6. Chemical-Nuclear Propulsion MAE 342 2016Solid-Fuel Rocket Motor Thrust is proportional to burning...
Transcript of 6. Chemical-Nuclear Propulsion MAE 342 2016Solid-Fuel Rocket Motor Thrust is proportional to burning...
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Chemical/Nuclear PropulsionSpace System Design, MAE 342, Princeton University
Robert Stengel
• Thermal rockets• Performance parameters• Propellants and propellant
storage
Copyright 2016 by Robert Stengel. All rights reserved. For educational use only.http://www.princeton.edu/~stengel/MAE342.html
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Chemical (Thermal) Rockets
• Liquid/Gas Propellant–Monopropellant
• Cold gas• Catalytic decomposition
–Bipropellant• Separate oxidizer and fuel• Hypergolic (spontaneous)
ignition• External ignition• Storage
– Ambient temperature and pressure
– Cryogenic– Pressurized tank
–Throttlable–Start/stop cycling
• Solid Propellant–Mixed oxidizer and fuel–External ignition–Burn to completion
• Hybrid Propellant–Liquid oxidizer, solid fuel–Throttlable–Start/stop cycling
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Cold Gas Thruster(used with inert gas)
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Moog Divert/Attitude Thruster and Valve
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Monopropellant Hydrazine Thruster
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• Catalytic decomposition produces thrust
• Reliable• Low performance• Toxic
Aerojet Rocketdyne
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Bi-Propellant Rocket MotorThrust / Motor Weight ~ 70:1
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Hypergolic, Storable Liquid-Propellant Thruster
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• Spontaneous combustion
• Reliable• Corrosive, toxic
Titan 2
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Pressure-Fed and TurbopumpEngine Cycles
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Gas-Generator Rocket Cycle, with Nozzle Cooling
Pressure-Fed Rocket Cycle
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Staged Combustion Engine Cycles
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Staged Combustion Rocket Cycle
Full-Flow Staged Combustion Rocket Cycle
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German V-2 Rocket Motor, Fuel Injectors, and Turbopump
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Combustion Chamber Injectors
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• Air Force legacy (1955)• Design undertaken before
vehicle or mission were identified
• Big engine, big problems• 16:1 nozzle expansion• 6.67 MN thrust
• F-1 turbopumps• Oxygen: 24,811 gal/min• RP-1: 15,741 gal/min
• F-1 injector• Combustion instability
• Significant theoretical work by Luigi Crocco and David Harrje, Princeton 13
Origins of the F-1
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USSR RD-107/8 Rocket MotorsRD-107
4 combustion chambers, 2 verniersRD-108
4 combustion chambers, 4 verniers
R-7 Base4-RD-107, 1-RD-108
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(used on Atlas V)
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Special Shuttle Main Engine (RS-25)
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SpaceX Merlin Family
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Merlin 1C (vacuum nozzle)
Merlin 1D (throttlable)
Merlin 1A (ablative nozzle)
Roll control from turbine exhaust
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• LOX/Liquefied natural gas
• United Launch Alliance has chosen as motor for the Vulcan launch vehicle
• Thrust = 2.5 MN (550,000 lb)
Blue Origin BE-4
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RD-181 and RD-191RD-181 RD-191
to be used on Orbital-ATK Antares
to be used on NPO Energomash Angara
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Solid-Fuel Rocket Motor
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Solid-Fuel Rocket Motor
Thrust is proportional to burning areaRocket grain patterns affect thrust profile
Propellant chamber must sustain high pressure and temperatureEnvironmentally unfriendly exhaust gas 23
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Hybrid-Fuel Rocket Motor
• SpaceShipOne motor– Nitrous oxide– Hydroxy-terminated polybutadiene (HTPB)
• Issues– Hard start– Blow back– Complete mixing of oxidizer and fuel toward completion of burn
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Rocket Thrust
Thrust = !mpropellantVexhaust + Aexit pexit − pambient( ) ≡ !m ceff
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ceff =Thrust!m
= Effective exhaust velocity
!m ≡Mass flow rate of on-board propellant
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Specific Impulse
• go is a normalizing factor for the definition• Chemical rocket specific impulse (vacuum)
– Solid propellants: < 295 s– Liquid propellants: < 510 s
Isp =Thrust!m go
=ceffgo
, Units = m / sm / s2 = seconds
go ≡ Gravitational acceleration at earth's surface
• Space Shuttle Specific Impulses–Solid boosters: 242-269 s–Main engines: 455 s–OMS: 313 s–RCS: 260-280 s
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Specific Impulse
Isp =Thrust!m go
=ceffgo
= CF c* go
= Vexhaustgo
when CF = 1, pe = pambient
Specific impulse is a product of characteristic velocity, c*, and rocket
thrust coefficient, CF
• Characteristic velocity is related to– combustion chamber performance– propellant characteristics
• Thrust coefficient is related to– nozzle shape– exit/ambient pressure differential
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The Rocket EquationIdeal velocity increment of a rocket stage, ΔVI (gravity and aerodynamic effects
neglected)
dVdt
= Thrustm
=!m ceffm
= −dm
dt Ispgom
Konstantin Tsiolkovsky
dVVi
Vf
∫ = −Ispgo dmm
mi
mf
∫ = −Ispgo lnm mi
mf
Vf −Vi( ) ≡ ΔVI = Ispgo lnmi
mf
⎛
⎝⎜⎞
⎠⎟≡ Ispgo ln µ
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Volumetric Specific Impulse
ΔVI = Ispgo lnµ = Ispgo lnmfinal +mpropellant
m final
⎛
⎝⎜⎞
⎠⎟= Ispgo ln 1+
mpropellant
m final
⎛
⎝⎜⎞
⎠⎟
= Ispgo ln 1+Densitypropellant •Volumepropellant
m final
⎛
⎝⎜⎞
⎠⎟
≈ goIspρ propellant •Volpropellant
m final
⎛
⎝⎜⎞
⎠⎟= go Ispρ propellant( )Volpropellantm final
Specific impulse
Volumetric specific impulse
Ispvol !VIsp = Ispρ propellant29
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Volumetric Specific Impulse
• For fixed volume and final mass, increasing volumetric specific impulse increases ideal velocity increment
•Saturn V Specific Impulses, vacuum (sea level)–1st Stage, 5 F-1 LOX-Kerosene Engines: 304 s (265 s)–2nd Stage, 5 J-2 LOX-LH2 Engines: 424 s (~360 s)–3rd Stage, 1 J-2 LOX-LH2 Engine: 424 s (~360 s) 30
Density, g/cc Isp, s, SL
VIsp, s (g/cc), SL Isp, s, vac
VIsp, s (g/cc), vac
LOX/Kerosene 1.3 265 345 304 395
LOX/LH2 (Saturn V) 0.28 360 101 424 119LOX/LH2 (Shuttle) 0.28 390 109 455 127
Shuttle Solid Booster 1.35 242 327 262 354
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Typical Values of Chemical Rocket Specific Impulse • Chamber pressure = 7 MPa (low by
modern standards)• Expansion to exit pressure = 0.1 MPa
Solid-Propellant Rockets
Double-Base Isp, sVIsp, kg-s/m^3 x 10^3
AFU 196 297ATN 235 376JPN 250 405
CompositeJPL 540A 231 383TRX-H609 245 431PBAN (SSV) 260 461
Liquid-Fuel Rockets
Monopropellant Isp, sVIsp, kg-s/m^3 x 10^3
Hydrogen Peroxide 165 238Hydrazine 199 201Nitromethane 255 290
Bipropellant
Fuel Oxidizer Isp, sVIsp, kg-s/m^3 x 10^3
Kerosene Oxygen 301 307Flourine 320 394Red Fuming Nitric Acid 268 369
Hydrogen Oxygen 390 109Flourine 410 189
UDMHNitrogen Tetroxide 286 339
Hybrid-Fuel RocketFuel Oxidizer Isp, sHTPB N2O 250
SSME
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Exhaust Velocity vs. Thrust Acceleration
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Rocket Characteristic Velocity, c*
c*= 1Γ
RoTcM
, where Γ = γ 2γ +1
⎛⎝⎜
⎞⎠⎟
γ +12 γ −1( )
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Ro = universal gas constant = 8.3×103 kg m2 s2 °K
Tc = chamber temperature, °KM = exhaust gas mean molecular weightγ = ratio of specific heats ~1.2-1.4( )
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Rocket Characteristic Velocity, c*
c*= pcAt
!m= exhaust velocity if CF = 1
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Rocket Thrust Coefficient, CF
CF typically 0.5 - 2
CF =ThrustpcAt
= λΓ 2γγ −1
⎛⎝⎜
⎞⎠⎟1− pe
pc
⎛⎝⎜
⎞⎠⎟
γ −1( ) γ⎡
⎣⎢⎢
⎤
⎦⎥⎥+ pe − pambient
pc
⎛⎝⎜
⎞⎠⎟AeAt
Thrust = λ !m ve + Ae pe − pambient( )λ : reduction ratio (function of nozzle shape)
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Thrust Coefficient, CF, vs. Nozzle Expansion Ratio
• xx
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Mixture Ratio, r
• Stoichiometric mixture: complete chemical reaction of propellants
• Specific impulse maximizedwith lean mixture ratio, r (i.e., below stoichiometric maximum)
r =!moxidizer
!mfuel
; !mfuel =!mtotal
1+ r; "leaner"< r < "richer"
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Effect of Pressure Ratio on Mass Flow
!m = Γ pcAtR oTc
M
In choked flow, mass flow rate is maximized
pepc
≤ 2γ +1
⎛⎝⎜
⎞⎠⎟
γ γ −1
≈ 0.53
Choked flow occurs when
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Combustion Instability
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Harrje, NASA SP-194, 1972
• Complex mix of species, phases, pressures, temperatures, and flows
• Cavity resonance
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Combustion Instability
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Stable Response to Disturbance
Unstable Response to Disturbance
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Shock DiamondsWhen pe ≠pa, exhaust flow is over- or underexpanded
Effective exhaust velocity < maximum value
Vikinghttps://www.youtube.com/watch?v=qiMSko4HBe8
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Rocket Nozzles
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Rocket Nozzles• Expansion ratio, Ae/At,
chosen to match exhaust pressure to average ambient pressure– Ariane rockets: Viking V for
sea level, Viking IV for high altitude
• Rocket nozzle types– DeLaval nozzle– Isentropic expansion nozzle– Spike/plug nozzles– Expansion-deflection nozzle
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Rocket Nozzles
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Linear Spike/Plug Nozzles
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Throttling, Start/Stop Cycling
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CECE Demonstrator PintleInjector
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Reaction Control Thrusters• Direct control of angular rate• Unloading momentum wheels or control-moment gyros• Reaction control thrusters are typically on-off devices using
– Cold gas– Hypergolic propellants– Catalytic propellant– Ion/plasma rockets
• Thrusters commanded in pairs to cancel velocity change
Apollo Lunar Module RCS Space Shuttle RCS
• Issues– Specific impulse– Propellant mass– Expendability
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RCS Thruster
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Divert and Attitude Control Thrusters
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https://www.youtube.com/watch?v=W8efpDBvTDE
https://www.youtube.com/watch?v=KBMU6l6GsdM
https://www.youtube.com/watch?v=JURQYH669_g
https://www.youtube.com/watch?v=71qgI6bddM8
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Nuclear Propulsion
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• Nuclear reaction produces thermal energy to heat inert working fluid– Solid core– Liquid core– Gaseous core
• High propellant temperature leads to high specific impulse
• Working fluid chosen for low molecular weight and storability
c*= 1Γ
RoTcM
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Solid-Core Nuclear Rocket
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• Operating temperature limited by – melting point of reactor materials– cracking of core coating– matching coefficients of expansion
• Possible propellants: hydrogen, helium, liquid oxygen, water, ammonia
• Isp = 850 – 1,000 sec• T / W ~ 7:1
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Project Rover, 1955-1972
Kiwi-B4-A Reactor/Rocket
NERVA Rocket, Isp ~ 900 sec
NERVA-Powered Mars Mission
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Liquid/Particle-Core Nuclear Rocket
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• Nuclear fuel mixed with working fluid• In principle, could operate above
melting point of nuclear fuel• Isp ~ 1,300 – 1,500 sec• Conceptual• Massive radioactive waste
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Open-Cycle Gas Core Nuclear Rocket
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• Toroidal circulation of working fluid confines nuclear fuel to center
• Fuel does not touch the wall• Conceptual• Massive radioactive waste• Isp ~ 3,000 – 5,000 sec
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Closed-Cycle Gas Core Nuclear Rocket
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• “Nuclear light bulb”• Nuclear fuel contained in quartz container• Isp ~ 1,500 – 2,000 sec• Conceptual
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Nuclear-Pulse (“Explosion”) Rocket - Project Orion
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“Physics packages” ejected behind the pusher plate
https://en.wikipedia.org/wiki/Project_Orion_(nuclear_propulsion)
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Next Time:Launch Vehicles
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Supplemental Material
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Propellant Tanks
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Propellant must be kept near the exit duct without bubbles during thrusting
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Ion/Plasma Thrusters
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Engine Propellant Required power Specific impulse ThrustkW s mN
NSTAR Xenon 2.3 3,300 to 1,700 92 maxNEXT[ Xenon 6.9 4,300 236 maxHiPEP Xenon 20–50 6,000–9,000 460–670
Hall effect Xenon 25 3,250 950FEEP Liquid Cesium 6×10−5–0.06 6,000–10,000 0.001–1
VASIMR Argon 200 3,000–12,000 ~5,000DS4G Xenon 250 19,300 2,500 max
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Variable Specific Impulse Magnetoplasma Rocket (VASIMR)
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PropellantRequired
powerSpecific impulse Thrust
kW s mNArgon 200 3,000–12,000 ~5,000
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DAWN Spacecraft
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Engine Propellant Required power Specific impulse ThrustkW s mN
NSTAR Xenon 2.3 3,300 to 1,700 92 max
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