Engine Design Reference

Everything on the Rocket Engineer Table, with the exact data and formulas the mod uses. If you only want to design an engine, the table's own readouts are enough; this page is for planning ahead and for people who like the maths.

Power packs

A power pack decides how efficiently propellant is burned.

Power pack	Combustion efficiency	Propellants it supports	How you get it
Open Cycle	0.70	Metha/LOX	Unlocked from the start
Ox Rich Flow (oxidiser-rich staged cycle)	0.90	Metha/LOX, LH2/LOX	Design blueprint, found on the Moon
Fuel Rich Flow (fuel-rich staged cycle)	0.90	Metha/LOX, LH2/LOX	Design blueprint, found on the Moon
Full Flow (full-flow staged cycle)	0.98	Metha/LOX, LH2/LOX	Design blueprint, found on the Moon
Catalyse Cycle	0.70	Premixed methalox (monopropellant)	Design blueprint, found on Mars

• Ox Rich and Fuel Rich currently have identical numbers. They differ in name only.
• The Catalyse Cycle has no assembly recipe in the current mod data, so you can unlock the design but cannot yet manufacture a power pack for it. See Known Issues.
• The data also contains a "number of pumps" field for each pack, but nothing in the code uses it. How many turbines a pack needs is set by its assembly recipe: the Open Cycle needs one, the three staged cycles need two.

Exhaust packs (nozzles)

Exhaust pack	Expansion ratio range	Propellants	How you get it	Notes
Bell Nozzle	2 – 100	all three	Unlocked from the start	The classic nozzle.
Aerospike	2 – 500	all three	Design blueprint, found on the Moon	A much wider ratio range, and lighter at the same ratio (see the mass formula).

Propellant types

Propellant	Mass ratio	Max ISP	Cp (J/kg·K)	γ (gamma)	Molar mass
Metha/LOX	2 LOX : 1 methane	460 s	4 500	1.1	18 g/mol
LH2/LOX	6 LOX : 1 hydrogen	540 s	3 000	1.2	10 g/mol
Premixed methalox	single fluid	600 s	4 500	1.1	18 g/mol
Metallic hydrogen	single fluid	6 000 s	18 280	1.3	2 g/mol (unusable, no pack allows it)

More about the fluids is on Fluids and Propellants.

Materials

Level	Material	Max pressure at 0 °C (bar)	Fails at (°C)
0	Andesite	3	900
1	Iron	50	1 200
2	Copper	80	2 200
3	Brass	150	2 500
4	Reinforced Copper	250	2 800
5	Copronickel	400	3 200
6	Monel	650	3 800
7	Inconel	900	4 500
8	Hastelloy	1 200	5 500

The rule: at a combustion temperature T (in °C), a material can hold up to

P_max(T) = P_max(0 °C) × (1 − T ÷ T_fail)

bar, and nothing at all once T reaches T_fail. Your design is valid when its chamber pressure is less than or equal to P_max(T).

Allowed pressure (bar) for common temperatures:

Material	1 000 °C	1 900 °C	2 500 °C	3 000 °C	4 500 °C	5 000 °C
Andesite	0	0	0	0	0	0
Iron	8	0	0	0	0	0
Copper	43	10	0	0	0	0
Brass	90	36	0	0	0	0
Reinforced Copper	160	80	26	0	0	0
Copronickel	275	162	87	25	0	0
Monel	478	325	222	136	0	0
Inconel	700	520	400	300	0	0
Hastelloy	981	785	654	545	218	109

(Values rounded down to whole bar, so a value that works out to exactly a whole number may be off by one in-game.)

Note

This differs from older descriptions that give a fixed "material level" from separate pressure and temperature steps. In this version the material limit is the sliding scale above.

Formulas

All of these run live in the Engineer Table screen. Symbols: η is the power pack efficiency, Isp_max the propellant's max ISP, Cp, γ and molar mass M come from the propellant table, V is the chamber volume in cubic metres (size in mB ÷ 1000), F is the thrust in newtons, and ε is the expansion ratio.

1. Combustion temperature (°C as displayed): T_c = 300 + 0.5 × (Isp_max × 9.81)² × η ÷ Cp
2. Gas constant: R_s = 8.314 ÷ (M ÷ 1000)
3. Exit Mach number from the area ratio: solve numerically, for Mach from 1 to 1 000, for ε = (1 ÷ Ma) × [ (2 ÷ (γ+1)) × (1 + (γ−1)/2 × Ma²) ] ^ ((γ+1) ÷ (2(γ−1)))
4. Exhaust temperature: T_e = T_c ÷ (1 + (γ−1)/2 × Ma²)
5. Exhaust velocity and ISP: v_e = Ma × √(γ × R_s × T_e), then ISP = v_e ÷ 9.81.
6. Throat area: A_t = V ÷ 6. The mod's own code comment says this assumes a characteristic length of 2 m and a chamber area three times the throat area.
7. Chamber pressure: P = (F ÷ ISP) × √(T_c × R_s ÷ γ) ÷ A_t × ((γ+1)/2) ^ ((γ+1) ÷ (2(γ−1))) in pascals, shown in bar (÷ 100 000).
8. Engine mass (kg): m = (ε ÷ (ε_max − ε_min) + 1) × V × 2000, where ε_min and ε_max are the nozzle's ratio range (Bell 2/100, Aerospike 2/500).
9. The blueprint stores efficiency = ISP ÷ Isp_max; the built engine's ISP is the whole number below Isp_max × efficiency, i.e. the ISP shown on the table (rounded down).

Note

In flight, the rocket treats each engine as if its ISP were multiplied by the server config ISPModifier (default 1.5). With defaults, an engine listed at 200 s behaves like a 300 s engine when working out fuel use and Δv. See Rocket Science Reference.

ISP by expansion ratio

Computed from the formulas above:

Design	Combustion T	ε = 5	ε = 10	ε = 25	ε = 50	ε = 100	ε = 500 (aerospike only)
Metha/LOX, Open Cycle	1 883 °C	224 s	248 s	274 s	289 s	303 s	329 s
Metha/LOX, Ox/Fuel Rich	2 336 °C	250 s	277 s	305 s	322 s	338 s	367 s
Metha/LOX, Full Flow	2 517 °C	259 s	287 s	316 s	335 s	351 s	381 s
LH2/LOX, Ox/Fuel Rich	4 509 °C	451 s	492 s	531 s	555 s	574 s	607 s
LH2/LOX, Full Flow	4 883 °C	470 s	512 s	553 s	577 s	597 s	631 s
Premixed, Catalyse	2 994 °C	283 s	313 s	345 s	365 s	382 s	415 s

(Rounded down as the game displays them; a value may differ by 1 s from what you see because of floating-point rounding.)

Worked examples

Chamber pressure with the expansion ratio at 50 (pressure barely depends on the power pack, but LH₂/LOX is slightly higher):

Propellant & pack	Thrust	Size	Pressure	Lowest material that works	Engine mass (bell)
Metha/LOX, Open Cycle	100 kN	1 000 mB	30 bar	Brass	3.0 t
Metha/LOX, Open Cycle	100 kN	5 000 mB	6 bar	Copper	15 t
Metha/LOX, Open Cycle	600 kN	3 000 mB	61 bar	Reinforced Copper	9.1 t
Metha/LOX, Open Cycle	600 kN	6 000 mB	30 bar	Brass	18 t
Metha/LOX, Full Flow	600 kN	6 000 mB	30 bar	Copronickel	18 t
LH2/LOX, Full Flow	600 kN	3 000 mB	64 bar	Hastelloy	9.1 t
LH2/LOX, Full Flow	100 kN	100 mB	322 bar	None (too hot and too tight)	n/a

Notice the trade-off: a hotter, more efficient engine needs a stronger metal for the same pressure, and the only way to relax the pressure is a bigger, heavier chamber.

Engine mass versus rocket weight

Engine mass is added to your rocket's dry mass, and your thrust must exceed the weight of the whole loaded rocket. A heavy low-pressure engine and a light high-pressure engine each have their place; see Rocket Science Reference for the full flight maths.