Propulsion sizing
Propulsion Sizing Calculator
How much propellant does the mission need? The Tsiolkovsky rocket equation turns a delta-v budget into propellant mass, wet mass, and tank volume.
// Tsiolkovsky rocket equation. propellant mass from dry mass, exhaust velocity from Isp, first-order tank + burn-time sizing. cold-gas / monoprop / bipropellant / electric. engineering trade-study accuracy.
This calculator solves the Tsiolkovsky rocket equation — Δv = ve · ln(m_wet / m_dry) — turning a delta-v (Δv) budget into propellant mass, wet mass, mass ratio, tank volume, and burn time, with exhaust velocity from specific impulse (Isp): ve = Isp · g0. The propellant families span cold gas, hydrazine monopropellant, MMH/NTO bipropellant, and electric, with Isp and density values in the ranges of SMAD ch. 17 (Wertz et al., 2011) and Sutton's Rocket Propulsion Elements — which makes it equally at home as a CubeSat propulsion calculator or a smallsat deorbit-burn sizer.
How this model works & what it omits
Every orbital manoeuvre — station-keeping, plane change, deorbit, transfer — costs a velocity increment, the delta-v (Δv). Sizing a propulsion system means turning that Δv budget into a propellant mass and, from there, into tank volume and burn time. The governing relation is the Tsiolkovsky rocket equation: Δv = ve · ln(m_wet / m_dry), where ve is the exhaust velocity. Solved for propellant mass from the dry mass it becomes m_p = m_dry · (e^(Δv/ve) − 1).
Exhaust velocity follows from specific impulse: ve = Isp · g0, with standard gravity g0 = 9.80665 m/s². Because propellant mass grows exponentially with Δv/ve, a higher-Isp propellant collapses the propellant requirement dramatically — the central trade in propulsion selection. Cold gas (Isp ≈ 50–90 s) is simple but heavy; hydrazine monopropellant (≈ 200–235 s) and MMH/NTO bipropellant (≈ 300–340 s) are denser and more efficient; electric propulsion (≈ 1000–3000 s) is the most mass-efficient but delivers only milli-newtons of thrust, so burns last weeks.
Tank volume is sized from the stored propellant density and an ullage fraction — the gas headspace left for pressurant and thermal expansion. The internal tank volume is V = m_p / ρ / (1 − ullage). Cold gas is the only family stored as a gas, so its low density (~56 kg/m³ at high pressure) makes its tanks far bulkier than the dense liquids. Burn time at constant thrust is t = m_p · ve / F; for low-thrust electric systems the tool also suggests a count of discrete burns, each capped at one day of thrusting.
What this tool does not capture: gravity and steering losses, finite-burn losses for non-impulsive manoeuvres, residual and trapped propellant, pressurant mass, feed-system and tankage structural mass, blow-down pressure decay, and throttling. The result is a first-order trade-study sizing — detailed mission design needs a full propulsion feed-system model and an integrated mass budget.
// pick a scenario, then dial delta-v / dry mass / Isp.
Mission
// total delta-v the propulsion system must deliver.
Propulsion
// propellant family sets density; Isp sets exhaust velocity.
Tank & burn
// ullage = gas headspace; thrust sets burn time.
Propulsion sizing
// Tsiolkovsky: m_p = m_dry (e^(Δv/ve) − 1)
52.2 kg
Propellant mass
252.2 kg
Wet mass
1.261×
Mass ratio
20.7%
Propellant fraction
2.16 km/s
Exhaust velocity
57.4 L
Tank volume
31.3 h
Total burn time
// wet mass = dry + propellant
// shareable URL encodes every input. no backend.
// Tsiolkovsky — propellant mass vs Δv
// ai-generated breakdown of what these numbers mean — with diagrams.
Common questions
How do I calculate propellant mass from delta-v?
Solve the rocket equation for propellant from the dry mass: m_p = m_dry · (e^(Δv/ve) − 1). For the CubeSat cold-gas preset — a 6 kg spacecraft needing 50 m/s of reaction-control delta-v — that comes to about 0.45 kg of propellant, but roughly 9 litres of tank. Enter your own dry mass, delta-v, and propellant family and the tool solves the same chain through tank volume and burn time.
What specific impulse (Isp) should I assume?
By family: cold gas ≈ 50–90 s, hydrazine monopropellant ≈ 200–235 s, MMH/NTO bipropellant ≈ 300–340 s, electric propulsion ≈ 1000–3000 s. Higher Isp means exponentially less propellant for the same delta-v — the central trade in propulsion selection — but the families differ enormously in thrust, complexity, and power demand.
Why does propellant mass grow exponentially with delta-v?
Because every kilogram of propellant you carry for the end of a burn must itself be accelerated through the start of it — the compounding that puts Δv/ve in the rocket equation's exponent. It is also why high Isp pays so well: the electric station-keeping preset pushes a 450 kg spacecraft through 1500 m/s on about 40 kg of xenon (an 8% propellant fraction) where a chemical system would need several times more.
Why are cold-gas tanks so bulky?
Cold gas is the only family stored as a gas, at a storage density around 56 kg/m³ — versus roughly a tonne per cubic metre for the liquid propellants. Since tank volume is V = m_p / ρ / (1 − ullage), the same propellant mass costs an order of magnitude more volume — a real constraint inside a CubeSat structure, even when the propellant mass itself looks small.
What does burn time mean for electric propulsion?
At constant thrust, t = m_p · ve / F — and because electric thrusters deliver only milli-newtons, the total adds up to weeks or months of thrusting. The tool therefore also suggests a count of discrete burns, each capped at one day: the electric preset's 1500 m/s works out to roughly a hundred day-long burns.
References
- // Wertz, J. R., Everett, D. F., Puschell, J. J. (eds.) (2011). Space Mission Analysis and Design / Space Mission Engineering: The New SMAD, ch. 17 — Spacecraft Propulsion.
- // Sutton, G. P., Biblarz, O. (2017). Rocket Propulsion Elements, 9th ed. Wiley.
- // Goebel, D. M., Katz, I. (2008). Fundamentals of Electric Propulsion: Ion and Hall Thrusters. JPL / Wiley.
- // Tsiolkovsky, K. E. (1903). Exploration of Outer Space by Means of Rocket Devices.
