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1 Page 1 School of Aeronautics and Astronautics Rocket Propulsion Fundamentals & Nuclear Thermal Propulsion Professor Steve Heister School of Aeronautics and Astronautics Outline Rocket Performance Fundamentals How much thrust do we get? Rocket Design Fundamentals How much propellant is required? Classification of Rocket Propulsion Systems Historical Accomplishments - the NERVA/Rover Program Nuclear Thermal Space Propulsion Design Studies Conclusions

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Page 1: Rocket Propulsion Fundamentals Nuclear Thermal … Propulsion Fundamentals & Nuclear Thermal Propulsion ... aerospace propulsion system ... propulsion)

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School of Aeronautics and Astronautics

Rocket Propulsion Fundamentals

& Nuclear Thermal Propulsion

Professor Steve Heister

School of Aeronautics and Astronautics

Outline

• Rocket Performance Fundamentals– How much thrust do we get?

• Rocket Design Fundamentals– How much propellant is required?

• Classification of Rocket Propulsion Systems• Historical Accomplishments - the

NERVA/Rover Program• Nuclear Thermal Space Propulsion Design

Studies• Conclusions

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School of Aeronautics and Astronautics

Classification of Rocket Propulsion Systems

School of Aeronautics and Astronautics

Rocket Propulsion in Space Systems(an abundance of uses)

• Launch Vehicles– Solid Rocket Motors (SRMs) and Liquid

Rocket Engines (LREs)• Upper Stage or Orbital Transfer Vehicles

– Solid or liquid propulsion– Nuclear thermal rockets

• Satellite Propulsion– Liquid or electric propulsion options– Nuclear electric rockets

• Spin/Despin Systems, Deorbit Systems– Generally solid or liquid propulsion

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School of Aeronautics and Astronautics

Russian TOPAZ Radio-isotope Thermoelectric Generator (RTG)

School of Aeronautics and Astronautics

There are two important parameters that pertain to any aerospace propulsion system What kind of “gas mileage” does system provide? How much does it weight?

The Specific Impulse, Isp, is the measure of the gas mileage of a rocket propulsion system

where I = total impulse = (N-S or lbf-s )Mp = total propellant mass/weight (lbf, kg)F = delivered thrust (lbf, N)

= propellant flowrate (lbf/s, N/S)

Note: If F, = const then I = F tb

mF

MIIspp �

========

����bt

0dtF

m�

m� bp t/Mm ====�

Rocket Propulsion Basics

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School of Aeronautics and Astronautics

The propellant mass fraction, measures the weight /structural efficiency of the system

“Useful” propellant (that which is burned to provideacceleration in desired direction.)

Sum of all inert masses associated with the propulsionsystem. Includes engines, tanks, pumps, lines, reactors, pressurant bottles, gas generators, insulation, etc.

The best possible design in this sense would be a consumablerocket made entirely from propellant

In general, increases with system size due to structuralefficiency.

,λλλλ

ip

pMM

M++++

====λλλλ

====pM

====iM

(((( ))))1====λλλλ

λλλλ

School of Aeronautics and Astronautics

The best system would have both high Isp and high λλλλ.

Unfortunately, these items are somewhat mutually exclusive

Hot propellants give good Isp but require additional insulation (lowers λλλλ).

Large nozzle gives good Isp but reduce λλλλ . High pressures give better Isp but require thicker-walled

structures. Our consumable rocket would give high λλλλ but lousy Isp.

Page 5: Rocket Propulsion Fundamentals Nuclear Thermal … Propulsion Fundamentals & Nuclear Thermal Propulsion ... aerospace propulsion system ... propulsion)

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School of Aeronautics and Astronautics

The Rocket (or Tsiolkovsky) Equation

m(t)

v(t)

F(t)

D = Drag (atmospheric)

m g = Weight

Newton’s 2nd Law dtdvmamgmDF ========−−−−−−−−

But and sogIspmF �==== dt/dmm −−−−====�

dvgdtdtmD

mdmgIsp ====−−−−−−−−−−−−

Initial Conditions:

Final Conditions:

Now integrate differential eq. To give

omm,0v,0t ============

fb mm,vv,tt ====∆∆∆∆========

����

School of Aeronautics and Astronautics

The Rocket/Tsiolkovsky Equation

(((( ))))���

�������� ���� �����

loss"tg"orgravity

tg

lossdrag

dtbt

0 mD

systempropulsionthebyimpartedgainvelocity

m/mlnIspg

payloadthesensedgain

velocityactual

v bfo

−−−−

−−−−−−−−����−−−−====∆∆∆∆

This equation is the fundamental expression used inRocket Design.

It links mission requirements , propulsion performance, and vehicle masses .

)V(∆∆∆∆)Isp( (((( ))))fo m/m

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•••• Advantages •••• Disadvantages Simple, high λλλλ Cannot be throttled Devices scale up to or turned off

high thrust in All propellant lies instraightforward the combustionmanner chamber

Ready at a moment’s Isp typically lower thannotice for LRE

Package very well Difficult to test

Solid Rocket Motors

School of Aeronautics and Astronautics

Solid Rocket Motor

Graphite Epoxy Motor (GEM) used as strap-on booster forDelta II launch vehicle

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Advantages Disadvantages

— Throttable — More complex, lower

— Can test prior to launch than SRMs

— Higher Isp than solid — Leaking & boiloff issues

rockets — Not as responsive, must

— Can serve as dual-use fill tanks prior to launch

(apogee engine + spacecraft --- Complex pumps often

propulsion) required

Liquid Rocket Engines (LREs)

λλλλ

School of Aeronautics and Astronautics

Bipropellant Liquid Rocket Engine

Space Shuttle Main Engine (SSME)

- utilizes Liquid Oxygen (LOX) andLiquid Hydrogen propellants

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School of Aeronautics and Astronautics

Advantages Disadvantages

— Higher Isp than LREs — Lower than LREs

— Eliminates handling of (requires power conditioning

of hazardous liquids unit)

— Enables precise impulse — Requires high power input

control — Possible spacecraft charging

problem

--- Very low thrust levels

Electric Propulsion (EP)

λλλλ

School of Aeronautics and Astronautics

Electric Propulsion Ion Thruster

Xenon Ion PropulsionSystem (XIPS) thrusterused on Hughes Satellites

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Advantages

• High thrust capability

• High Isp, roughly double that attainable using chemical propulsion

• Uses only a single fluid – simplifies plumbing/pumps

Disadvantages

• It’s a nuke – cultural perceptions influence acceptability

• Testing is difficult – exhaust scrubbers required, disposal of used

hardware an issue

• Substantial inert mass required for shielding (Lowers )

• Engine is difficult to shut down – long tailoff

Nuclear Thermal Propulsion

λλλλ

School of Aeronautics and Astronautics

General Classification of Rocket Propulsion Systems

Engine Thrust/Wght. Isp Range,SEC

ThrustDuration

λλλλ

SRM 1-10 250-310 1-200 sec Highest

LRE 1-10 300-475 1-1000 sec Moderate

Nuke Th. 1-5 750-1000 hours Poor

Electric <0.001 300-10,000 months Poor

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Brief History

Rover/NERVA Program (1955-1973)

• NERVA = Nuclear Engine for Rocket Vehicle Application

• 20 Reactors Designed, Built & Tested from 1955-1973

- Total Cost $1.4 B

- Reactors Kiwi, Phoebus, NRX, Pewee

• 1100 - 4100 MW Reactors

• 55,000 – 210,000 (Phoebus) lbf Thrust Levels

• 2550 – 2700 K Fuel Temps

• Isp to 850 sec (Pewee)

• “Burn” Durations 1 – 2 Hours

Space Nuclear Thermal Propulsion (SNTP)

• Late 1987-1993

• 1000 MW core, T = 3000 K, Isp = 1000 sec, PwR Density = 40 Mw/l

School of Aeronautics and Astronautics

Fission Process

• This Rxn Generates 260 Mev of Energy

• Neutrons Control Rxn

• Gamma Rays are a Pain!

•1-2% Fission Fragments Enter Fuel

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Cross-Section of NERVA Reactor

Control Drums expose neutron absorbing material such as B, Ha to slow rxn

Reflector fabricated from Be or Li

School of Aeronautics and Astronautics

NERVA 50,000 lbfXE Flight Engine Installed in test stand

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NERVA Reactor Fuel Element Design

School of Aeronautics and Astronautics

Fuel Bead Design

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Phoebus Reactor in Transport

School of Aeronautics and Astronautics

Design Study Conducted by Grumman Aerospace

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Engine Cycle Comparisons

School of Aeronautics and Astronautics

Engine Thrust/Weight Vs. Power Level

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Vehicle Performance Comparisons

School of Aeronautics and Astronautics

Vehicle Mass Comparisons

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School of Aeronautics and Astronautics

Impulse Loss During Shutdown

School of Aeronautics and Astronautics

Another cool picture of NERVA

I believe this is a mock-up of a flight unit

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Conclusions

• Nuclear Thermal Propulsion Provides Attractive Performance for Space Missions

• Current Political Environment is not Conducive to Development of Engines of this Type

•Technical Challenges/Requirements Include

• High Temperature Materials (Current SOA about 3000 K)

• High Core Power Densities (Current SOA about 40 MW/liter)

• Low Weight Shielding

• Potential Flow Instabilities - Pressure drop coupling to heat transfer from core

School of Aeronautics and Astronautics

Vehicle Performance Specs