# Introducing the Launcher Rocket Calculator

## An open-source orbital launch vehicle payload calculator by Launcher

8 min readJun 23, 2020

When a company announces a new orbital rocket design, it is nearly impossible for anyone, including media or investors, to verify the payload mass claim, compare it to other rockets, or judge how realistic the payload mass claim is.

The Launcher Rocket Calculator was created to help solve this problem and increase transparency into the assumptions behind historical, existing, and future rockets. It can also be used for analysis or due diligence on proposed new rocket designs.

# The calculator currently features estimated presets based on public sources for:

• Launcher Rocket-1, Light, XL and Nano
• NASA Saturn V
• Yuzhnoye Zenit
• ULA Delta-IV
• ULA Atlas V 401
• Orbital (Northrop Grumman now) Antares 100
• Northrop Grumman Antares 230
• Northrop Grumman Pegasus
• SpaceX Falcon 1e
• SpaceX Falcon 9 V1.0
• SpaceX Falcon 9 V1.1
• SpaceX Falcon 9 FT
• SpaceX Starship
• Rocket Lab Electron
• Blue Origin New Glenn
• Virgin Orbit Launcher One
• Relativity Terran 1
• Astra Rocket-3
• Firefly Alpha
• ABL RS1

# The importance of payload to lift-off mass ratio

The maximum payload mass is the main driver of a rocket’s revenue potential. It may not be intuitive that the payload mass of any useful orbital rocket ranges from about 1% for the lowest performance vehicle (e.g. Rocket Lab Electron) to about 4.5% for the highest performance launch vehicles (e.g. Falcon 9 FT, Zenit) of the total rocket lift-off mass.

Example orbital rocket mass breakdown by category:

• 90% for propellant mass
• 8% for rocket mass (includes tanks, pressurization system, engines, avionics and fairing)
• 2% left for payload mass

For small launchers being developed by new startups, size is the main driver of vehicle cost. With today’s 3D printing and vertical integration methods, the cost will be about the same for two different startups in the same country to manufacture and launch two 30,000 kg lift-off mass vehicles — regardless of performance.

This makes the payload-to-lift-off mass ratio the key driver to either generate more revenue and margin or lower market prices in the same class of expandable vehicles. Reusability is another key factor to increase margin or lower prices, but it currently can only be achieved partially and through iterations funded by paying customer flights (i.e., we are not likely to see a new vehicle reach orbit on its first flight, then successfully return all stages and fairing and be available for a second flight.)

Here are the key factors that affect the payload-to-lift-off mass ratio. If you are performing due diligence on a proposed new rocket design, the following list, along with the calculator, can help you ask the right questions to better understand how aggressive or realistic a proposed rocket is.

## Engine Performance

The specific impulse (ISP) for each engine is analogous to the automotive ‘miles per gallon’ performance metric. If your rocket engine can generate the same thrust with less propellant, this saved propellant mass can be used to increase the payload — and for the second stage, this results in a payload increase even greater than just the propellant saving, as you can reclaim the tank mass that would have been needed for this additional propellant. Given that about 90% of a launch vehicle is propellant, this is the most important factor. An increase of just a few percentage points in an engine’s specific impulse can double the payload capacity.

For a two-stage rocket, you cannot evaluate the payload claim without knowing three important numbers:

• First stage engine specific impulse, at sea level: when the engine ignites on the ground at the start of the flight. This is the “1st stage Isp sea level or at the start altitude” parameter in the calculator.
• First stage engine specific impulse, in the vacuum of space. The performance at the end of the first stage flight. This is the “1st stage Isp vacuum” parameter in the calculator.
• Second stage engine specific impulse, in the vacuum of space. Even if using the same engine as the first stage (such as SpaceX Merlin), this engine will have a larger nozzle optimized for the vacuum of space and perform better than the first stage engine in space. This is the “2nd stage Isp” parameter in the calculator.

Often only a single performance number will be provided on the rocket company website, and usually the highest: when the engine is operating in the vacuum of space. However, the first stage engine with a shorter nozzle will typically perform 15% lower when the flight starts at sea level and at the end of its flight in space. Add another 10% of specific impulse for when you transition to the larger nozzle of the second stage version. Given that small increases of engine performance can have a substantial impact on the payload, ensuring that the right numbers are known is critical in evaluating any orbital rocket payload claim.

Engine performance is driven by the following key factors:

• Propellant Choice — Defines a realistic specific impulse range that can be attained. It can be complicated to compare engines with different propellants apples to apples. For example, high-performing fuels, such as Hydrogen (H₂) and Methane (CH₄) also require larger and thus heavier tanks due to their lower density.
• Engine Cycle — In an open cycle configuration, about 5% of the propellant is ‘wasted’ in the gas generator to generate the power needed to drive the turbopump turbine. Closed cycles such as oxidizer-rich staged-combustion are the highest performance, as no propellant is consumed to drive the turbopump turbine. Electric cycles have high specific impulse since no propellant is used to drive the pumps, but have the disadvantage that they can only generate relatively low chamber pressure and have extra battery mass which does not decrease as their power is consumed (Empty battery pack ejection can help mitigate but is limited to a single or double event, and these ejection systems require heavy systems and cables to switch over and eject). More cycle information can be found in this Wikipedia article.
• Combustion Pressure — Directly increases performance and reduces the engine mass and size. High combustion pressure is hard to achieve, especially in a closed cycle, given the pump pressure required.
• Oxidizer and Fuel Mixture Ratio — Each propellant mixture, at a given combustion pressure, has an optimal mixture ratio for maximum performance — typically a little lower than the stoichiometric ratio. However, this is also the ratio at which the maximum heat fluxes are generated, thereby increasing the regenerative cooling demand. The optimal mixture ratio is typically reached with the use of a combustion chamber with a copper alloy liner because it offers the highest thermal conductivity to regeneratively cool the combustion chamber. Inconel combustion chambers offer a better strength to weight ratio than copper and is more readily available from 3D printing service providers.
• Combustion Efficiency (c*) — Is a measure of how well the propellants are mixed and combusted. Low efficiency leads to wasted propellant which exits the nozzle un-reacted. An orbital launch vehicle would need to reach at least 90%, and 98%+ is considered world-class.
• Size and Mass — In addition to specific impulse, size and mass are key attributes that impact the final payload mass. Compact and light hardware is desired.

## Propellant Choice & Stage Mass

Hydrogen (H₂) and Methane (CH₄) have higher specific impulse potential than Kerosene. However, given their lower density, they require larger and thus heavier fuel tanks. Also, engines using H₂ or CH₄ are more complex and expensive to develop due to increased leak potential, more complex pumps, cryogenic issues, and risk of air/fuel explosion should a leak occur. In some cases, this combustion performance gain can have a significant impact on the payload mass ratio — especially for in-space propulsion stages where the tank mass can be maximized. For the first stage at sea level, the performance benefit can be a wash, as the tank mass cannot be fully optimized due to the need for additional mass to structurally support the upper stage, payload and fairing.

## Unused Propellants

Liquid rockets use fuel and oxidizer, each with their own tank. Towards the end of a stage burn, neither the fuel tank nor the oxidizer can run out. If either runs out, in the best case scenario the engine will shut down before the required time and impulse is reached. In the worst case, the engine turbopump will explode. Both scenarios will result in mission failure. As a result, a margin of extra propellant is planned and required. More margin is safer but less efficient in terms of payload mass loss. Typically the first flight of a new vehicle would have more margin of about 3%, but as the vehicle is proven there’s an opportunity to reduce this wasted mass further. This can be achieved with improved internal tank baffling to minimize propellant slosh, sophisticated fuel level sensors and an optimal engine mixture ratio controller to reach about 1%, which is considered world-class. Any new vehicle designer planning 0%-1% for the first flight is not likely to be announcing a realistic payload for their vehicle. This is a good first question to ask in due diligence and interviews.

## Dry Mass

How light are the tanks, engine, fairing and avionics? For the final stage, any mass saved can be directly added to the payload on a one-to-one basis. Large rockets have the advantage with dry mass. The smaller the vehicle, the harder it is to decrease the dry to wet mass ratio. Smaller rocket disadvantages include more atmospheric aerodynamic losses, sensors and computers do not shrink with the rocket size, smaller and thinner tanks are harder to optimize for low mass, smaller engines and pumps have lower efficiency and lower thrust-to-weight ratio, etc.

This calculator is open-source and available on Github. We hope to receive comments on any errors and improvement contributions and suggestions directly on Github or via the Discord chat channel.

You can verify the calculator accuracy by using widely available public information for proven rockets such as NASA Saturn V. Any proposed paper rocket in the calculator uses the same calculations. The tool is also open source for further verification or proposed enhancements.

## Here are just a few examples of what you can do with our calculator:

• Check historical and future launch vehicle performance details
• Increase or decrease engine performance by 10% and see the payload impact
• Change the launch location and target orbit for any rocket and see how it impacts the payload
• Enable Air Launch (plane) on any rocket and see the payload impact
• Build your own paper rocket design based on existing engines

## Not yet supported:

• Automatically changing the dry mass based on propellant choices
• Side booster configurations — this is why SLS, Atlas V, Space Shuttle are not yet available
• Orbital refueling — something that will help evaluate Starship for Artemis Moon mission
• Re-usability and its payload impact due to added legs, propellant, etc.

# About the Launcher Rocket Calculator author

Igor Nikishchenko, Chief Designer, Launcher

Igor previously served as Deputy Chief Designer in the Liquid Propulsion Department at Yuzhnoye, the Ukrainian state-owned company that designed the world-leading Zenit launch vehicle. Most recently, Igor worked for Avio, a key contractor for the European Space Agency’s Ariane and Vega launchers. Igor has over 30 years of experience in high-performance liquid rocket engine development.