Launch Vehicle Selection Calculator
Payload capacity comparison across launch providers.
About this calculator
This calculator runs the Tsiolkovsky rocket equation in reverse to size a launch stage around a fixed payload rather than a fixed rocket. It first converts specific impulse (Isp) and standard gravity into exhaust velocity, then computes the mass ratio m0/mf = exp(Δv / ve) — the multiple by which the vehicle's mass must shrink from full propellant tanks to burnout for a given delta-v budget. Treating your payload mass as the final mass after burnout (a simplification that ignores staging, since a real multi-stage rocket splits this ratio across stages), it derives initial mass and therefore propellant mass directly. Structural mass then comes from your dry mass fraction — the fraction of a stage's non-payload mass that is tanks, engines, and airframe rather than propellant — rearranged algebraically from propellant mass.
Summing payload, propellant, and structure gives total launch mass, from which payload fraction (the efficiency metric engineers actually compare across vehicles) and launch cost (payload mass × cost per kg) fall out. The built-in delta-v references are real mission budgets: about 9.4 km/s to reach low Earth orbit from the ground (most of which fights gravity and drag, not just orbital velocity), 12 km/s to geostationary transfer orbit, and 15.9-16.5 km/s for direct lunar or Mars trajectories. Because this treats the whole vehicle as one idealized stage, it will overstate the propellant burden compared to a real multi-stage design, where each stage only needs to accelerate what's above it — use the payload fraction as a comparative, order-of-magnitude figure across scenarios rather than an exact substitute for a staged mission design tool.
Inputs
Results
Propellant mass (kg)
104,021
Total launch mass (kg)
120,579
Launch cost ($M)
13.6
How to Use This Calculator
- Enter payload mass (kg) and required delta-V (km/s) for your target orbit.
- Set propellant specific impulse (sec) and dry mass fraction (%) for the launch vehicle stage.
- Enter cost per kg to orbit ($) for the candidate launch vehicle.
- Review mass ratio, propellant mass, structural mass, total launch mass, and payload fraction.
- Compare payload fraction across launch vehicle options to identify the most mass-efficient choice.
How the result changes with Specific impulse (sec)
| Specific impulse (sec) | Propellant mass (kg) | Total launch mass (kg) | Launch cost ($M) |
|---|---|---|---|
| 156 | 2,325,602 | 2,589,002 | 13.6 |
| 233 | 300,917 | 339,352 | 13.6 |
| 467 | 33,938 | 42,709 | 13.6 |
| 778 | 12,141 | 18,490 | 13.6 |
What each input means
- Payload mass (kg)
- Mass of the payload to be delivered to the target orbit.
- Required delta-v (km/s)
- Total velocity change needed. LEO ~9.4, GTO ~12, Moon ~15.9, Mars ~16.5 km/s from surface.
- Specific impulse (sec)
- Engine Isp. RP-1/LOX ~311s, LH2/LOX ~450s, ion ~3000s, solid ~260s.
- Dry mass fraction (%)
- Structural mass as percentage of (structure + propellant). Modern rockets: 5-12%.
- Cost per kg to orbit ($)
- Launch cost per kg of payload. Falcon 9: ~$2,720/kg, Atlas V: ~$13,200/kg, Electron: ~$26,000/kg.
What each result means
- Mass ratio (m₀/m_f)
- Tsiolkovsky mass ratio: initial mass / final mass.
- Propellant mass (kg)
- Required propellant mass from the rocket equation.
- Structural mass (kg)
- Rocket dry structure mass (tanks, engines, fairings).
- Total launch mass (kg)
- Payload + propellant + structure.
- Payload fraction (%)
- Payload as percentage of total launch mass.
- Propellant fraction (%)
- Propellant as percentage of total launch mass.
- Exhaust velocity (km/s)
- Effective exhaust velocity = Isp x g₀.
- Launch cost ($M)
- Estimated launch cost in millions of dollars.
How this is calculated
Worked example, using the default values
- Identify Input Parameters4 parametersPayload mass (kg) = 5000, Required delta-v (km/s) = 9.4, Specific impulse (sec) = 311, Dry mass fraction (%) = 10 = 5 input(s) provided
- Calculate Propellant massPropellant mass = m_initial - m_final104021 = 104021
- Calculate Total launch massTotal launch mass = payloadMassKg + propellantMassKg + structuralMassKg120579 = 120579
- Calculate Launch costLaunch cost = (payloadMassKg * costPerKgUSD) / 1_000_00013.6 = 13.6
- Calculate Mass ratio21.804 = 21.804
- Calculate Structural massStructural mass = (propellantMassKg / (1 - dryFraction)) * dryFraction11558 = 11558
Engine last updated . Checked against 2 independently-derived tests — how we verify calculators. Built by Paul Gunder, a software engineer, not a licensed financial, medical, or legal professional.
Frequently Asked Questions
Why does the calculator treat payload mass as the rocket equation's "final mass" instead of adding structural mass to it first?
The code explicitly simplifies a real multi-stage rocket into a single idealized stage: m_final is set equal to payloadMassKg, then mass ratio and propellant mass are derived from that. Structural mass is worked out afterward from the resulting propellant mass and your dry mass fraction, rather than being included in m_final up front — which is why the explainer notes this will overstate propellant needs compared to a properly staged design.
Why does raising specific impulse (Isp) shrink required propellant mass so much?
Isp feeds directly into exhaust velocity (Isp × 9.80665 m/s²), and the mass ratio is exp(Δv / exhaust velocity) — since delta-v sits in the exponent's denominator, a higher exhaust velocity shrinks that exponent and therefore the mass ratio non-linearly. Going from a kerosene engine's ~311s to an ion engine's ~3000s can turn a mass ratio in the double digits into one barely above 1, which is why ion propulsion needs so little propellant for the same delta-v.
How is structural mass calculated from the dry mass fraction?
It's rearranged algebraically from propellant mass: structuralMassKg = (propellantMassKg / (1 - dryFraction)) × dryFraction, where dryFraction is your dry mass fraction percentage divided by 100. This reflects that dry mass fraction is defined as structural mass over the sum of structural-plus-propellant mass, not over total launch mass including payload.
Why would a real launch vehicle need less propellant than this calculator predicts for the same payload and delta-v?
Because the model treats the entire delta-v budget as one stage's job, when real rockets split it across multiple stages that each only have to accelerate the mass above them. A staged design lets each stage drop its own dead structural weight before the next stage fires, which is intrinsically more propellant-efficient than one giant single-stage burn — so treat this calculator's propellant and mass figures as a conservative, order-of-magnitude comparison tool, not a final design number.
Related Calculators
The questions that sit next to this one — chosen by subject, including calculators filed under a different category.
Rocket Propulsion Calculator
Calculate rocket delta-V, exhaust velocity, and burn time using the Tsiolkovsky rocket equation.
Aerospace EngineeringSpacecraft Mass Budget Calculator
Estimate spacecraft dry mass, propellant mass, and payload fraction from subsystem mass percentages.
Space TechnologySpace Mission Cost Estimator Calculator
Total mission cost from spacecraft mass and destination.
Amateur RocketryRocket Motor Selection Calculator
Determine the right NAR motor class for your rocket based on mass, target altitude, and drag estimates using energy-based delta-V and impulse calculations.
Space TechnologyCubeSat Budget Calculator
Build cost, launch cost, and operations for CubeSat missions.
More in Science & Physics.