rocket propulsion fundamentals nuclear thermal … propulsion fundamentals & nuclear thermal...
TRANSCRIPT
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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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Classification of Rocket Propulsion Systems
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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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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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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.
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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
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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
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Bipropellant Liquid Rocket Engine
Space Shuttle Main Engine (SSME)
- utilizes Liquid Oxygen (LOX) andLiquid Hydrogen propellants
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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
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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
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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
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NERVA 50,000 lbfXE Flight Engine Installed in test stand
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NERVA Reactor Fuel Element Design
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Fuel Bead Design
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Phoebus Reactor in Transport
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Design Study Conducted by Grumman Aerospace
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Engine Cycle Comparisons
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Engine Thrust/Weight Vs. Power Level
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Vehicle Performance Comparisons
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Vehicle Mass Comparisons
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Impulse Loss During Shutdown
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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
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Vehicle Performance Specs