Overview of Propulsion Systems

argumentwildlifeUrban and Civil

Nov 16, 2013 (3 years and 9 months ago)

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MAE 5391: Rocket Propulsion

Overview of
Propulsion Systems





2

Rocket Technologies




3

Propulsion Technology Options


Thermodynamic Systems (TE KE)


Cold Gas Thrusters


Liquids


Monopropellants


Bipropellants


Solids


Hybrids


Nuclear (NE TE KE)


Electric Systems


Electrothermal (Resistance Heating)


Electrostatic (Ion with E field F=qE)


Electromagnetic (plasma with B field F=JxB)


With the exception of electrostatic and electromagnetic, all use
concept of gas at some temp flowing though a
converging/diverging nozzle!



4

Chemical Limitations


Why we have thermo!

]
)
(
1
[
)
1
(
2
/
1
0
0













p
p
M
T
R
V
e
u
exit









V
exit
= nozzle exit velocity (m/s)

R
u
= universal gas constant (8314.41 J/kmol*K)

T
0
= chamber temperature (K)

P
e
= exit pressure (Pa)

P
0
= chamber pressure (Pa)

M= molecular mass of gas (kg/kmol)


= ratio of specific heats (no dimensions)





5

Cold Gas

Gas

Molecular

Weight

Specific

Impulse (sec)

Air

28.9

74

Argon

39.9

57

CO
2

44.0

67

Helium

4.0

179

Hydrogen

2.0

296

Nitrogen

28.0

80

Methane

16.0

114

1.5 litre X 600 bar
Nitrogen tanks
Fill/drain valve
Two stage regulator
(feed pressure ~ 4
bar)
Thruster (0.01
N,
1.3 *10
-5
kg/s,
throat diameter
0.0133 cm)
Stop valve
Microsat cold gas
propulsion system
layout proposal
Cold Gas: Expand a pressurized gas through a nozzle






6

Liquid Monopropellant

Parameter

Value

Catalyst

LCH 227/202

Steady
-
state thrust (N)


11.1
-

31.2

Isp

(sec)

228
-

235

Propellant specific gravity

1.023

Average Density
Isp

( sec)

236.8

Rated total impulse (Nsec)

124,700

Total pulses

12,405

Minimum impulse bit (Nsec)

0.56

Feed pressure (bar)

6.7
-

24.1

Chamber pressure (bar)

4.5
-

12.4

Nozzle expansion ratio

61:1

Mass flow rate (gm/sec)

5.0
-

13.1

Valve power

27 W max @ 28 VDC

Thruster mass (kg)

0.52

3 N
2
H
4



4 NH
3

+ N
2

+ 336,280 joules

MonoProp: Decompose a single
propellant and expand the exhaust
through a nozzle



7

Liquid Bi
-
Propellant

Storable

Isp 250
-
320 sec


finert=0.03
-
.13

Cryogenic Isp 320


452 sec


finert=0.09
-
0.2

BiProp: Combust (burn) two propellants (fuel +
oxidizer) in a combustion chamber and expand
exhaust through a nozzle


Finert = 0.04
-
0.2

Finert=0.11
-
0.31



8

Solids


Composite

propellant, consisting of an
oxidizing agent, such as
ammonium
nitrate

or
ammonium perchlorate

intimately mixed with an organic or
metallic fuel and binder.

Thrust function of burn area, Isp = 250
-
300 sec

Finert=0.06
-
0.38, 2/3 of motors have fiinert below 0.2

Advantages

Simple

Reliable

High density Isp

No chamber cooling


Disadvantages

Cracks=disaster

Can’t restart

Hard to stop

Modest Isp



9

When solids go bad!



10

Hybrids

Isp= 290
-
350 sec

Finert=0.2

Hybrid: Bipropellant system with liquid oxidizer (usually) and a solid fuel

Catalyst
Pack

Combustion
Chamber

Nozzle

Test Stand

Load Cell

Fuel
Element

H
2
O
2
/PE
Hybrid Test
Set
-
Up

Polyethylene fuel rod



11

Nuclear Thermal Propulsion

NERVA Program


Thrust = 890,000N


Isp = 838 sec


Working fluid = Hydrogen


Test time = 30 minutes


Stopped in 1972


Finert=0.5
-
0.7 (shielding)



12

Electrothermal
-
Resistojets

Nozzle
Thermocouple tapping
Stainless steel outer
1225W Cartridge heater
Water inlet
Heater thermocouple
Power input
Sintered stainless
water distribution ring
Sintered stainless filter
Pressure tapping
SiC Heat transfer medium
Cutaway of Mark- III Resistojet
Working
Fluid

Thrust (mN)

Isp (sec)

Power (W)

Cp (kJ/kg K)

Tc (K)

hydrogen

37

546

100

14.32

1000

water

93

219

100

2.3

1000

nitrous oxide

141

144

100

1.0

1000

Electrothermal
--

electrical energy is used to directly heat a working fluid. The resulting hot
gas is then expanded through a converging
-
diverging nozzle to achieve high exhaust
velocities. These systems convert thermal energy to kinetic energy





13

Electrothermal
-
Arcjets

In an arcjet, the working gas is injected in a chamber through which an
electric arc is struck. The gas is heated to very high temperature (3000


4000 K), Arc temp =10,000K on average, and much greater in certain
regions in the arc.



Power = 1.8 kW, Isp = 502, Thrust = 0.2N, Propellant = hydrazine



14

Electrostatic
-
Ion Propulsion


Electrostatic
--

electrical energy is directly converted into
kinetic energy. Electrostatic forces are applied to charged
particles to accelerate the propellant.

Deep Space 1 = 4.2 kW, Thrust = 165 mN, Isp = 3800 sec

7000 hours of operation is becoming the standard!



15

Electromagnetic
-
MPD Thruster


Electromagnetic
--


electromagnetic forces directly
accelerate the reaction mass. This is done by the
interaction of electric and magnetic fields on a highly
ionised propellant plasma.

NH
3

MPD, F=23 mN, Isp= 600 sec, P=430 W


Stuttgart, Isp=5000sec, F=100N, P=6 MW, hydrogen



16

Pulsed Plasma Thrusters

C
trigger
C
Main
R
trigger
Center
Electrode
Intermediate
Electrode
Outer
Electrode
Teflon Annulus
PPU
Spacecraft
Ground
Isp = 500
-
1500 sec

P = 1


100 W

Thrust = 5
m
N/W



17

Hall Effect Thruster

Power = 50W


25kW

Isp = 500


3000 sec

Thrust = 5 mN
-

1N



18

Propulsion System “Cost”


Performance issues


Mass


Volume


Time (thrust)


Power


Safety


Logistics


Integration


Technical Risk


The “best” (lowest “cost”)
option optimizes these issues
for a given set of mission
requirements