Propulsion System Overview

Report
MAE 5391: Rocket Propulsi
Overview of Propulsion Sys
Rocket Technologies
2
Propulsion Technology Options

Thermodynamic Systems (TE


KE)
Cold Gas Thrusters
Liquids
• Monopropellants
• Bipropellants




Nuclear (NE TE
Electric Systems




Solids
Hybrids
KE)
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!
3
Chemical Limitations

Why we have thermo!
V exit 
2   Ru T 0
(  1)  M
 [1  (
pe
)
 1 / 
]
p0
Vexit= nozzle exit velocity (m/s)
Ru= universal gas constant (8314.41 J/kmol*K)
T0= chamber temperature (K)
Pe= exit pressure (Pa)
P0= chamber pressure (Pa)
M= molecular mass of gas (kg/kmol)
= ratio of specific heats (no dimensions)
4
Cold Gas
Cold Gas: Expand a pressurized gas through a nozzle
Gas
M icro sat c old gas
prop ulsion syste m
layo ut propo s al
1.5 litre X 60 0 b a r
N itro ge n ta nk s
T w o s ta ge re gula to r
(fee d p re s su re ~ 4
ba r)
T hru ste r (0 .01
N,
1.3 *10 -5 kg /s,
th ro a t d iam eter
0.0 13 3 cm )
F ill/d rain v alv e
S top va lve
Molecular
Weight
Specific
Impulse (sec)
Air
28.9
74
Argon
39.9
57
CO2
44.0
67
Helium
4.0
179
Hydrogen
2.0
296
Nitrogen
28.0
80
Methane
16.0
114
5
Liquid Monopropellant
MonoProp: Decompose a single
propellant and expand the exhaust
through a nozzle
Parameter
Catalyst
Value
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
Thruster mass (kg)
27 W max @ 28 VDC
0.52
3 N2H4  4 NH3 + N2 + 336,280 joules
6
Liquid Bi-Propellant
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
Storable Isp 250-320 sec
finert=0.03-.13
Cryogenic Isp 320 – 452 sec
finert=0.09-0.2
7
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.
Advantages
Simple
Reliable
High density Isp
No chamber cooling
Disadvantages
Cracks=disaster
Can’t restart
Hard to stop
Modest Isp
Thrust function of burn area, Isp = 250-300 sec
Finert=0.06-0.38, 2/3 of motors have fiinert below 0.2
8
When solids go bad!
9
Hybrids
Hybrid: Bipropellant system with liquid oxidizer (usually) and a solid fuel
Isp= 290-350 sec
Finert=0.2
Load Cell
Catalyst
Pack
Test Stand
Fuel
Element
Combustion
Chamber
Nozzle
Polyethylene fuel rod
H2O2/PE
Hybrid Test
Set-Up
10
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)
11
Electrothermal-Resistojets
C u ta w a y o f M a rk- III R e sisto je t
T h e rm o c o u p le ta p p in g
1 2 2 5 W C a rtrid g e h e a te r
S ta in le s s s te e l o u te r
N o z z le
W a te r in le t
H e a te r th e rm o c o u p le
S in te re d s ta in le s s filte r
P re s s u re ta p p in g
S iC H e a t tra n s fe r m e d iu m
S in te re d s ta in le s s
w a te r d is trib u tio n rin g
P o w e r in p u t
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
12
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
13
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!
14
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.
NH3 MPD, F=23 mN, Isp= 600 sec, P=430 W
Stuttgart, Isp=5000sec, F=100N, P=6 MW, hydrogen
15
Pulsed Plasma Thrusters
Outer
Electrode
Rtrigger
Intermediate
Electrode
Center
Electrode
Teflon Annulus
Ctrigger
CMain
Spacecraft
Ground
PPU
Isp = 500-1500 sec
P = 1 – 100 W
Thrust = 5mN/W
16
Hall Effect Thruster
Power = 50W – 25kW
Isp = 500 – 3000 sec
Thrust = 5 mN- 1N
17
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
18

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