Aersp 401A Spacecraft Propulsion Subsystem (rocket science in 15 minutes)

Spacecraft Propulsion Subsystem • Uses of onboard propulsion systems – Orbit Transfer • LEO to GEO • LEO to Solar Orbit – Drag Makeup – Attitude Control – Orbit Maintenance

Spacecraft Propulsion Subsystem • Typical Mission Requirements – Orbit Transfer • Perigee Burn -- 2,400 m/s • Apogee Burn -- 1,500 - 1,800 m/s – Drag Makeup -- 60 - 1,500 m/s – Attitude Control -- 3 - 10% of total propellant – Orbit Maintenance • Orbit Correction -- 15 - 75 m/s (per year) • Stationkeeping -- 50 - 60 m/s

Spacecraft Propulsion Subsystem • Basics of Rocketry – Rocket -- Any propulsion system that carries its own reaction mass.

– Δv = u e ln(M initial /M final ) • Δv is the spacecraft velocity change • u e is the rocket exhaust velocity • M initial and M final are the spacecraft mass before and after the rocket firing, respectively

Spacecraft Propulsion Subsystem • Basics of Rocketry – τ = (Δm/Δt)u e +(p e -p a )A e • τ is the engine thrust • (Δm/Δt) is the mass flow rate of propellant • u e is the rocket exhaust velocity • p e and p a are the exhaust and ambient pressure, respectively • A e is the nozzle exit area • Most thrust for a “perfectly expanded” nozzle

Spacecraft Propulsion Subsystem • Basics of Rocketry – u eq = u e • u eq + [(p e -p a )/(Δm/Δt)]A e is the “equivalent exhaust velocity” • u eq = u e for a perfectly expanded nozzle – τ = (Δm/Δt)u eq – I sp = u eq /g • Specific impulse is a measure of thrust per propellant mass flow rate • g is always gravity at Earth’s surface, not local

Spacecraft Propulsion Subsystem – Chemical Rockets • Performance is energy limited • Propellant Selection – Maximum Performance – Density – Storage (i.e. cryogenic) – Heat transfer properties – Toxicity and corrosivity – Viscosity – Availability (cost)

Spacecraft Propulsion Subsystem – Chemical Rockets • Cold Gas Systems – pressurized gas flowing through a nozzle, no reaction – very low performance -- 30-70s I sp – very simple, inexpensive system • Monopropellant Liquid Systems – Single substance with a catalyst – hydrazine, hydrogen peroxide with metal catalysts -- silver, rhodium, platinum – physically simple system – 200-225s I sp

Spacecraft Propulsion Subsystem – Chemical Rockets • Bipropellant Liquid Systems – liquid fuel -- hydrocarbons, kerosene or alcohol based – liquid oxidizer -- oxygen, nitrogen tetroxide – more complex pumping/feed systems – better performance -- 300-450s I sp • Solid Propellants – Matrix of fuel and oxidizer – simple system – single burn, no throttling – moderate performance 275s I sp

Spacecraft Propulsion Subsystem – Electric Propulsion • Performance – Input Power = τI sp g/(2η) – η is efficiency (Kinetic Energy/Input Power) • Electrothermal – Electrical energy is used to heat the propellant to high temperature, and then gas is expanded through a nozzle.

– Resistojet » Ammonia, Water » ~300s I sp

Spacecraft Propulsion Subsystem – Electric Propulsion • Electrothermal (cont.) – Arcjet » Ammonia, Hydrazine » ~500-600s I sp • Electrostatic – Electrical energy is used to accelerate charged particles with a static electric field – Ion Engine » Xenon, Krypton » 2,500-10,000s Isp

Spacecraft Propulsion Subsystem – Electric Propulsion • Electromagnetic – Combination of steady or transient electric and magnetic fields used to accelerate charged particles – Pulsed plasma thruster » Teflon » 850-1200s I sp

Spacecraft Propulsion Subsystem – System Selection and Sizing (Table 17.2) 1) Determine propulsion functions -- table 17.1

2) Determine Δv and thrust levels needed -- sec. 7.3, sec. 10.3

3) Determine subsystem options -- ch. 17 4) Estimate Isp, thrust, mass, volume for each option 5) Establish baseline subsystem

Spacecraft Propulsion Subsystem – References • Hill and Peterson, Mechanics and Thermodynamics of Propulsion • Sutton, Rocket Propulsion Elements • Micci and Ketsdever, eds., Micropropulsion for Small Spacecraft.

• Aersp 430, 530