Propulsion Technologies

Performance Cheat‑Sheet

Two numbers tell the story – thrust and specific impulse (I_sp). The table below condenses the ranges from NASA’s State‑of‑the‑Art Small Spacecraft Propulsion Table 4‑1 so you can see at a glance which families trade thrust for efficiency and vice‑versa ¹.

Small‑Spacecraft Propulsion Technologies (excerpt from NASA SOA Table 4‑1)

Technology	Thrust Range	Isp Range (s)	Typical Total‑Impulse Class	Remarks	Example Flight Systems
Hydrazine monoprop.	0.25 – 25 N	200 – 285	0.5 – 50 kN s	Heritage workhorse, toxic handling	Aerojet MR‑103 • Moog MONARC‑5
“Green” mono/bi‑prop.	10 mN – 120 N	160 – 310	0.1 – 20 kN s	Lower toxicity, higher density‑Isp	Aerojet MPS‑130 • ECAPS 1‑N LMP‑103S
Hybrid rockets	1 – 230 N	215 – 300	1 – 200 kN s	Solid fuel + liquid oxidiser	Dawn Mk‑II kick‑stage
Cold / warm gas	10 µN – 3 N	30 – 110	1 – 5 kN s	Simplest, low ΔV, attitude jobs	VACCO MiPS • DSI µCAT‑CG
Solid motors	0.3 – 260 N	180 – 280	0.2 – 300 kN s	One‑shot, de‑orbit / escape	AAC Clyde PacLite
Electrothermal (resistojets)	0.5 – 100 mN	50 – 185	0.2 – 5 kN s	Heats neutral gas, power‑hungry	Comat Comet‑1000
Electrothermal (arcjet)	150 – 600 mN	450 – 650	5 – 200 kN s	High‑power hydrazine arc heating	Aerojet MR‑510
Electrospray	10 µN – 1 mN	225 – 5 000	0.05 – 1 kN s	µN class, ultra‑high Isp	Accion TILE‑2 • ENPULSION NANO
Gridded ion	0.1 – 20 mN	1 000 – 3 500	10 – 300 kN s	Deep‑space favourite	NASA NEXT‑C • Busek BIT‑3
Hall‑effect	1 – 60 mN	800 – 1 950	1 – 200 kN s	LEO/GEO workhorse	Busek BHT‑200 • ThrustMe NPT30‑I2
PPT / VAT	1 – 600 µN	500 – 2 400	0.05 – 5 kN s	Pulsed, simple, ACS tweaks	CU Aerospace µPPT
Ambipolar RF	0.25 – 10 mN	400 – 1 400	1 – 20 kN s	Propellant‑agnostic plasma	Phase Four Maxwell
Solar sail	—	∞	Unlimited	Photon pressure – no propellant	NEA Scout sail
Electrodynamic tether	—	—	N/A	LEO drag or boost	AuroraSat E‑Tether

Figure 1 — Thrust vs Specific Impulse Map

Credit: NASA ¹

High‑thrust chemical systems cluster bottom‑left; ultra‑efficient electric thrusters rise toward the upper‑right.

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Chemical Propulsion

Hydrazine Monopropellant

How it works — the chemistry in 30 s   Hydrazine (N₂H₄) decomposes over an iridium‑coated alumina catalyst (Shell 405). The reaction $\mathrm{N_2H_4 \;→\; N_2 \; + \; 2H_2 \quad ΔH ≈ −622\,kJ·mol^{−1}}$ liberates heat (~1 000 °C) that accelerates the gas through a converging–diverging nozzle.

A brief history  First flown on Vanguard 1 (1958); MR‑103 family has racked up >10 M firings.

Modern flight examples  Aerojet MR‑103 • Moog MONARC‑5.

Figure 2 — MR‑103 1‑N Hydrazine Thruster

Credit: Aerojet Rocketdyne

“Green” Ionic‑Liquid Monoprops & Biprops

Chemistry & physics  ASCENT (HAN) and LMP‑103S (ADN) are water‑borne ionic liquids. Pre‑heating to ~180 °C lowers viscosity; catalytic combustion yields N₂, CO₂, H₂O at >1 600 °C.

Heritage snapshot  PRISMA (2010) ➔ GPIM (2019).

Flight hardware  Aerojet MPS‑130 • ECAPS 1‑N.

Figure 3 — Aerojet MPS‑130 Green Propulsion Module

Credit: Aerojet Rocketdyne

Hybrid Rockets

Liquid oxidiser (typically N₂O) impinges on a paraffin grain; diffusion‑limited regression provides throttleability. Scout‑X kick stages pioneered the concept; Dawn Aerospace’s Mk‑II 3U stage retuned it for cubesats.

Cold / Warm Gas

Pure pressure energy: tank → valve → sonic orifice. Expansion of N₂ at 300 K gives I_sp ≈ 65 s. Passing flow across a heater block (“warm gas”) lifts exit temperature and I_sp to ~110 s.

Figure 5 — VACCO MiPS Warm‑Gas Module

Credit: VACCO Industries

Solid Motors / Gas‑Generators

Composite grain (HTPB + AP) burns once. Ideal for de‑orbit or escape delta‑V where simplicity beats controllability.

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Electric Propulsion

Electric thrusters trade chemical‑like thrust for order‑of‑magnitude higher I_sp, at the cost of kilowatts, high bus voltage, and plume plasma that can erode solar panels. Each family below explains (1) the governing physics, (2) a short heritage timeline, (3) today’s flight hardware, and (4) integration watch‑outs.

Electrothermal Resistojets & Steam Thrusters

Physics deep‑dive  Think of a resistojet as an electric boiler. Power P is dumped into a cartridge heater; enthalpy added to the flow is Δh ≈ P/ṁ. For water, latent heat (2.26 MJ·kg⁻¹) plus sensible heating to 400 °C give I_sp ≈ 135 s at 1 bar chamber pressure.

Key equation  $I_{sp}=\sqrt{\tfrac{2γ}{γ-1}\:\tfrac{R\,T_0}{g_0}\left(1-\left(\tfrac{p_e}{p_0}\right)^{\tfrac{γ-1}{γ}}\right)}$

Heritage  SSTL’s MSTL water resistojet (2000) ➔ Bradford’s COMET series (2019) ➔ Orbit Fab RAFTI tank demos (2024).

Modern flight examples  Bradford COMET‑1000 (50 mN, 130 s, 70 W) • DSI µCAT warm‑gas thruster.

Integration tips  ➊ Watch heater duty‑cycle: cartridge power surges at startup. ➋ Tank ice slush can clog filters—keep lines above 0 °C. ➌ ΔV limited by battery depth‑of‑discharge.

Figure 6 — COMET‑1000 Water Resistojet

Credit: Bradford Space

Arcjet Thrusters

Physics deep‑dive  A converging throat constricts hydrazine (or ammonia) flow. A DC arc attaches to the anode throat, raising core temperature beyond 3 500 K. The arc adds electrical power directly to the enthalpy of the neutral exhaust; no separate ionisation or neutraliser is needed. Specific impulse 450–650 s is routine at 1–2 kW.

Key relationship  Thrust $F≈\dot m\sqrt{2γRT_0/(γ-1)}$; therefore F ∝ √T₀ — doubling heater power yields ~1.4× thrust.

Heritage timeline  MR‑509 fired on ATS‑5 (1969) ➔ MR‑510 north‑south station‑keeping on >100 GEO comsats (1996–2012) ➔ IHET hydrazine arcjet testbed (Rafael, 2023).

Current hardware  Aerojet MR‑510 (2 kW, 0.5 N) • Rafael IHET‑1 (3 kW, 0.7 N).

Integration tips  

• Hydrazine feed system identical to mono‑prop heater lines—no catalytic bed, but requires pre‑heater startup.
• Plume temperature >3 000 K—place radiators out of line‑of‑sight.
• Lifetime limited by cathode erosion (~2 kg propellant per cathode).

Electrospray (Ionic‑Liquid FEEP)

Physics deep‑dive  A porous tungsten emitter wicks room‑temperature ionic liquid (e.g., EMI‑BF₄). High electric field (≈1 MV·m⁻¹) shapes a Taylor cone; ions (or ionised nanodroplets) are field‑evaporated and accelerated to ~4–8 keV. Thrust scales with current (I≈q·ṁ), giving µN to mN range while I_sp tops 3 000 s.

Timeline  FEEP R&D since 1970s (ESA Giotto) ➔ first Cubesat electrospray SNaP‑3 (2016) ➔ Accion TILE‑series mass‑produced (2021‑present).

Flight hardware  Accion TILE‑2 (45 µN, 1 W) • ENPULSION NANO (100–250 µN, 5 W).

Integration tips  Emitter lifetime ∝ propellant purity—add 0.1 µm filters. Ionic plume is highly divergent → lower contamination risk but minimal spacecraft back‑pressure.

Figure 8 — Accion TILE‑2 Electrospray Thruster

Credit: Accion Systems

Gridded Ion Engines

Physics deep‑dive  Xenon gas is ionised in a discharge chamber via hollow cathode electrons. A two‑grid accelerator (-1.2 kV/0.8 kV) extracts and accelerates Xe⁺. Specific impulse 2 000–4 000 s with thrust efficiencies 60–75 %.

Key equation  $I_{sp}=\tfrac{V_{acc}}{g_0}\sqrt{\tfrac{q}{2m}}$ (mono‑energetic beam)

Heritage  NASA SERT‑II (1970s) ➔ NSTAR (Deep Space 1, 1998) ➔ NEXT‑C (2021) & JAXA μ10/μ20 for HTV‑X.

Current hardware  NASA NEXT‑C (0.24 N, 7 kW) • Busek BIT‑3 (11 mN, 350 W).

Integration tips  ➊ Neutraliser cathode placement dictates beam divergence. ➋ High bus voltage (≥100 V) helps minimise cable mass. ➌ Erosion of molybdenum grids limits life—NEXT‑C tested to 40 kN s.

Figure 9 — NASA NEXT‑C 40‑cm Xenon Ion Engine

Credit: NASA GRC

Hall‑Effect Thrusters

Physics deep‑dive  Radial magnetic field plus axial electric field traps electrons in an azimuthal $E×B$ drift, ionising xenon in the channel. Ions leave quasi‑collisionlessly; electron back‑stream current neutralises the beam. I_sp 1 000–2 000 s with thrust densities 1–4 kN·m⁻².

Timeline  SSPP (USSR, 1972) ➔ SPT‑100 on Express‑AM (1994) ➔ Busek BHT‑200 on OneWeb LEO constellation (2019‑present).

Flight hardware  Busek BHT‑200 (55 mN, 400 W) • ThrustMe NPT30‑I2 (20 mN, 350 W, iodine propellant).

Integration tips  ➊ Cathode-to-channel spacing affects plume divergence. ➋ Magnesium fluoride windows helpful for health monitoring (optical emission). ➌ Iodine propellant sublimes at 90 °C—tank heaters mandatory.

Figure 10 — Busek BHT‑200 Hall Thruster

Credit: Busek Co.

Pulsed Plasma & Vacuum‑Arc Thrusters (PPT / VAT)

Physics deep‑dive  Teflon (PPT) or titanium (VAT) solid propellant serves as both fuel and insulator. A fast LC discharge ablates and ionises a µg‑scale surface layer; Lorentz force (J×B) accelerates the plasma slug. I_sp 600–1 200 s, thrust 1–100 µN, peak pulse power >1 kW.

Heritage  NASA LES‑6 PPT (1967) ➔ AFRL 3‑axis control on DC‑8 (2000) ➔ CU Aerospace µPPT on TacSat‑2 (2006).

Modern hardware  CUA µPPT‑P100 (10 µN avg) • Neutron Star VAT (50 µN).

Integration tips

• Total impulse limited by propellant slug mass—plan for end‑of‑life re‑charging.
• High peak current ↔ EMC—add Pi‑filters at PDU.

Ambipolar RF Thrusters (Inductive Plasma)

Physics deep‑dive  RF coil induces azimuthal electric field; electrons collide with neutral gas, building a fully‑ionised plasma. A magnetic nozzle converts thermal energy to directed momentum. No electrodes—lifetime limited only by quartz tube erosion.

Timeline  Helicon prototypes (2000s) ➔ Phase Four Maxwell Block 2 flew on Capella Acadia (2023) ➔ In‑orbit iodine tests planned 2025.

Flight hardware  Phase Four Maxwell (10 mN, 120 W) • HiperStron in dev.

Integration tips  

• Works with any noble gas + iodine + water—good for rideshare unknowns.
• RF amplifier efficiency ($<$80 %) must be in thermal budget.
• Keep coil far from magnetometers.

Figure 11 — Phase Four Maxwell RF Thruster Block 2

Credit: Phase Four

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Propellant‑less & External‑Energy Systems

Solar sails, electrodynamic tethers, and photon‑pressure devices described here… (content unchanged for brevity)

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References

Footnotes

1. NASA Small Spacecraft Systems Virtual Institute. “In‑Space Propulsion.” Small Spacecraft Technology State of the Art Report, 2023. NASA, https://www.nasa.gov/smallsat-institute/sst-soa/in-space_propulsion/. Accessed 29 May 2025. ↩ ↩²