NovaSolver›Ion Thruster (Electric Propulsion) Isp & Thrust Simulator Back
Space Engineering
Ion Thruster (Electric Propulsion) Isp & Thrust Simulator
Size a gridded electrostatic ion thruster: see exit velocity, specific impulse Isp, thrust and thrust-to-power ratio in real time. Vary the beam voltage, beam current, propellant atomic mass, charge state and propellant utilization to explore high-Isp regimes that chemical rockets cannot reach.
Parameters
Propellant
Sets the atomic mass automatically (slider remains editable)
Beam voltage V_b
V
Grid potential difference. Mostly governs Isp
Beam current I_b
A
Ion current carried by the beam. Mostly governs thrust
Propellant atomic mass m
amu
Charge state q
+
Singly-charged (1+) dominates; multiply-charged species reduce efficiency
Propellant utilization η
%
Effective utilization including neutral leak
Results
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Exit velocity v_e (km/s)
—
Specific impulse Isp (s)
—
Effective Isp (with η, s)
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Thrust T (mN)
—
Beam power (W)
—
Thrust/power (mN/kW)
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Ion thruster cross-section — acceleration animation
Ions produced in the discharge chamber are accelerated between the screen and accel grids and ejected as a blue beam. Electrons supplied by the downstream neutralizer neutralize the beam current.
From energy conservation (1/2)m_iv_e² = qV_b, exit velocity v_e, divided by g_0 = 9.81 m/s² for the specific impulse. Thrust T is the product of ion mass flow rate ṁ_i and v_e.
Ion mass flow rate ṁ_i follows from beam current I_b; propellant utilization η converts it to the total mass flow. η < 1 means the effective Isp is below v_e/g_0.
Thrust-to-power scales inversely with v_e. High Isp means large Δv for the same propellant, but thrust is ṁ·v_e and the power budget limits it (higher Isp → lower T/P).
Ion Thruster (Electric Propulsion) — Specific Impulse and Thrust
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Professor, I heard that the Hayabusa ion thrusters only produce about 10 mN. That can't even lift a coin on Earth — how can they move a spacecraft?
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Good question. On the ground gravity is 9.8 m/s² and you fight air drag, so 10 mN moves nothing. But space is microgravity and vacuum, so as long as you can accelerate "a little bit, for a long time", you eventually build up a large velocity change Δv. Hayabusa fired its μ10 ion engine for a total of 40,000 hours and pulled off a sample return from asteroid Itokawa that a chemical rocket simply could not do. The trick is not the thrust magnitude but the high specific impulse Isp.
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That Isp value moves a lot when I drag the sliders. A chemical rocket is around 450 s, but the default here shows 4793 s — that's an order of magnitude!
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Right. A chemical rocket is capped by the thermal speed of combustion gas (max ~4.5 km/s), but an ion thruster accelerates individual ions in an electric field, so at a beam voltage of 1.5 kV a Xe ion reaches 47 km/s. The formula is simply v_e = √(2qV/m): raise V and v_e grows, and Isp is just v_e divided by g_0. As long as you have a few kW of electrical power, getting 10× the Isp of a chemical engine is easy.
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Then just crank the voltage and get Isp of 10,000 s, right? What's the catch?
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That is the central trade-off of electric propulsion. If the power P is fixed, T/P = 2η/v_e — raising v_e drops thrust. Default settings give T/P ≈ 30 mN/kW; push voltage to 5 kV and Isp rises but thrust-to-power collapses. Add the fact that higher voltage stresses grid insulation and sputter lifetime, and the practical optimum Isp comes from the mission Δv requirement and the solar-array power you can carry.
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When I switched the propellant from Xe to Ar, the thrust dropped by more than half. Why does a lighter atom give less thrust?
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At the same voltage, v_e = √(2qV/m) so v_e gains about √(m_Xe/m_Ar) ≈ 1.8×, but thrust T = (I_b/qe)·m·v_e is proportional to m, which only gets back m^(1/2). So T falls as √m. Argon is cheap and stores well per volume, but the lower thrust-to-power makes it less suited for large GEO transfers or long maneuvers. SpaceX picked Kr for Starlink to cut propellant cost to a tenth of Xe, and Ar is now being studied for Starship-class missions.
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I'm also curious about propellant utilization η. Going from 85% to 95% raises the effective Isp a lot. What is it exactly?
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η is "the fraction of neutral gas supplied that gets ionized and leaves as the beam". The rest leaks out as neutrals with zero thrust. Low η means you are literally throwing away fuel, so effective Isp = T/(ṁ_total·g_0) drops. Designers optimize the discharge-chamber magnetic confinement and tighten the upstream pressure to raise η, but push too far and the discharge becomes unstable, so 80-92% is the usual operating band. NASA's NSTAR and Japan's μ10 both live in that range.
Frequently Asked Questions
Specific impulse is the exit velocity divided by standard gravity (Isp = v_e/g₀). A chemical rocket is limited by combustion thermal speed (max around 4.5 km/s), while an ion thruster accelerates individual ions in an electric field — a 1.5 kV beam voltage pushes Xe ions to about 47 km/s. From the energy balance (1/2)m·v² = qV, v_e = √(2qV/m), so increasing voltage or using a lighter atom increases v_e, and Isp reaches a few thousand to ten thousand seconds.
Even small thrust T = ṁ·v_e is useful because it can fire continuously for very long times and needs very little propellant mass. From Tsiolkovsky Δv = Isp·g₀·ln(m₀/m_f), a 10× higher Isp gives the same Δv with under one-tenth the propellant. Deep-space probes (Hayabusa, Dawn) and all-electric GEO satellites integrate mN-level thrust over months or years to reach several km/s of Δv, drastically cutting launch mass and cost.
Xe meets every requirement: (1) its atomic mass of 131 with single-charge state gives a good thrust-to-power ratio, (2) its first ionization energy of 12.13 eV makes plasma generation easy, (3) it is monoatomic with no molecular dissociation, (4) it can be stored as a liquid at room temperature under pressure, and (5) it is chemically inert. Alternatives have appeared — cheaper Kr (used on Starlink), lighter Ar with higher v_e, solid-storage iodine (I) for CubeSats — but Xe still dominates for the Isp-versus-thrust-density balance.
Propellant utilization η = ṁ_ion / ṁ_total is the fraction of neutral gas supplied to the discharge chamber that gets ionized and leaves as the beam. The rest leaks out as neutrals and produces no thrust. It never reaches 100% because of (1) neutrals diffusing downstream before being ionized, (2) wall recombination, and (3) energy loss to multiply-charged ions. NSTAR and μ10-class operating points typically run at η = 80-92%. Lower η means a bigger propellant tank for the same thrust and longer mission time.
Real-World Applications
Deep-space probes: NASA's Deep Space 1 (1998, NSTAR 2.3 kW, Xe, Isp 3100 s) was the first major flight, followed by Dawn (orbiting asteroid Vesta and dwarf planet Ceres) and JAXA's Hayabusa and Hayabusa2 (microwave-discharge μ10, Xe, Isp ~3000 s). Each spent years firing continuously to accumulate several km/s of Δv and reach destinations that would be impossible for chemical rockets.
GEO transfer and station-keeping: "All-electric" geostationary satellites (Boeing 702SP, Airbus Eurostar Neo and others) replace chemical thrusters with electric propulsion and cut launch mass by 40%. The GTO-to-GEO transfer takes around six months, but the much smaller propellant tank allows two satellites to share a single launch slot, reducing operating cost. Hall thrusters dominate this segment, with Isp around 1600-2000 s.
LEO mega-constellations: Every SpaceX Starlink satellite carries a krypton Hall thruster for orbit raising, station-keeping, drag compensation, and end-of-life deorbit. Choosing krypton — far cheaper than xenon — is what makes manufacturing thousands of satellites per year viable, and future generations may switch to argon to push propellant cost even lower.
Micro-electric propulsion for CubeSats: CubeSat-class missions now use iodine (I) miniature ion engines (e.g. France's ThrustMe NPT30) that store propellant as a solid, or field-emission electric propulsion (FEEP). The tiny tank volume of solid iodine lets even a 6U CubeSat perform real orbit changes — a breakthrough enabling new mission classes.
Common Misconceptions and Pitfalls
The biggest misconception is that "high Isp is automatically the best propulsion choice". There is always a trade-off between Δv = Isp·g₀·ln(m₀/m_f) and burn time t = Δv·m̄/T. Doubling Isp halves propellant mass but, at fixed power, halves the thrust as well, so the same Δv takes twice as long. Whether six months for a GEO transfer is acceptable is a business decision for the satellite operator. For deep-space missions ("plenty of time, tight mass budget") high Isp wins; for time-critical near-Earth maneuvers, medium-Isp / high-thrust solutions or chemical-electric hybrids are chosen.
Next, treating "beam power = total engine power consumption". This tool assumes a 70% conversion efficiency (P_total = P_beam/0.7), but a real spacecraft must also feed the discharge, the neutralizer, the magnetic coils and the losses in the Power Processing Unit (PPU). Total bus power demand is typically 1.3 to 1.5 times the beam power. NSTAR, for example, ran a 2.3 kW beam with PPU input of about 2.5 kW. Always budget the bus side with an efficiency multiplier.
Finally, "ignoring grid lifetime when projecting burn time". Neutral xenon escaping the discharge chamber is re-ionized and sputters the screen and accel grids, enlarging their apertures over time until the thruster stops working. NSTAR achieved 30,000 hours, μ10 around 18,000 hours, and commercial Hall thrusters about 10,000 hours. The thrust and Isp shown by this tool are instantaneous; performance degradation as a function of accumulated burn (fluence) must be characterised by separate life tests. NASA and JAXA practice is to demonstrate twice the design life on the ground before flight.
How to Use
Enter beam voltage (200–3000 V typical for gridded ion thrusters; higher voltage increases exit velocity)
Input beam current (0.5–2.0 A range; proportional to mass flow rate and thrust)
Specify propellant mass flow rate (mg/s; xenon typical at 10–50 mg/s for deep-space missions)
Set charge state (usually +1 for singly ionized xenon; +2 for doubly ionized propellant increases exhaust velocity)
Read exit velocity, Isp, effective Isp (accounting for efficiency), thrust in mN, beam power in watts, and thrust-to-power ratio
Worked Example
Dawn spacecraft ion thruster baseline: 1200 V beam voltage, 1.3 A beam current, 5 mg/s xenon mass flow, charge state +1. Exit velocity ≈ 30 km/s, Isp ≈ 3050 s, thrust ≈ 92 mN, beam power ≈ 1560 W, thrust-to-power ≈ 59 mN/kW. Increasing voltage to 3000 V and charge state to +2 raises exit velocity to 47 km/s and Isp to 4800 s, typical for advanced gridless (Hall-effect) thrusters.
Practical Notes
Xenon propellant: atomic mass 131 amu; use charge state +1 unless modeling advanced ion sources. Iodine (127 amu, lower ionization cost) emerging for CubeSats.