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Propulsion Overview

MAE155ADr. Nacouzi

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Agenda

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

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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

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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)

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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

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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...

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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

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Cold Gas Systems• Cold Gas Systems: Involves the expulsion of high pressure

gas

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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

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Bi-propellant Systems• Includes both fuel and oxidizer in separatetanks•Propellant managementthrough pressurant orturbine•Cryogenic systems muchmore complex due to temperature control

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Pump Fed System: Liquid Engine Shematic

Ref: University of Maryland

• High flow rate• Complex, heavy systems

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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

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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

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Different Solid Grain Designs

Chamber pressure, i.e., Thrust, can be tailored for mission

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More Solid Grain Designs Comparison

Progressive: Chamber pressure increases during burnRegressive: Chamber pressure decreases during burnNeutral: Approximately constant chamber pressure

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Solid Propellant Burn Rate

Burning rate, r, is the recession rate of asolid propellant and has units of lengthper 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 propellantcomposition as well as other parameters.

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Hybrid Rockets ConceptHybrid propulsion systems involve the injection of the oxidizerinto a solid fuel. Main purpose is for throttle control. Simpler thana bipropellant system, Isp slightly lower.

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Thrust Vector ControlVarious approaches to thrust vector control,TVC, are shown here

Side Injectionis also calledLITVC (LiquidInjection TVC).Works by producingan asymmetricalnozzle flow, throughan oblique shock,causing a nozzleside force.

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Micropropulsion System (MEMS)

Micro satellite ~ 1kg 15 Microthrusters on a chipimpulse ~ 0.00001 N-sUsed in DACS for Micro Sat

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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

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Motor Thrust

F ~ Thrust Pc~ Chamber PressurePe~ Exit PressurePa~ Ambient PressureAt~ Throat AreaAe~ Nozzle Exit AreaVe~ Gas Exit Velocitydm/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

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Basic Propulsion EquationsFamiliar 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

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Basic Propulsion EquationsPropulsion measure of performance:

Characteristic velocity, C* (C star) is a measure of the energy availablefrom the combustion chamber

C* = Pc At/ (dm/dt)

Combining Cf & C*, we get: Isp = T/ (dm/dt) g = Cf C* / gTherefore, 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 ~ ratio of specific heats

Cf2

2

1

2

1

1

1 1

PePc

1

Pe PaPc

AeAt

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Basic Gas Dynamics EqtsIdeal nozzle performance is based on isentropic relations, calculatedvalues are within a few percent of actual. Further improvementscan be made using correction factors.Temperature as a function of Mach number and (gama) is given by:

T0 T 1 0.5 1 M2 M

Pressure and density relationships are similarly given by: P0 PT0T

1

T0

0 T0T

1

1

T0

Where ( )0 is stagnation or chamber conditions

; Cstar is given by: CstarRTc

2

1

1

2 1( )

2

&

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Nozzle FlowIdeal Nozzle Relationships• Assumes isentropic flow• Nozzle Area Expansion given by = Ae/At

(converging/diverging nozzle) where Ae is the nozzle exit area and At is the nozzle throat area

Chamber, PcM~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)

dmdt

Pc At

RTc

2

1

1

1

0.5

dmdt

Pc At

RTc

2

1

1

1

0.5

Flow rate:

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Nozzle Performance• Exhaust Velocity is given by (Ref: Rocket Propulsion

Elements, G. Sutton):

=> Exhaust Velocity for an ideally expanded nozzle ==> x

~ Ideal Cycle Efficiency

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Nozzle Performance

Area Expansion Ratio for a given exit pressure and gama:

Approximate Eta for Gama =1.2

00.10.20.30.40.50.60.70.8

1 10 100 1000 10000

P0/PeEt

a

Approximate valuefor the ideal cycleefficiency =>(Ref. Sutton)

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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

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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...

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Conclusions

• Example Problem• Questions & Discussions...

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Topocentric Coordinate System (SEZ)