Chemical/Nuclear Propulsion!

Chemical/Nuclear Propulsion! Space System Design, MAE 342, Princeton University! Robert Stengel Thermal rockets Performance parameters Propellants and propellant storage Copyright 2016 by Robert Stengel.
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Chemical/Nuclear Propulsion! Space 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. 1 Chemical (Thermal) Rockets Liquid/Gas Propellant!Monopropellant Catalytic ignition / chemical decomposition Cold gas!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 2 Cold Gas Thruster! (used with inert gas) Moog Divert/Attitude Thruster and Valve 3 Monopropellant Hydrazine Thruster Aerojet Rocketdyne Catalytic decomposition produces thrust Reliable Low performance Toxic 4 Bi-Propellant Rocket Motor Thrust / Motor Weight ~ 70:1 5 Hypergolic, Storable Liquid- Propellant Thruster Titan 2 Spontaneous combustion Reliable Corrosive, toxic 6 Pressure-Fed and Turbopump Engine Cycles Pressure-Fed Rocket Cycle Gas-Generator Rocket Cycle, with Nozzle Cooling 7 Staged Combustion Engine Cycles Staged Combustion Rocket Cycle Full-Flow Staged Combustion Rocket Cycle 8 German V-2 Rocket Motor, Fuel Injectors, and Turbopump 9 Combustion Chamber Injectors 10 12 11 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 Origins of the F-1 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 14 USSR RD-107/8 Rocket Motors RD combustion chambers, 2 verniers RD combustion chambers, 4 verniers R-7 Base 4-RD-107, 1-RD (used on Atlas V) 16 Special Shuttle Main Engine (RS-25) 17 Merlin 1A (ablative nozzle) SpaceX Merlin Family Merlin 1C (vacuum nozzle) Merlin 1D (throttlable) Roll control from turbine exhaust 18 Blue Origin BE-4 LOX/Liquefied natural gas United Launch Alliance has chosen as motor for the Vulcan launch vehicle Thrust = 2.5 MN (550,000 lb) 19 RD-181 and RD-191 RD-181 RD-191 to be used on Orbital- ATK Antares to be used on NPO Energomash Angara 20 Solid-Fuel Rocket Motor 21 22 Solid-Fuel Rocket Motor Thrust is proportional to burning area Rocket grain patterns affect thrust profile Propellant chamber must sustain high pressure and temperature Environmentally unfriendly exhaust gas 23 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 24 Rocket Thrust Thrust =!m propellant V exhaust + A exit ( p exit! p ambient ) !m c eff c eff = Thrust = Effective exhaust velocity!m!m! Mass flow rate of on-board propellant 25 I sp = Thrust!m g o = c eff g o, Specific Impulse Units = m / s m / s 2 = seconds g o! Gravitational acceleration at earth's surface g o is a normalizing factor for the definition Chemical rocket specific impulse (vacuum)! Solid propellants: 295 s! Liquid propellants: 510 s Space Shuttle Specific Impulses!Solid boosters: s!main engines: 455 s!oms: 313 s!rcs: s 26 Specific Impulse Specific impulse is a product of characteristic velocity, c*, and rocket thrust coefficient, C F I sp = Thrust!m g o = c eff g o = C F c * g o = V exhaust g o when C F = 1, p e = p ambient Characteristic velocity is related to! combustion chamber performance! propellant characteristics Thrust coefficient is related to! nozzle shape! exit/ambient pressure differential 27 Konstantin Tsiolkovsky The Rocket Equation Ideal velocity increment of a rocket stage, V I (gravity and aerodynamic effects neglected) dv dt = Thrust m =!m c eff m =! dm dt I spg o m V f m! dv = I sp g dm o! m = I sp g o ln m f mi V i m f m i $ ( V f! V i ) #V I = I sp g o ln m i & % m f ' ) ( I spg o ln µ 28 Volumetric Specific Impulse Specific impulse !V I = I sp g o ln µ = I sp g o ln m final + m propellant % $ ' # & = I g ln 1+ m % propellant sp o $ ' # & m final = I sp g o ln 1+ Density Volume % propellant propellant $ ' # & ) ( g o I propellant Vol propellant % sp $ ' # & = g o m final m final m final ( I sp ) propellant ) Vol propellant m final Volumetric specific impulse I spvol! VI sp = I sp! propellant 29 Volumetric Specific Impulse For fixed volume and final mass, increasing volumetric specific impulse increases ideal velocity increment Density, g/ cc Isp, s, SL VIsp, s (g/cc), SL VIsp, s (g/cc), Isp, s, vac vac LOX/Kerosene LOX/LH2 (Saturn V) LOX/LH2 (Shuttle) Shuttle Solid Booster Saturn V Specific Impulses, vacuum (sea level)!1 st Stage, 5 F-1 LOX-Kerosene Engines: 304 s (265 s)!2 nd Stage, 5 J-2 LOX-LH2 Engines: 424 s (~360 s)!3 rd Stage, 1 J-2 LOX-LH2 Engine: 424 s (~360 s) 30 Typical Values of Chemical Rocket Specific Impulse Chamber pressure = 7 MPa (low by modern standards) Expansion to exit pressure = 0.1 MPa SSME Liquid-Fuel Rockets Monopropellant Isp, s VIsp, kgs/m^3 x 10^3 Hydrogen Peroxide Hydrazine Nitromethane Bipropellant Fuel Oxidizer Isp, s VIsp, kgs/m^3 x 10^3 Kerosene Oxygen Flourine Red Fuming Nitric Acid Hydrogen Oxygen Flourine UDMH Nitrogen Tetroxide Solid-Propellant Rockets Double-Base Isp, s VIsp, kgs/m^3 x 10^3 AFU ATN JPN Composite JPL 540A TRX-H PBAN (SSV) Hybrid-Fuel Rocket Fuel Oxidizer Isp, s HTPB N2O Exhaust Velocity vs. Thrust Acceleration 32 Rocket Characteristic Velocity, c* c* = 1! R o T c M, where! = # 2 & $ % +1' ( +1 2 )1 ( ) R o = universal gas constant = 8.3!10 3 kg m 2 s 2 K T c = chamber temperature, K M = exhaust gas mean molecular weight ( ) = ratio of specific heats ~ Rocket Characteristic Velocity, c* c* = p c A t!m = exhaust velocity if C F = 1 34 Rocket Thrust Coefficient, C F C F = Thrust p c A t =! % 2# ( + & ' # $1) * 1$ % p ( e - & ' p c ) *,- (# $1) #. 0 / 0 + % p $ p ( e ambient & ' p c ) * ( ) Thrust =!!m v e + A e p e p ambient! : reduction ratio (function of nozzle shape) A e A t C F typically Thrust Coefficient, C F, vs. Nozzle Expansion Ratio xx 36 r =!m oxidizer!m fuel ;!m fuel =!m total 1+ r Mixture Ratio, r Stoichiometric mixture: complete chemical reaction of propellants Specific impulse maximized with lean mixture ratio, r (i.e., below stoichiometric maximum) ; leaner r richer 37 Effect of Pressure Ratio on Mass Flow In choked flow, mass flow rate is maximized!m =! p ca t R o T c M p e p c! Choked flow occurs when # 2 & % ( $ +1' )1 * Combustion Instability Complex mix of species, phases, pressures, temperatures, and flows Cavity resonance Harrje, NASA SP-194, Combustion Instability Stable Response to Disturbance Unstable Response to Disturbance Harrje, NASA SP-194, Shock Diamonds When p e #p a, exhaust flow is over- or underexpanded Effective exhaust velocity maximum value https://www.youtube.com/watch?v=qimsko4hbe8 Viking! 41 Rocket Nozzles 42 Rocket Nozzles Expansion ratio, A e /A t, 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 43 Rocket Nozzles 44 Linear Spike/Plug Nozzles 45 Throttling, Start/Stop Cycling CECE Demonstrator Pintle Injector 46 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 Issues! Hypergolic propellants! Specific impulse! Catalytic propellant! Propellant mass! Ion/plasma rockets! Expendability Thrusters commanded in pairs to cancel velocity change Apollo Lunar Module RCS Space Shuttle RCS RCS Thruster 47 Divert and Attitude Control Thrusters https://www.youtube.com/watch?v=w8efpdbvtde https://www.youtube.com/watch?v=71qgi6bddm8 https://www.youtube.com/watch?v=kbmu6l6gsdm https://www.youtube.com/watch?v=jurqyh669_g 48 Nuclear Propulsion c* = 1! R o T c M 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 49 Solid-Core Nuclear Rocket 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 I sp = 850 1,000 sec T / W ~ 7:1 50 Project Rover, NERVA Rocket, I sp ~ 900 sec NERVA-Powered Mars Mission Kiwi-B4-A Reactor/Rocket 51 52 Liquid/Particle-Core Nuclear Rocket Nuclear fuel mixed with working fluid In principle, could operate above melting point of nuclear fuel I sp ~ 1,300 1,500 sec Conceptual Massive radioactive waste 53 Open-Cycle Gas Core Nuclear Rocket Toroidal circulation of working fluid confines nuclear fuel to center Fuel does not touch the wall Conceptual Massive radioactive waste I sp ~ 3,000 5,000 sec 54 Closed-Cycle Gas Core Nuclear Rocket Nuclear light bulb Nuclear fuel contained in quartz container I sp ~ 1,500 2,000 sec Conceptual 55 Nuclear-Pulse ( Explosion ) Rocket - Project Orion Physics packages ejected behind the pusher plate https://en.wikipedia.org/wiki/project_orion_(nuclear_propulsion) 56 Next Time:! Launch Vehicles! 57 Supplemental Material! 58 Propellant Tanks Propellant must be kept near the exit duct without bubbles during thrusting 59 Ion/Plasma Thrusters Engine Propellant Required power Specific impulse Thrust kw s mn NSTAR Xenon 2.3 3,300 to 1, max NEXT[ Xenon 6.9 4, max HiPEP Xenon ,000 9, Hall effect Xenon 25 3, FEEP Liquid Cesium 6 10# ,000 10, VASIMR Argon 200 3,000 12,000 ~5,000 DS4G Xenon ,300 2,500 max 60 Variable Specific Impulse Magnetoplasma Rocket (VASIMR) Propellant Required power Specific impulse Thrust kw s mn Argon 200 3,000 12,000 ~5, DAWN Spacecraft Engine Propellant Required power Specific impulse Thrust kw s mn NSTAR Xenon 2.3 3,300 to 1, max 62
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