Details About Various Fuels propulsion

1. 1
Propulsion Overview
MAE155A
Dr. Nacouzi

2. 2
Agenda
• Introduction to Propulsion
• Propulsion Systems: Liquids, Solids, other
• Basic Propulsion Performance
• Essential Isentropic Equations
• Nozzle Design and Performance
• Example

3. 3
Propulsion Overview
• Launch & Space Propulsion Systems
– Propulsion system comprise the main
component of a launch system. It can also be a
significant (in terms of wt) component of the
space vehicle (SV), depending on the SV’s
mission
• Delivers SV to proper orbit
• Supports or provides means for interplanetary travel
• Key component of many SV attitude control
systems

4. 4
Types of Propulsion Systems
• Cold Gas: Pressurized Gas Expulsion, low Isp
• Chemical Propulsion Systems: ‘Controlled
Explosions’
– Liquid Systems: High Isp, throttle control,
complex, most liquid propellants are toxic
• Mono-propellant, usually used for SV ACS & Orbit
maintenance, e.g. Hydrazine (long shelf life)
• Bi-propellant, used for SV ACS and orbit
maintenance as well as launch vehicle propulsion,
e.g. MMH-Hydrazine, Cryogenic (LH2/LO2)
• Dual Modes: Bi-propellant systems that can be used
as mono-propellant (to minimize impulse bit)

5. 5
Types of Propulsion Systems
• Solid: Grain includes both fuel and oxidizer bound
together, requires external ignition source (e.g., HTPB).
– Simpler design, easier to handle, long shelf life
– Low detonability, lower Isp than high energy bi-propellants,
include metals to increase Isp, difficult to throttle control.
– Consists of single or multiple pulses (restart option). May include
thrust termination system.
• Hybrid: Liquid oxidizer with solid fuel, throttle control
• Gel Propellant: Safer storage than liquids, easier to
throttle than solids, however viscosity makes flow
management difficult, sensitive to temperature changes

6. 6
Types of Propulsion Systems
• Electric Propulsion: Used for space travel and
orbit maintenance, very high Isp, low thrust and
high energy input requirement
– Electrothermal: Heats propellant using electric power
(solar, nuclear or stored)
– Electrostatic: Ion propulsion, involves ionizing gas and
accelerating it to very high velocity by electrostatic
fields
– Electromagnetic: Plasma is accelerated by electric
current and magnetic field
• Other Propulsion Devices: Solar Sails,
Laser propulsion...

7. 7
Propulsion System Applications
Propulsion
Type
Launch
Vehicle
Orbital
Transfer
Orbital
Maint. &
Maneuvering
ACS Typical Isp
(sec)
Cold Gas X X 25 - 75
Solid X X 260 - 310
Liquid-
Mono
X X 210 - 250
Liquid Bi- X X X X 300 - 400
Dual Mode X X X X 220 - 350
Hybrid X X X 250 - 350
Electric X X 300 - 4000
Note: Electric propulsion effective for interplanetary travel since
high thrust is not typically needed

8. 8
Cold Gas Systems
• Cold Gas Systems: Involves the expulsion of high pressure
gas

9. 9
Liquid Propulsion System
• Mono-propellant systems: Usually hypergolic fuel
(no external spark needed), catalysis used to
initiate chemical reaction
• Bi-propellant systems: Fuel and oxidizer stored
separately, mixed in combustion chamber at pre-
determined mixture ratio and react hypergolically
– Propellant pressurization can be regulated, i.e., external
pressurant, or blowdown, i.e., in propellant tank

10. 10
Bi-propellant Systems
• Includes both fuel
and oxidizer in separate
tanks
•Propellant management
through pressurant or
turbine
•Cryogenic systems much
more complex due to
temperature control

11. 11
Pump Fed System: Liquid
Engine Shematic
Ref: University of Maryland
• High flow rate
• Complex,
heavy systems

12. 12
Solid Propulsion
• Solid propulsion systems have good
performance, are easy to handle, withstand
shock, and less complex than liquid systems
(few moving parts)
• Difficult to check solid stages for internal
cracks (although X ray may be used for
small motors)
– Cracks and failed bonds, can cause catastrophic
termination due to increased burning surface
area

13. 13
Solid Propellant Rocket Motor
Schematic
- Casing holds and protects propellant
- Thermal insulation applied on both outside
and inside the casing to protect from both
external and internal heating

14. 14
Different Solid Grain Designs
Chamber pressure, i.e., Thrust,
can be tailored for mission

15. 15
More Solid Grain Designs Comparison
Progressive: Chamber pressure increases during burn
Regressive: Chamber pressure decreases during burn
Neutral: Approximately constant chamber pressure

16. 16
Solid Propellant Burn Rate
Burning rate, r, is the recession rate of a
solid propellant and has units of length
per time. The burning rate is estimated
from:
r = a Pcn
where,
a~ empirical constant fn of initial grain Temp &
n~ burning rate pressure exponent
Note that r is also a function of the propellant
composition as well as other parameters.

17. 17
Hybrid Rockets Concept
Hybrid propulsion systems involve the injection of the oxidizer
into a solid fuel. Main purpose is for throttle control. Simpler than
a bipropellant system, Isp slightly lower.

18. 18
Thrust Vector Control
Various approaches to thrust vector control,TVC, are shown here
Side Injection
is also called
LITVC (Liquid
Injection TVC).
Works by producing
an asymmetrical
nozzle flow, through
an oblique shock,
causing a nozzle
side force.

19. 19
Micropropulsion System (MEMS)
Micro satellite ~ 1kg 15 Microthrusters on a chip
impulse ~ 0.00001 N-s
Used in DACS for Micro Sat

20. 20
Ideal Rocket Assumptions
• Ideal rocket equations are usually used to estimate
the performance of a rocket. Assumptions:
– Homogeneous & single (gaseous) phase
products
– Perfect gas, adiabatic & isentropic
– Steady state, axial flow with uniform
distribution
– No chemical reactions past chamber, boundary
layer, i.e., friction effects are ignored

21. 21
Motor Thrust
F ~ Thrust
Pc~ Chamber Pressure
Pe~ Exit Pressure
Pa~ Ambient Pressure
At~ Throat Area
Ae~ Nozzle Exit Area
Ve~ Gas Exit Velocity
dm/dt ~ Gas Mass Flow Rate
F = dm/dt x Ve + (Pe - Pa) x Ae,
Veq = Ve + (Pe - Pa) x Ae
dm/dt
Sonic Line
@ Throat
(M = 1)
Exit Mach > 1

22. 22
Basic Propulsion Equations
Familiar specific impulse relationship: Isp = T/[(dm/dt) g]
Total Impulse, I, is given by: I tT
⌠

⌡
d:=
We know, T = Veq x dm/dt = Isp (dm/dt) g
=> Isp = Veq/g
Other definitions for Propulsion Measures of Performance:
Thrust Coefficient, Cf: measure of nozzle performance efficiency
Cf = T/(Pc At) ; Cf ~ fn(nozzle design, chamber conditions)
where,
T ~ Thrust, Pc ~ Chamber Pressure, At ~ Nozzle Throat Area
=> I ~ T x t

23. 23
Basic Propulsion Equations
Propulsion measure of performance:
Characteristic velocity, C* (C star) is a measure of the energy available
from the combustion chamber
C* = Pc At/ (dm/dt)
Combining Cf & C*, we get: Isp = T/ (dm/dt) g = Cf C* / g
Therefore, given Cf and C*, the performance of the rocket can be evaluated.
Cf is given by:
( )e ~ conditions at exit, ( )c in chamber, ( )a is ambient
γ ~ ratio of specific heats
Cf
2 γ
2
⋅
γ 1−
2
γ 1+






γ 1+
γ 1−
⋅ 1
Pe
Pc






γ 1−
γ
−








⋅








Pe Pa−
Pc
Ae
At
⋅+:=

24. 24
Basic Gas Dynamics Eqts
Ideal nozzle performance is based on isentropic relations, calculated
values are within a few percent of actual. Further improvements
can be made using correction factors.
Temperature as a function of Mach number and γ (gama) is given by:
T0 T 1 0.5 γ 1−( ) M
2
+ := M
Pressure and density relationships are similarly given by: P0 P
T0
T






γ
γ 1−
:=
T0
ρ0 ρ
T0
T






1
γ 1−
:=
T0
Where ( )0 is stagnation or
chamber conditions
; Cstar is given by: Cstar
γRTc
γ
2
γ 1+






γ 1+
2 γ 1−( )
⋅
:=
2
&

25. 25
Nozzle Flow
Ideal Nozzle Relationships
• Assumes isentropic flow
• Nozzle Area Expansion given by ε = Ae/At
(converging/diverging nozzle) where Ae is the nozzle exit
area and At is the nozzle throat area
Chamber, Pc
M~0
Nozzle Exit, Ae (Me>1)
Nozzle Throat, At (M=1)
- Flow is choked @ throat, M=1
- Pa (ambient) is < Pt (Pt is
the throat pressure)
dm
dt
Pc At⋅
γ
RTc






2
γ 1+






γ 1+
γ 1−






⋅








0.5
⋅:=
dm
dt
Pc At⋅
γ
RTc






2
γ 1+






γ 1+
γ 1−






⋅








0.5
⋅:=
Flow rate:

26. 26
Nozzle Performance
• Exhaust Velocity is given by (Ref: Rocket Propulsion
Elements, G. Sutton):
=> Exhaust Velocity for
an ideally expanded nozzle ==> x√η
η ~ Ideal Cycle Efficiency

27. 27
Nozzle Performance
Area Expansion Ratio for a given exit pressure and gama:
Approximate Eta for Gama =1.2
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
1 10 100 1000 10000
P0/Pe
Eta
Approximate value
for the ideal cycle
efficiency =>
(Ref. Sutton)

28. 28
Nozzle Performance Considerations
• An ideally expanded nozzle has its exit pressure
equal to the operating ambient pressure (1) =>
A nozzle operating in a vacuum would have an
infinite expansion ratio…
– Overexpanded nozzle (2): Pe < Pa
• Oblique shock waves outside of exit plane
• For higher Pa, flow separation & oblique
shock waves are formed inside the nozzle
– Underexpanded nozzle (3): Pe > Pa
• Expansion waves @ exit plane to equalize
pressure

29. 29
Nozzle Performance Considerations
• Fixed geometry nozzles cannot be designed to be
optimal through their whole flight regime. Nozzle
is underexpanded at ignition and overexpanded at
burnout. Must be optimized for best overall
performance.
• When testing at sea level, nozzles are usually
overexpanded (especially for upper stages).
Adjustments to results and/or test article must be
performed to account for ambient pressure
differences...

30. 30
Conclusions
• Example Problem
• Questions & Discussions...

31. 31
Topocentric Coordinate System (SEZ)