star tracker | Images star field, matches to star catalogue | 1–10 arcsec | High-accuracy 3-axis |
| Gyroscope (rate) | Measures angular rate (mechanical / ring-laser / fibre-optic) | rate, bias drifts | Control feedback; dead-reckoning between updates |
| Magnetometer | Measures local magnetic field vector | ~0.5° | LEO only (needs field) |
| GPS receiver (LEO) | Carrier-phase attitude (multi-antenna) | ~0.1° | Modern LEO buses |
2. Attitude Control (actuation)
| Actuator | Principle | Pros | Cons |
|---|---|---|---|
| Reaction / momentum wheel | Motor-driven flywheel; torque by changing wheel speed | Precise, no propellant | Rate saturation; needs momentum dumping |
| Control moment gyro (CMG) | Gimballed spinning rotor redirects momentum vector | High torque authority | Complex, heavy, costlier |
| Magnetic torquer (torque rod) | Current loop interacts with Earth's B-field | No propellant | Works only in LEO; torque ⊥ B |
| Thrusters (cold gas / hydrazine) | Expellant mass produces reaction torque | Any axis, momentum dumping | Consumes propellant; plume contamination |
| Gravity-gradient boom | Long boom aligns with gravity gradient | Simple, passive | Weak; only 2-axis; LEO |
Momentum dumping: τext = dL/dt ⇒ dumped momentum ΔL = ∫τmag dt = ∫ (m × B) dt
Environmental torques (gravity-gradient, solar radiation, aerodynamic, magnetic) build wheel momentum; periodic unloading against magnetic torquers (LEO) or thrusters (GEO) is required.
3. Typical Control Modes
- Acquisition: stabilise after separation (often spin or Sun-pointing safe mode).
- Normal (nominal): nadir/Earth-pointing with array tracking; control loops at 1–10 Hz with sensor fusion (e.g., star tracker + gyros via Kalman filter).
- Station-keeping / manoeuvre: momentum desaturation before and after burns.
- Safe mode: Sun-pointing, minimal power, awaiting ground recovery.
Worked Example 1 — Gravity-gradient torque scale: For a comsat with inertia difference (Iy − Ix) ≈ 1,000 kg·m² at GEO, the maximum gravity-gradient torque is Tgg = (3μ/2R³)(Iy − Ix)sin2θ ≈ 7×10-6 N·m — tiny, but over a 15-year mission it accumulates wheel momentum that must be dumped weekly against magnetic torquers (LEO) or thrusters (GEO).
SUBSYSTEM 3 — ELECTRICAL POWER SYSTEM (EPS)
Power Subsystem
The EPS generates, stores, conditions and distributes electrical power. For GEO comsats the design case is eclipse (Earth's shadow, up to 72 min around equinox); for LEO it is the rapid day/night cycling (e.g., 16 eclipses/day at ~35 min each in a 400 km, 51.6° orbit).
1. Power Generation — Solar Arrays
- Photovoltaic cells: triple-junction GaInP/GaAs/Ge cells reach 30±% efficiency (BOL, 28°C); silicon cells ~18–20% are cheaper for CubeSats.
- Array types: body-mounted (small sats) < deployable rigid panels < flexible blanket arrays (ISS, large GEO).
- Degradation: arrays lose ~2–3% of power per year from ultraviolet radiation, particle bombardment and micrometeoroids; designs are sized for EOL power, typically 70–80% of BOL.
- Sun-pointing via the SADM maximises flux; off-pointing of even 10° costs ~1.5% power.
2. Energy Storage — Batteries
- Nickel–hydrogen (NiH₂): the GEO comsat workhorse; 40,000+ cycles capability, 60–80 Wh/kg, pressure vessel doubles as housing.
- Lithium-ion: higher energy density (100–160 Wh/kg), now standard on LEO/CubeSats; managed with cell balancing and strict depth-of-discharge (DOD) limits.
- Depth of discharge vs. lifetime: NiH₂ survives ~40% DOD for 15 years GEO duty; Li-ion at 30% DOD achieves thousands of LEO cycles. Battery sizing is a direct lifetime trade.
3. Power Conditioning and Distribution
- Unregulated bus (28 V) for small sats; regulated 50/70/100 V buses on large comsats for lower harness current (P = VI).
- Peak power tracking (PPT) or direct-energy transfer (DET) from array; shunt dissipators dump excess Sun-side power.
- Pyro drives, heaters, and load-shedding: the EPS must survive and prioritise loads during anomalies — safe mode assumes payload OFF.
Battery capacity: C = Pload × Tecl / (DOD × Vbus × ηdis)
C (Ah), Pload (W), Tecl (h), Vbus (V), ηdis ≈ 0.9 discharge efficiency
Solar array sizing (EOL): Psa = Pecl × Tecl / (Tsun × ηreg) + Pday
Energy balance: the array must recharge the eclipse energy plus run daylight loads; then divide by EOL efficiency (≈0.75) and cosine off-pointing factor.
Worked Example 2 — GEO eclipse battery: Payload+bus load 5 kW, eclipse 1.2 h, bus 50 V, DOD 40%, η = 0.9.
C = 5000 × 1.2/(0.40 × 50 × 0.9) = 333 Ah → two 175 Ah NiH₂ strings for redundancy.
C = 5000 × 1.2/(0.40 × 50 × 0.9) = 333 Ah → two 175 Ah NiH₂ strings for redundancy.
SUBSYSTEM 4 — THERMAL CONTROL
Thermal Control Subsystem (TCS)
In vacuum there is no convection — heat leaves only by radiation. The TCS keeps every unit within its allowable temperature range (–10 to +50 °C for most electronics; much tighter for batteries, propellant lines, and optical payloads).
1. The Governing Equation
ε σ A T4 = α S Aproj + Qinternal + Qalbedo + QIR
Steady-state radiative balance: emitted IR = absorbed sunlight + internal dissipation + Earth albedo & IR (σ = 5.67×10-8 W/m²K4). At 1 AU, S ≈ 1361 W/m².
2. Passive Techniques (no power, high reliability)
- Surface finishes: white paint / optical solar reflectors (OSRs) give low solar absorptivity α and high IR emissivity ε (α/ε ≈ 0.1–0.3) for radiators; black anodising raises absorption for cold units.
- Multi-layer insulation (MLI): 10–20 layers of aluminised Kapton/Mylar with spacer mesh; effective conductivity ≈ 0.03–0.06 W/m·K — blankets protect equipment bays and propellant lines.
- Radiators: dedicated panels sized by Q/εσT&sup4;; louvres (variable-emittance blinds) modulate rejection without power.
- Heat pipes / constant-conductance arms: two-phase transport moving heat from dissipating boxes to radiator edges; loop heat pipes for high-power payloads.
- Thermal doublers, fillers, straps: spread and conduct heat across jointed structures.
3. Active Techniques (powered, used where passive is insufficient)
- Electric heaters: survival heaters on batteries, thrusters, propellant lines (often thermostatic, drawing from the EPS during eclipse/safe mode).
- Mechanical pumped loops / fluid loops: ammonia loop (ISS) for high heat-flux transport.
- Thermo-electric coolers (TECs): for detectors needing below-ambient temperatures.
- Louvres (motorised) and thermal switches / cryocoolers for IR instruments.
Worked Example 3 — Radiator sizing: A TWTA dissipates 300 W that must be rejected at 300 K with painted radiator (ε = 0.85, two-sided):
A = Q/(2εσT&sup4;) = 300/(2 × 0.85 × 5.67×10-8 × 3004) = 300/780 ≈ 0.38 m² (view factors and degradation add margin → ~0.5 m² flight).
A = Q/(2εσT&sup4;) = 300/(2 × 0.85 × 5.67×10-8 × 3004) = 300/780 ≈ 0.38 m² (view factors and degradation add margin → ~0.5 m² flight).
SUBSYSTEM 5 — PROPULSION
Propulsion Subsystem
Propulsion provides the velocity increments (ΔV) for orbit raising (e.g., GTO → GEO), station-keeping, momentum dumping and end-of-life disposal. Total mission ΔV budget is a primary sizing parameter.
1. The Rocket Equation
ΔV = g0 Isp ln(m0/mf)
Tsiolkovsky: g0 = 9.81 m/s², Isp (s) = specific impulse, m0/mf = initial/final mass ratio.
2. Thruster Types
| Type | Propellant / principle | Isp (s) | Thrust | Typical use |
|---|---|---|---|---|
| Cold gas | Pressurised N₂/He expansion | 50–75 | mN–N | Micro-sats, attitude |
| Monopropellant hydrazine | Catalytic decomposition on Ir/Al₂O₃ bed | 220–235 | 0.5–25 N | Station-keeping, attitude (workhorse of comsats) |
| Bipropellant MMH/N₂O₄ | Hypergolic ignition | 300–340 | 10 N–500 N | Orbit raising, large manoeuvres |
| Solid (AKM) | Composite solid grain, single burn | 280–300 | kN | Apogee kick motor (legacy GTO→GEO) |
| Electric (Hall / ion) | Xenon ionised & accelerated electrostatically | 1,500–3,500 | mN–0.5 N | Orbit raising & station-keeping (huge Isp → 50%+ propellant saving) |
~50 m/s/yr
GEO N-S + E-W station-keeping (ΔV)
~11 m/s/yr
LEO 400 km drag makeup (solar-cycle dependent)
1,800 m/s
GTO → GEO apogee raise (chemical)
10–15 yr
Comsat lifetime limited by propellant
Electric orbit raising (EOR): replacing chemical apogee burns with Hall thrusters cuts propellant by half but raises GTO→GEO transfer time from hours to ~5 months — a classic systems trade between mass and time-to-revenue.
SUBSYSTEM 6 — TELEMETRY, TRACKING & COMMAND (TT&C)
TT&C Subsystem
TT&C is the satellite's lifeline to the ground: it reports health (telemetry), carries commands, and provides ranging for orbit determination. It must be simple, robust, and able to command the satellite even when everything else has failed.
Functions
- Telemetry: housekeeping data — voltages, currents, temperatures, pressures, status bits — collected by the CDH and downlinked, typically at 1.5–4 kbps (low rate, high reliability).
- Command: uplinked instructions decoded and distributed by the CDH; protected by authentication and, for critical commands, multiple confirmations / interlocks (e.g., arming before firing pyros).
- Ranging / orbit determination: tone or code ranging through the TT&C transponder; coherent transponders (turn-around ratio 240/221, 749/719 typical) allow two-way Doppler and range for precise orbit tracking.
Typical Design
- Transponder: transparent turn-around (bent-pipe), S-band uplink ~2 GHz / downlink ~2.2 GHz (or C-band).
- Antennas: low-gain omnidirectional antennas guarantee a link in any attitude (safe mode); the main payload antennas may serve double duty for high-rate telemetry.
- Redundancy: dual transponders with cross-strapping; survival heaters and processors powered from a dedicated “essential bus.”
- Independent of the payload: TT&C stays alive with the payload switched off.
Two-way Doppler: Δf = (2vr/c) f0 ⇒ vr = c Δf / (2 f0)
Radial velocity from measured Doppler shift on the coherent carrier; integrated with ranging arcs for orbit determination.
SUBSYSTEM 7 — COMMUNICATION PAYLOAD (TRANSPONDER)
Communication Payload
The payload receives uplink signals, processes them (frequency conversion, amplification, channelisation) and retransmits them on the downlink. Architectures fall into two families:
- Transparent (bent-pipe) transponder: RF in → low-noise amplification → down-conversion → channel filtering → high-power amplification → RF out. Signal content is untouched; only the carrier is shifted. Simple, low delay, but uplink noise and interference pass straight through.
- Regenerative (processing) repeater: demodulation and remodulation on board — the satellite acts as a base-station in the sky; improves link budget (error correction before downlink), enables beam routing and interference isolation, at the cost of complexity.
Payload Chain (per channel)
| Stage | Function | Key device / parameter |
|---|---|---|
| Receive antenna | Collect uplink; spatial discrimination | Reflector + feed; G/T = Grx − 10log Tsys |
| LNA | Amplify weak uplink with minimal noise | HEMT/TLA; Tsys = Tant + Trx; NF < 1 dB |
| Down-converter | Shift carrier to downlink band | Mixer + LO; e.g., C-band 6/4 GHz, Ku 14/12 GHz, Ka 30/20 GHz |
| Channel filter (IMUX/OMUX) | Separate FDM channels; define bandwidth | 36 MHz or 72 MHz typical per transponder |
| HPA | Boost to downlink power | TWTA (60–250 W, high efficiency, linearised) or SSPA (better linearity, lower power) |
| Transmit antenna | Focus EIRP onto coverage zone | Shaped reflector; gain G = η(πD/λ)² |
EIRP = Ptx + Gtx (dBW); C/N₀ = EIRP + (G/T)rx − Lfs − Latm + 228.6 − 10log10 B
Link budget at the receive input: 228.6 = 10log(k), k = 1.38×10-23 J/K. Requirement: C/N₀ ≥ threshold + margin (≈ 1–2 dB, plus rain fade at Ku/Ka).
Worked Example 4 — Downlink C/N₀: GEO satellite EIRP 45 dBW, earth-station G/T = 30 dB/K, slant range 37,500 km, 12 GHz, clear sky.
Lfs = 20log(4πd/λ) = 20log(4π × 3.75×107/0.025) = 205.5 dB → C/N₀ = 45 + 30 − 205.5 + 228.6 = 98.1 dB·Hz — far above any modem threshold (≈ 85–90 dB·Hz); the margin (10+ dB) is what absorbs rain fade.
Lfs = 20log(4πd/λ) = 20log(4π × 3.75×107/0.025) = 205.5 dB → C/N₀ = 45 + 30 − 205.5 + 228.6 = 98.1 dB·Hz — far above any modem threshold (≈ 85–90 dB·Hz); the margin (10+ dB) is what absorbs rain fade.
SUBSYSTEM 8 — COMMAND & DATA HANDLING (CDH)
Command and Data Handling (On-Board Computer)
The CDH (or OBDH) is the spacecraft's brain: a fault-tolerant processor architecture that collects telemetry, executes commands, manages the attitude control loops, handles timekeeping, and autonomously responds to anomalies.
Key Functions
- Housekeeping acquisition: polling of all subsystem status via data buses (MIL-STD-1553, CAN, SpaceWire).
- Command validation and distribution: decode, authenticate, check pre-conditions (e.g., Sun angle, thermal limits) before actuation; time-tagged command queues for scheduled events (station-keeping, eclipse entry).
- Control law execution: runs the ADCS control loops; sensor fusion (e.g., extended Kalman filter combining star trackers and gyros).
- Autonomy & fault protection: FDIR (Fault Detection, Isolation and Recovery) — monitors limits, switches to redundant units, reconfigures, and if needed enters safe mode without ground intervention (critical when out of contact).
- Data storage / playback: solid-state recorders (SSRs) buffering payload and telemetry data for later downlink (imaging missions, store-and-forward).
- Timekeeping: on-board clock (often GPS-disciplined in LEO) for time-tagging and coherent operations.
Design Features
- Radiation tolerance: total-dose hardening and latch-up protection; error-detection-and-correction (EDAC) on memories, watchdog timers.
- Redundancy: dual cold- or warm-spare computers with cross-strapping; voting in triple-modular-redundant (TMR) designs for highest reliability.
- Deliberate conservatism: flight-proven processors (e.g., RAD750 family) rather than fastest COTS, trading performance for reliability.
SUBSYSTEM 9 — ANTENNA SUBSYSTEM
Antenna Subsystem
Antennas provide the spatial interface between the satellite and Earth. Payload antennas must shape the beam to match the coverage zone exactly — wasting power outside the zone is directly lost revenue.
Common Antenna Types
| Type | Gain / beamwidth | Use |
|---|---|---|
| Omni / helix / turnstile | 0–6 dBi, hemispherical | TT&C, beacon (any-attitude coverage) |
| Horn | 10–20 dBi, 20–40° | Feeds for reflectors, global-beam comsats |
| Parabolic reflector (shaped) | 30–45 dBi, 0.5–2° | Contoured beams: one per coverage country/region |
| Reflector + feed array | Multiple shaped beams from one aperture | High-throughput satellites (HTS): 50–200+ beams |
| Phased array | Electronic beam steering, no moving parts | LEO mobile (Iridium), multi-beam, agile shaping |
G = η(πD/λ)²; beamwidth ≈ 70λ/D degrees
Reflector sizing: a 2 m Ku-band (14 GHz) dish → G ≈ 44 dBi, beamwidth ≈ 1.5° — small enough to reuse the same frequency in an adjacent country (frequency reuse by spatial separation).
Coverage shaping is the business model: a “contoured beam” reflector concentrates EIRP on Kenya/East Africa and suppresses spillover into the ocean; multi-beam HTS payloads reuse each band dozens of times — capacity scales roughly with the number of beams.
System Design Budgets & Interactive Calculators
Satellite design is bookkeeping: every subsystem contributes rows to a budget, and the spacecraft must close all of them simultaneously. The three master budgets are mass, power and ΔV; the payload adds the link budget.
1. Power Subsystem Sizing Calculator
Enter mission parameters to size the eclipse battery and estimate the solar array EOL power requirement.
2. Simplified Link Budget Calculator
Compute clear-sky C/N₀ for a GEO downlink and compare with a typical modem threshold.
Summary of Key Points
- A satellite = bus (structure, ADCS, power, thermal, propulsion, TT&C, CDH) + payload (antennas + repeater). The bus enables; the payload earns.
- Structure survives launch (6–10 g, vibration) and preserves alignment; CFRP gives stiffness with low thermal distortion. Deployment mechanisms are single-point-failure-critical.
- ADCS: sense with star trackers/horizon sensors/gyros; act with reaction wheels (precision), magnetic torquers (LEO, no propellant), thrusters (any axis). Environmental torques accumulate wheel momentum → periodic dumping.
- EPS: sized on the eclipse case; NiH₂/Li-ion batteries at 30–40% DOD; arrays derated for ~2–3%/yr degradation — design to EOL power.
- TCS: radiation-only heat rejection; passive MLI/radiators/OSRs first, heaters and louvres when passive is insufficient; α/ε ratio is the master surface parameter.
- Propulsion: Tsiolkovsky governs; hydrazine for station-keeping, bipropellant for orbit raising, electric propulsion for Isp > 1500 s with time penalty. Propellant usually ends the mission.
- TT&C: always-on lifeline; omnidirectional antennas + redundant transponders; ranging/Doppler for orbit determination.
- Payload: bent-pipe vs. regenerative; chain = LNA → converter → IMUX → HPA (TWTA/SSPA) → OMUX → shaped-beam antenna; link closes when C/N₀ exceeds threshold plus margin.
- CDH: fault-tolerant computer running control loops, command validation, FDIR autonomy and SSR storage; radiation-hardened, EDAC-protected, redundant.
- Master budgets (mass, power, ΔV, link) must all close simultaneously at worst-case/EOL conditions — with margin.
Review Questions
Q1. Distinguish the bus from the payload of a communications satellite, listing the subsystems that belong to each.
Show answer
The bus keeps the spacecraft alive and pointed: structure & mechanisms, ADCS, EPS, thermal, propulsion, TT&C, CDH. The payload performs the mission: receive/transmit antennas and the repeater (LNA, converters, channel filters, HPAs). Ground segment is separate: TT&C stations, control centre, user earth stations.Q2. A reaction wheel saturates during a mission. Explain why this happens and describe two ways to recover the attitude-control authority.
Show answer
Constant environmental torques (gravity-gradient, solar radiation) integrate wheel momentum to its speed limit. Recovery = momentum dumping: (1) magnetic torquers against Earth's field (LEO, propellant-free); (2) thruster firings (GEO). Both apply an external torque to transfer momentum off the spacecraft.Q3. Size the battery for a LEO satellite: 400 W load, eclipse 35 min, bus 28 V, Li-ion DOD 30%, discharge efficiency 90%.
Show answer
C = 400 × (35/60) / (0.30 × 28 × 0.9) = 233/7.56 = 30.8 Ah → specify ~35–40 Ah with margin; 16 eclipses/day implies ~6000 cycles/yr — hence low DOD for cycle life.Q4. Why are GEO comsats designed to an “EOL power” rather than “BOL power,” and what causes the difference?
Show answer
Arrays degrade ~2–3%/yr from UV, particle radiation (both Van Allen transit on the way up and the on-station environment), and micrometeoroid pitting; over 15 years ≈ 25% loss. Batteries also fade. Since the mission must perform at year 15, the array is sized so that EOL power still meets the load, meaning BOL power is deliberately oversized.Q5. Explain the physical basis of the α/ε ratio in thermal design and give typical values for a radiator and for a solar array.
Show answer
α (solar absorptivity) controls heat absorbed from the Sun; ε (IR emissivity) controls heat radiated away. Radiators want low α/ε (≈ 0.1–0.3, white paint/OSR) to stay cold; solar cells need high α (≈ 0.8) to absorb light and run warm (α/ε ≈ 0.8–0.9) — which is why arrays need radiators behind them for high-power sats.Q6. Using Tsiolkovsky, how much propellant fraction is needed for a ΔV of 1,800 m/s with a hydrazine thruster (Isp = 225 s)? With a Hall thruster (Isp = 2,000 s)?
Show answer
m0/mf = exp(ΔV/(g₀Isp)): hydrazine → exp(1800/2207) = 2.26 → propellant fraction 56%; Hall → exp(1800/19620) = 1.096 → only 8.8%. This is the mass case for electric orbit raising — at the cost of months of transfer time.Q7. Why must the TT&C subsystem remain operational with the payload switched off, and how is this achieved in practice?
Show answer
TT&C is the only means to diagnose and recover the satellite; if it depended on the payload, a payload failure would be fatal. Achieved by omni antennas (link in any attitude), a separate essential power bus, redundant cross-strapped transponders, and survival heaters — safe mode assumes payload off, Sun-pointing, minimal loads.Q8. Compare TWTA and SSPA for the high-power amplifier of a transponder channel.
Show answer
TWTA: high efficiency (≈ 60–70% at saturation), high power (100–250 W), mature; but needs high-voltage supplies, is bulky and less linear (needs linearisers/operating back-off). SSPA (GaN): excellent linearity, instant on, lower voltage, rugged; lower per-device power and efficiency at very high power — chosen for multi-carrier linearity, TWTAs for raw downlink power economics.Q9. A contoured-beam antenna covers East Africa; the operator wants to add capacity. Explain two payload-level options and their trade-offs.
Show answer
(1) Add transponder channels (wider bandwidth) — limited by spectrum filings and power; (2) move to a multi-beam/HTS payload — many narrow beams reuse the same frequencies N times and concentrate EIRP, multiplying capacity per Hz; costs: more HPAs/feeds or a phased array, higher DC power (EPS grows), and complex beam-forming on board.Q10. State the role of FDIR in the CDH and give a typical escalation sequence.
Show answer
FDIR = Fault Detection, Isolation, Recovery: the autonomy layer that keeps the satellite safe without ground contact. Typical escalation: (1) detect out-of-limit telemetry / watchdog; (2) retry or switch to redundant unit (isolation); (3) reconfigure loads / safing (payload off, arrays Sun-pointed); (4) persist → safe mode with beacon + omni TT&C, awaiting ground commanding; critical events (imminent collision, power crisis) may trigger immediate safe mode.References and Further Reading
- Richharia, M. (2014). Satellite Communication Systems. McGraw-Hill — payload and link budgets.
- Fortescue, P., Swinerd, G., & Stark, J. (2011). Spacecraft Systems Engineering (4th ed.). Wiley — the standard bus-subsystem reference.
- Larson, W. J., & Wertz, J. R. (Eds.) (1999). Space Mission Analysis and Design (3rd ed.). Microcosm Press — subsystem sizing trades.
- Pratt, T., Bostian, C., & Allnutt, J. (2003). Satellite Communications (2nd ed.). Wiley — transponders and link budgets.
- Wertz, J. R. (1978/2001). Spacecraft Attitude Determination and Control. Springer/D. Reidel.
- Gilmore, D. (Ed.) (2002). Spacecraft Thermal Control Handbook. Aerospace Press.
- EEEN 567 Course Lecture Notes, Department of Electrical and Electronic Engineering, Egerton University.