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Tsiolkovsky Rocket Equation Calculator (Δv)

The Tsiolkovsky rocket equation (1903) is the fundamental law of spaceflight: the ideal velocity change Δv = Isp · g₀ · ln(m₀ ⁄ m_f). Enter specific impulse (seconds) and the initial / final masses (any consistent mass unit); the tool returns Δv, exhaust velocity v_e, mass ratio and propellant fraction — the starting point for orbital mission planning, two-stage rocket sizing and Kerbal Space Program runs.

Ideal Δv

Exhaust velocity v_e = Isp · g₀
Mass ratio m₀ ⁄ m_f
Propellant burnt m₀ − m_f
Propellant fraction
Δv = Isp · g₀ · ln(m₀ ⁄ m_f), where g₀ = 9.80665 m/s². This is the "ideal" Δv — gravity losses, drag and thrust-vector losses are not included. Real Δv-to-LEO budgets are around 9.4 km/s.

Formula

Δv = Isp · g₀ · ln(m₀ ⁄ m_f) v_e = Isp · g₀ g₀ = 9.80665 m/s²

Frequently asked

Why is Isp measured in seconds — does that "second" mean anything in time?

Half history, half dimensional convenience. The more direct quantity is exhaust velocity v_e (m/s): Δv = v_e · ln(m₀⁄m_f). But early rocketeers preferred the ratio "pounds of thrust produced per pound of propellant burnt per second", which gives Isp = F ⁄ (ṁ · g₀); dimensionally that is (N) ⁄ ((kg/s) · (m/s²)) = s. Physically it says "1 unit of propellant mass can sustain 1 unit of weight-force as thrust for Isp seconds". So 250 s Isp ≈ 1 kg of propellant gives 1 kgf (9.80665 N) of thrust for 250 s. Because g₀ = 9.80665 m/s² is a fixed constant in both SI and US customary units, Isp in seconds is unit-free across systems — handy for cross-comparison. Isp (s) is the practical figure of merit; v_e (m/s) is what actually shows up in the physics.

Why do ion thrusters have such high Isp but SpaceX does not use them to reach orbit?

Because Δv and thrust are two different metrics. An ion thruster's Isp can reach 3000–5000 s (huge per-kg efficiency) and that is decisive for small probes — Deep Space 1, Dawn, BepiColombo all used ion drives. But ion-thruster thrust is tiny: the NSTAR engine produces only ~92 mN, about 9.4 grams-force. A launch vehicle must overcome its own weight (thrust-to-weight > 1), and chemical rockets deliver millions of newtons. To lift the 549 t Falcon 9 first stage you would need hundreds of thousands of ion engines — impossible. Ion drives also need months-to-years of continuous burn to accumulate Δv, which only works in vacuum away from strong gravity. The classic division: chemical rockets handle the high-thrust 0→LEO segment, electric propulsion handles the high-Isp LEO→deep-space segment.

Why is the Δv budget to LEO ~9.4 km/s when orbital speed is only ~7.8 km/s — where does the extra 1.6 km/s go?

The Tsiolkovsky equation gives the ideal Δv only (vacuum, point mass). Three real-launch losses must be added: (1) Gravity loss: during vertical climb thrust fights gravity rather than adding to velocity, typically 1.0–1.5 km/s. Modern reusable boosters (Falcon 9, Starship) with steep gravity-turn trajectories and high thrust-to-weight see less; older heavy launchers can lose up to 1.5–2 km/s. (2) Drag loss: traversing the lower atmosphere costs 0.1–0.3 km/s depending on shape and ascent profile. (3) Steering loss: angling the thrust vector wastes some Δv, around 0.1 km/s. Sum ≈ 1.5–2 km/s on top of the 7.8 km/s circular-orbit speed, giving a budget around 9.3–9.6 km/s. Saturn V / Apollo budgeted ~9.6 km/s; modern Falcon 9 is closer to 9.3 km/s. Note: re-entry and landing burns (e.g. Starship) are a separate Δv outside this 9.4 km/s envelope.

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