Aersp 401ASpacecraft 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 = ueln(Minitial/Mfinal)• Δv is the spacecraft velocity change

• ue is the rocket exhaust velocity

• Minitial and Mfinal are the spacecraft mass before and after the rocket firing, respectively

Spacecraft Propulsion Subsystem

• Basics of Rocketry– τ = (Δm/Δt)ue +(pe-pa)Ae

• τ is the engine thrust

• (Δm/Δt) is the mass flow rate of propellant

• ue is the rocket exhaust velocity

• pe and pa are the exhaust and ambient pressure, respectively

• Ae is the nozzle exit area

• Most thrust for a “perfectly expanded” nozzle

Spacecraft Propulsion Subsystem

• Basics of Rocketry– ueq = ue + [(pe-pa)/(Δm/Δt)]Ae

• ueq is the “equivalent exhaust velocity”

• ueq = ue for a perfectly expanded nozzle

– τ = (Δm/Δt)ueq

– Isp = ueq/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 Isp

– 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 Isp

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 Isp

• Solid Propellants– Matrix of fuel and oxidizer

– simple system

– single burn, no throttling

– moderate performance 275s Isp

Spacecraft Propulsion Subsystem

– Electric Propulsion• Performance

– Input Power = τIspg/(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 Isp

Spacecraft Propulsion Subsystem

– Electric Propulsion• Electrothermal (cont.)

– Arcjet» Ammonia, Hydrazine

» ~500-600s Isp

• 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 Isp

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