Satellite Launching: A Comprehensive Study Guide

By the end of this study guide, the student should be able to:

1. Explain the purpose and functions of satellite launch vehicles and classify the major families of launchers.
2. Apply the Tsiolkovsky rocket equation to determine the velocity capability and mass budget of a launch vehicle.
3. Describe the multistage ascent profile from lift-off to payload separation, including gravity, drag and steering losses.
4. Compute circular and escape orbital velocities and plan a Hohmann transfer from low Earth orbit (LEO) to geostationary orbit (GEO).
5. Compare launch sites worldwide and justify site selection using latitude, azimuth and safety criteria.
6. Identify the electrical engineering subsystems involved in a launch: telemetry, tracking & command (TT&C), flight computers, power conditioning and pyrotechnic circuits.
7. Analyse the mechanical and thermal launch environments (vibration, acoustic noise, acceleration) that a satellite must survive.
8. Evaluate modern trends including reusable launch vehicles, rideshare missions and small-satellite launchers.

A satellite launch is the engineered process of delivering a spacecraft from the surface of the Earth to its intended orbit (or beyond) using a launch vehicle — a rocket whose sole mission ends when the payload is released on the correct trajectory. For an electrical engineering student, launching is where several disciplines converge: control systems steer the vehicle, communication systems maintain the telemetry and command link, power electronics feed the avionics, and signal processing extracts navigation solutions from sensors.

Every orbital launch is fundamentally an energy management problem. To reach a 400 km circular orbit, a vehicle must lift the payload against gravity (potential energy), accelerate it to roughly 7.7 km/s (kinetic energy), and overcome atmospheric drag and steering losses along the way. Because a rocket must also carry its own propellant — and propellant is needed to lift propellant — the required vehicle is far larger than the payload: a typical ratio is that the payload is only 2–4% of the lift-off mass.

Why study launching in a satellite engineering course?

• Mission design starts at launch: the orbit achievable, the mass of the satellite, and even its structural design are dictated by the launcher.
• The satellite must survive the launch before it can do any work in orbit (see Section 9 on launch environment).
• Electrical systems are central: guidance, telemetry, sequencing, power and ordnance are all EE domains.
• Cost and scheduling: launch is typically 20–40% of total mission cost; launch delays cascade into spacecraft design freezes.

3.1 Classification by number of stages and reuse

3.2 Classification by lift capability

3.3 Classification by propellant type

3.4 Major launch systems (selected)

• Falcon 9 (SpaceX): two-stage LOX/RP-1, ~22.8 t to LEO; first stage routinely recovered; dominates commercial launch today.
• Ariane 6 (ESA/Arianespace): European heavy launcher replacing Ariane 5, with restartable Vinci upper stage (LOX/LH₂).
• Long March families (CNSA): workhorse of China's programme; LM-5 heavy lift supports space-station and deep-space missions.
• PSLV / LVM3 (ISRO): India's reliable medium/heavy vehicles; PSLV famous for high mission success and rideshare records.
• Soyuz-2 (Roscosmos): heritage R-7 design; launched from Baikonur, Plesetsk and Kourou.
• Starship / SLS: super-heavy class enabling crewed lunar return (Artemis) and interplanetary logistics.

4.1 Thrust

Thrust is produced by expelling propellant at high velocity (Newton's third law), plus the pressure thrust at the nozzle exit:

F = thrust (N) • ṁ = propellant mass flow (kg/s) • v_e = effective exhaust velocity (m/s) • p_e, p_a = nozzle exit and ambient pressure (Pa) • A_e = exit area (m²)

4.2 Specific impulse

Specific impulse, I_sp, measures propellant efficiency — thrust per unit weight-flow of propellant:

g₀ = 9.81 m/s² (standard gravity). Higher I_sp means more velocity per kilogram of propellant.

4.3 The Tsiolkovsky (ideal rocket) equation

The fundamental sizing equation of any launch vehicle relates achievable velocity change to exhaust velocity and the ratio of initial to final mass:

Δv = velocity change (m/s) • m₀ = initial (wet) mass • m_f = final (dry + payload) mass • R = mass ratio m₀/m_f

4.4 Why stages are needed

Once a stage's tanks are empty, their structure is dead mass. Discarding it (staging) improves the mass ratio of what remains. Optimum staging sizes each stage so that Δv contributions roughly balance structural and propellant fractions. Typical allocations:

Total ideal Δv ≈ 9.3 – 10 km/s is required from the vehicle to reach LEO, because roughly 1.5 – 2 km/s is lost to gravity and drag — the ideal (equation) Δv must cover both.

4.5 Launch mass budget (typical two-stage example)

5.1 The ascent profile

An orbital launch is not a vertical climb — it is a controlled pitch-over that trades altitude for horizontal velocity. A typical profile:

1. Vertical rise (0–15 s): clear the launch pad and towers before any steering.
2. Pitch-over / gravity turn: the vehicle slowly rotates toward the horizon so thrust (and later gravity) gradually builds horizontal velocity — an energy-efficient compromise between fighting gravity and gaining speed.
3. Max-Q: maximum dynamic pressure, q = ½ρv². Throttle-down protects the vehicle and payload; after this, engines throttle back up.
4. MECO & staging: Main Engine Cut-Off, stage separation, second-stage ignition.
5. Second-stage burn to orbit: performed above the sensible atmosphere, near-vacuum nozzle, possibly coast-and-restart to reach the precise target orbit.
6. Payload separation: the satellite is released on the target orbit; the stage usually performs a collision-avoidance / de-orbit manoeuvre.

5.2 Velocities that define the mission

μ_Earth = 3.986 × 10¹⁴ m³/s² • r = radius from Earth's centre (m). v_circ at 400 km ≈ 7.67 km/s; escape velocity ≈ 10.9 km/s there.

5.3 Launch azimuth and the benefit of Earth's rotation

Launching eastward exploits Earth's rotation, providing a free velocity bonus:

The target orbit inclination sets the launch azimuth (for a direct ascent, cos i = cos lat · sin Az, ignoring Earth's rotation). Sites far from the equator cannot reach low inclinations directly, which is why GTO missions favour Kourou.

5.4 From LEO to GEO: the Hohmann transfer

Geostationary satellites are usually injected into a Geostationary Transfer Orbit (GTO) — a highly elliptical orbit with perigee in LEO and apogee at GEO altitude (35 786 km). An apogee motor (or electric propulsion over months) then circularises at GEO.

r₁ = perigee radius (LEO) • r₂ = apogee radius (GEO = 42 164 km from Earth's centre)

Worked reference (400 km LEO → GEO): Δv₁ ≈ 2.45 km/s, Δv₂ (at apogee) ≈ 1.47 km/s, transfer time ≈ 5.26 hours. Try the calculator in Section 12.

A space launch is a months-long campaign, not just ten minutes of flight. The condensed sequence:

7.1 Selection criteria

• Latitude: low latitude maximises eastward rotational assist and access to low inclinations (hence Kourou's dominance for GTO).
• Azimuth corridors: downrange must be ocean or sparsely populated land (debris risk); this is why launches head east over ocean.
• Logistics & safety: propellant transport, buffer zones, lightning climatology (Cape Canaveral's afternoon storms drive many scrubs).
• Tracking & telemetry: downrange ground stations, TDRS relay satellites, tracking ships.
• Political/economic: national programmes vs commercial spaceports (e.g., spaceports proliferating in the USA, UK, and emerging sites worldwide).

Modern launch vehicles are "flying computer networks with engines attached". The main EE subsystems:

8.1 A typical telemetry chain (block-diagram view)

Launch is the most severe mechanical event in a satellite's life. Structural and electrical designers must qualify the spacecraft against:

Standard verification: workmanship (random vibe + thermal cycle), qualification (1.25× environments, protoflight or dedicated QM model), and acceptance testing of the flight unit. Environmental test standards: SMC-S-016, GSFC-STD-7000, ECSS-E-ST-32.

10.1 Falcon 9 rideshare and reusability economics

First-stage recovery and reflight (10+ flights per booster in routine service) reduced marginal launch cost dramatically, enabling rideshare programmes where many smallsats fly as secondary payloads. Electrical engineering highlight: landing uses grid fins (aerodynamic control), throttleable Merlin engines, and real-time GNSS/INS fusion — closed-loop control landing a 25-storey building on a drone ship.

10.2 Ariane 6 vs Ariane 5

Ariane 5 famously flew two-satellite GTO missions to GEO (dedicated dual launch). Ariane 6 adds a restartable Vinci cryogenic upper stage and a re-ignitable third stage option (A64 vs A62), restoring European autonomous access to space while cutting cost per kg.

10.3 Small launchers and air-launch

Electron (Rocket Lab) pioneered the dedicated smallsat class; LauncherOne (air-launched from a Boeing 747) showed mobile launch concepts. Air launch gives latitude/azimuth freedom and avoids weather scrubs at a fixed pad — at the cost of payload mass.

10.4 Emerging themes to watch

• Starship-class fully reusable super-heavy vehicles → order-of-magnitude cost changes and new payload classes.
• Mega-constellations (Starlink, OneWeb, Kuiper) driving weekly launch cadence — manufacturing-rate, not mission-rate, engineering.
• Green propulsion (methalox, hydrogen peroxide) and responsive/rapid launch for defence and disaster response.
• On-orbit servicing and space tugs changing the value chain after separation.

Use these to check the worked examples and explore "what-if" questions from your tutorials.

Attempt each question, then reveal the answer. Questions are typical of EEEN 567 continuous assessment.

Summary of Key Takeaways

1. Launch vehicles convert propellant into orbital energy; the payload is typically only 2–4% of lift-off mass because of the logarithmic rocket equation.
2. Ascent is a gravity turn to orbit velocity (~7.7 km/s at 400 km), with ~1.5–2 km/s lost to gravity and drag, hence ~9.3–10 km/s ideal Δv from the vehicle.
3. Site latitude sets the rotational assist and reachable inclinations; Kourou's 5°N latitude is a major economic advantage for GTO.
4. GEO missions use a Hohmann transfer via GTO (Δv₁ ≈ 2.45, Δv₂ ≈ 1.47 km/s from 400 km), or direct insertion by a cryogenic upper stage.
5. Launch is an electrical engineering showcase: GNC, flight computers, PCM telemetry, power conditioning, pyrotechnics and the safety-critical FTS.
6. The payload must be verified against vibration, acoustic, shock and thermal environments — environmental testing is a contractual deliverable, not an option.