Spacecraft Design : Propulsion Systems Design : Propulsion Systems ... From high orbit to Mars ⇒...

Spacecraft Design :Propulsion Systems

Olivier Leonard

University of LiegeTurbomachinery Group

October 2009

October 2009 à áá à ' n ? 6

Spacecraft systems

ControlBalistic

PayloadPropulsion

Telecommunications

Structures

Energy

October 2009 Slide 1/33 à áá à ' n ? 6

Spacecraft propulsion systems

n Introduction : missions and performance

n Thrust generated by a rocket engine

n Specific impulse and velocity increment

n Cold gas systems

n Mono- and bi-propellant systems

n Solid propellant systems

n Electric systems


Flight envelope


Rocket engines : various missions and performance

n From the ground to low orbit⇒ ∆V ≥ 9500 m/s

n From low orbit to high orbit⇒ ∆V ≈ 4200 m/s

n From high orbit to Mars⇒ ∆V ≈ 3400 m/s

n Escaping solar system requires an additional ∆V ≈ 8500 m/s

n On orbit control and positioning⇒ ∆V ≈ 20...400 m/s

n Launch vehicles engines providing huge thrust levels for extended periods(200 tons during 8 min for the SSME)

n Apogee and perigee engines providing moderate thrust levels(a few tons during a few seconds)

n On orbit control and positioning providing small thrust levels(mN to 10 N, pulsed operation for the whole vehicle lifespan)


Rocket engines : classification


Rocket engines : Thrust, power, acceleration

T = qmVe = Mαg0 P =1

2qmV

2e =

1

2TVe

P

M=

αg0

2Ve


Rocket engines : non chemical engines

n An external source (nuclear, solar energy) provides energy to the propellant

n The propellant may be directly or indirectly heatedand accelerated through a nozzle

n Ionized propellants are accelerated through an electrical field

n Electric rocket engines have high exhaust velocities but low specific power

n Nuclear rocket engines offer high power levels but have limited exhaustvelocities

n Solar energy is unlimited but offers a very low effective power density

⇒ Restricted to small accelerations : orbit control, spacecraft attitudes


Rocket engines : chemical engines

n A chemical reaction provides the energy

n Liquid propellants are the most energetic

n Turbopumps pressurize the combustionchamber but limit the available power

n Solid propellant engines (boosters) are simpleand compact but are energy limited

n Chemical engines are powerful but provide ratherlow Ve due to chemical and temperature limitations

n Well suited for launch vehicles

n Instabilities (chocks in the nozzle, pogo effects)


Thrust generated by a rocket engine

Thrust is the result of all internaland external forces due to pressureand viscous effects developedby the fluid on all componentsof the the engine

T =

ZΣ

p nfpax dS − F

fpv,i − F

fpv,e

T =

ZΣi

(pi − pa) nfpax dS +

ZΣe

(pe − pa) nfpax dS − F

fpv,i − F

fpv,e

At the test bench :

T =

ZΣi

(pi − pa) nfpax dS − F

fpv,i


Thrust generated by a rocket engine

Momentum balance yields(a contribution is positivewhen exerted from left to right) :

−Fpfv,i − (pe − pa)Ae +

ZΣi

(pi − pa) npfax dS = (ρV

2A)e

reulting in

T = qmVe + (pe − pa)Ae Tsl = qmVe + (pe − psl)Ae

Th = Tsl + (psl − ph)Ae T0 = Tsl + pslAe


Specific impulse

Effective ejection velocity C

T = qmC C = Ve +

»pe − pa

qm

–Ae = Ve

»1 +

1

γM2e

„1−

pa

pe

«–

Specific impulse

Isp =

Z τ

0T (t)dtZ τ

0qm(t)dt

≈T

qm= C [m/s]

Isp =

Z τ

0T (t)dt

g0

Z τ

0qm(t)dt

≈T

g0 qm=

C

g0[s]


Specific impulse


Velocity increment

Force balance

mdV

dt= T −D −mg = C qm −D −mg

dV = −Cdm

m−

D

mdt− g dt with dh = V dt

Velocity increment

∆V

C= ln

m0

m0 −mp−

Z τ

0

D

Cmdt−

Z τ

0

g

Cdt

Neglecting gravity and viscous effects :

∆V

C= ln

m0

m0 −mp

m0

m0 −mp= e

∆V/C


Secondary spacecraft propulsion

Typical functions of a secondary propulsion system:

n final orbit acquisition from the orbit established by the launch vehiclen orbit controln attitude control

Main options:

n cold gas systemsn monopropellant hydrazinen bi-propellant systemsn solid propellantsn electric systems


Cold gas systems

n Cold gas systems are based oniinert gas : nitrogen, argon, freon, propane

n Typical thrust level : 20 mN

n Typical specific impulse : 50 s

n Typical functions : high pointingaccuracies (0.1 degree)

n Gas is stored at high pressureand fed to small thrusters

n Propellant are selected for their storagesimplicity and plume compatibilitywith sensitive surfaces


Monopropellant and bi-propellant systems

n Hydrazine is stored as a liquid pressurized by an inert gas (N2 or He)

n Storage properties are similar to water

n Anhydrous hydrazine N2H4 decomposes through an exothermic process

n Decomposition is enhanced by a resistively heated metal catalyst(Pt-Ir-Al2O3)

n Hot gas products N2, H2 and NH3 expands through a nozzle


Monopropellant and bi-propellant systems

n Typical thrust level > 10 N

n Typical specific impulse : 200 s

n Typical functions : attitude control and station keeping(geostationary spacecrafts)

n Bi-propellants such as N2O4/MMH offer higher performance (Isp = 300)and mass reductions


Monopropellant systems

n Astrium CHT

n Thrust : 1 N

n Specific impulse : 210 s

n Propellant : Hydrazine

n Burn time : 50 hours

n Length : 17 cm

n Attitude control, orbit control and station keepingof small satellites and deep space probes

n Herschel, Globalstar


Monopropellant systems

n Astrium CHT

n Thrust : 400 N


n Propellant : Hydrazine

n Burn time : 30 minutes

n Length : 32 cm

n Attitude control

n Arianne V


Bipropellant systems

n Astrium S

n Thrust : 10 N


n Fuel : MMH

n Oxidizers : N2O4

n Attitude control and orbit control of large satellitesand deep space probes

n Venus Express, Arabsat


Bipropellant systems

n Astrium S

n Thrust : 400 N


n Fuel : MMH

n Oxidizers : N2O4

n Apogee orbit injectionof GEO satellitesand for planetary orbitmaneuvers of deep spaceprobes

n Venus Express, Artemis


Solid propellant systems

n Typical mission : from apogee of an elliptical transfer orbit to a circulargeostationary orbit, ∆V ≈ 2000 m/s

n Example : Intelsat V, on station mass ≈ 1000 kg, fuel mass ≈ 900 kg,engine mass ≈ 1000 kg

n Typical performance : 70 kN during 40 s, Isp ≈ 280 s


Solid propellant systems

n ATK Star

n Thrust : 27 kN


n Burn time : 34 s

n Length : 1.3 m

n Mass : 361 kg

n Apogee motor

n GOES, GPS


Electric systems

n Total spacecraft mass M = Mpay + Mpp + Me

n Power plant output W = qmV 2e /2

n Power plant output and weight are proportional Mpp = β W

n Exhaust mass flow is constant Me = qm tb

Me =M −Mpay

1 +βV 2

e

2tb

Mpp =M −Mpay

1 +2tb

βV 2e

Vc =

s2tb

β

∆V

C≈

∆V

Ve= ln

266641 +

„Ve

Vc

«2

Mpay

M+

„Ve

Vc

«2

37775October 2009 Slide 24/33 à áá à ' n ? 6

Electric systems


Electric systems

n For any payload the optimum is close to Ve ≈ Vc

n Low propellant mass requires high Ve, high Vc, long burn time and highspecific power

n Example :l payload ratio Mpp/M = 0.5l orbit raising maneuver with ∆V = 5000 m/sl Ve ≈ Vc ≈ 17000 m/sl electric engine powered by solar arrays : β ≈ 20 kg/kWl burn time tb ≈ 30 daysl thrust acceleration α ≈ ∆V/tb ≈ 2 10−4 g0


Electric systems : the resistojet

n The propellant is heated bypassing over a (tungsten)heating element

n Propellant : H2 (difficult to store),N2, NH3 (corrosive),hydrazine, teflon

n Exhaust velocity is limited bytemperature to ≈ 10 km/s

n Typical thrust : 1 N


Electric systems : electrostatic thrusters

n A cathode extracts electrons from the propellant which is ionized

n The ions are accelerated by an electric field

n Propellant : Hg, Ar, Xe

n Exhaust velocity is limited by temperature to ≈ 10 km/s

n Typical performance : T = 10 mN, Ve = 30 km/s


Electric systems : electrostatic thrusters


Low-Thrust systems

n Astrium RITA

n Thrust : 150 mN


n Propellant : Xenon

n Beam voltage : 1200 V

n Brun time : > 20000 h

n Mass : 154 kg

n Station keeping, orbit transfer,deep space trajectories

n Artemis


Propellant consumption for N/S station keeping


TOC

n Spacecraft systems v

n Spacecraft propulsion systems v

n Flight envelope v

n Rocket engines : various missions and performance v

n Rocket engines : classification v

n Rocket engines : Thrust, power, acceleration v

n Rocket engines : non chemical engines v

n Rocket engines : chemical engines v

n Thrust generated by a rocket engine v

n Specific impulse v

n Velocity increment v

n Secondary spacecraft propulsion v

n Cold gas systems v

n Monopropellant and bi-propellant systems v

n Solid propellant systems v

n Electric systems v

n Electric systems : the resistojet v

n Electric systems : electrostatic thrusters v


Spacecraft Design : Propulsion Systems Design : Propulsion Systems ... From high orbit to Mars ⇒...

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