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2/12/20 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. http://www.princeton.edu/~stengel/MAE342.html 1 1 Chemical (Thermal) Rockets • Liquid/Gas Propellant –Monopropellant • Cold gas • Catalytic decomposition –Bipropellant • Separate oxidizer and fuel • Hypergolic (spontaneous) • Solid Propellant ignition –Mixed oxidizer and fuel • External ignition –External ignition • Storage –Burn to completion – Ambient temperature and pressure • Hybrid Propellant – Cryogenic –Liquid oxidizer, solid fuel – Pressurized tank –Throttlable –Throttlable –Start/stop cycling –Start/stop cycling 2 2 1 2/12/20 Cold Gas Thruster (used with inert gas) Moog Divert/Attitude Thruster and Valve 3 3 Monopropellant Hydrazine Thruster Aerojet Rocketdyne • Catalytic decomposition produces thrust • Reliable • Low performance • Toxic 4 4 2 2/12/20 Bi-Propellant Rocket Motor Thrust / Motor Weight ~ 70:1 5 5 Hypergolic, Storable Liquid- Propellant Thruster Titan 2 • Spontaneous combustion • Reliable • Corrosive, toxic 6 6 3 2/12/20 Pressure-Fed and Turbopump Engine Cycles Pressure-Fed Gas-Generator Rocket Rocket Cycle Cycle, with Nozzle Cooling 7 7 Staged Combustion Engine Cycles Staged Combustion Full-Flow Staged Rocket Cycle Combustion Rocket Cycle 8 8 4 2/12/20 German V-2 Rocket Motor, Fuel Injectors, and Turbopump 9 9 Combustion Chamber Injectors 10 10 5 2/12/20 11 11 12 12 6 2/12/20 Origins of the F-1 • 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 13 14 14 7 2/12/20 USSR RD-107/8 Rocket Motors RD-107 RD-108 4 combustion chambers, 2 verniers 4 combustion chambers, 4 verniers R-7 Base 4-RD-107, 1-RD-108 15 15 (used on Atlas V) 16 16 8 2/12/20 Special Shuttle Main Engine (RS-25) 17 17 SpaceX Merlin Family Merlin 1A Merlin 1C Merlin 1D (ablative nozzle) (vacuum nozzle) (throttlable) Roll control from turbine exhaust 18 18 9 2/12/20 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 19 RD-181 and RD-191 RD-181 RD-191 to be used on Orbital- to be used on NPO ATK Antares Energomash Angara 20 20 10 2/12/20 Solid-Fuel Rocket Motor 21 21 22 22 11 2/12/20 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 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 24 12 2/12/20 Rocket Thrust Thrust = m! propellantVexhaust + Aexit ( pexit − pambient ) ≡ m! ceff Thrust c = = Effective exhaust velocity eff m! m! ≡ Mass flow rate of on-board propellant 25 25 Specific Impulse Thrust ceff m / s Isp = = , Units = 2 = seconds m! go go m / s go ≡ Gravitational acceleration at earth's surface • go 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: 242-269 s –Main engines: 455 s –OMS: 313 s –RCS: 260-280 s 26 26 13 2/12/20 Specific Impulse Specific impulse is a product of characteristic velocity, c*, and rocket thrust coefficient, CF Thrust ceff Isp = = m! go go = CF c* go V = exhaust when C = 1, p = p g F e ambient o • Characteristic velocity is related to – combustion chamber performance – propellant characteristics • Thrust coefficient is related to – nozzle shape – exit/ambient pressure differential 27 27 The Rocket Equation Ideal velocity increment of a rocket stage, Konstantin ΔVI (gravity and aerodynamic effects Tsiolkovsky neglected) dm dV Thrust m! c Ispgo = = eff = − dt dt m m m Vf m f dm m f dV = −Ispgo = −Ispgo ln m ∫ ∫ m mi Vi mi ⎛ mi ⎞ (Vf −Vi ) ≡ ΔVI = Ispgo ln⎜ ⎟ ≡ Ispgo ln µ ⎝ m f ⎠ 28 28 14 2/12/20 Volumetric Specific Impulse Specific impulse ⎛ m final + mpropellant ⎞ ⎛ mpropellant ⎞ ΔVI = Ispgo ln µ = Ispgo ln⎜ ⎟ = Ispgo ln⎜1+ ⎟ ⎝ m final ⎠ ⎝ m final ⎠ ⎛ Densitypropellant •Volumepropellant ⎞ = Ispgo ln⎜1+ ⎟ ⎝ m final ⎠ ⎛ ρ propellant •Volpropellant ⎞ Volpropellant ≈ goIsp ⎜ ⎟ = go (Ispρ propellant ) ⎝ m final ⎠ m final Volumetric specific impulse I VI I spvol ! sp = spρ propellant 29 29 Volumetric Specific Impulse • For fixed volume and final mass, increasing volumetric specific impulse increases ideal velocity increment Density, VIsp, s (g/cc), VIsp, s (g/cc), g/cc Isp, s, SL SL Isp, s, vac vac LOX/Kerosene 1.3 265 345 304 395 LOX/LH2 (Saturn V) 0.28 360 101 424 119 LOX/LH2 (Shuttle) 0.28 390 109 455 127 Shuttle Solid Booster 1.35 242 327 262 354 •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 30 15 2/12/20 Typical Values of Chemical SSME Rocket Specific Impulse • Chamber pressure = 7 MPa (low by modern standards) • Expansion to exit pressure = 0.1 MPa Liquid-Fuel Rockets VIsp, kg- Solid-Propellant Rockets VIsp, kg- Monopropellant Isp, s s/m^3 x 10^3 Hydrogen Peroxide 165 238 Double-Base Isp, s s/m^3 x 10^3 Hydrazine 199 201 AFU 196 297 Nitromethane 255 290 ATN 235 376 JPN 250 405 Bipropellant VIsp, kg- Composite Fuel Oxidizer Isp, s s/m^3 x 10^3 JPL 540A 231 383 Kerosene Oxygen 301 307 TRX-H609 245 431 Flourine 320 394 PBAN (SSV) 260 461 Red Fuming Nitric Acid 268 369 Hybrid-Fuel Rocket Hydrogen Oxygen 390 109 Fuel Oxidizer Isp, s Flourine 410 189 Nitrogen HTPB N2O 250 UDMH Tetroxide 286 339 31 31 Exhaust Velocity vs. Thrust Acceleration 32 32 16 2/12/20 Rocket Characteristic Velocity, c* γ +1 1 R T ⎛ 2 ⎞ 2(γ −1) c* = o c , where Γ = γ Γ M ⎝⎜ γ +1⎠⎟ 3 2 2 Ro = universal gas constant = 8.3×10 kg m s °K Tc = chamber temperature, °K M = exhaust gas mean molecular weight γ = ratio of specific heats (~1.2-1.4) 33 33 Rocket Characteristic Velocity, c* p A c* = c t = exhaust velocity if C = 1 m! F 34 34 17 2/12/20 Rocket Thrust Coefficient, CF (γ −1) γ Thrust ⎛ 2γ ⎞ ⎡ ⎛ p ⎞ ⎤ ⎛ p − p ⎞ A C = = λΓ ⎢1− e ⎥ + e ambient e F p A ⎝⎜ γ −1⎠⎟ ⎝⎜ p ⎠⎟ ⎝⎜ p ⎠⎟ A c t ⎣⎢ c ⎦⎥ c t Thrust = λ m! ve + Ae ( pe − pambient ) λ :reduction ratio (function of nozzle shape) CF typically 0.5 - 2 35 35 Thrust Coefficient, CF, vs. Nozzle Expansion Ratio • xx 36 36 18 2/12/20 Mixture Ratio, r m! oxidizer m! total r = ; m! fuel = ; "leaner" < r < "richer" m! fuel 1+ r • Stoichiometric mixture: complete chemical reaction of propellants • Specific impulse maximized with lean mixture ratio, r (i.e., below stoichiometric maximum) 37 37 Effect of Pressure Ratio on Mass Flow In choked flow, mass Choked flow flow rate is maximized occurs when Γ p A γ γ −1 m! = c t p ⎛ 2 ⎞ e ≤ ≈ 0.53 R oTc ⎜ ⎟ pc ⎝ γ +1⎠ M 38 38 19 2/12/20 Combustion Instability • Complex mix of species, phases, pressures, temperatures, and flows • Cavity resonance Harrje, NASA SP- 194, 1972 39 39 Combustion Instability Stable Response to Disturbance Unstable Response to Disturbance 40 Harrje, NASA SP-194, 1972 40 20 2/12/20 Shock Diamonds When pe ≠pa, exhaust flow is over- or underexpanded Effective exhaust velocity < maximum value Viking https://www.youtube.com/watch?v=qiMSko4HBe8 41 41 Rocket Nozzles 42 42 21 2/12/20 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 43 43 Rocket Nozzles 44 44 22 2/12/20 Linear Spike/Plug Nozzles 45 45 Throttling, Start/Stop Cycling CECE Demonstrator Pintle Injector 46 46 23 2/12/20 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 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 48 24 2/12/20 Nuclear Propulsion 1 R T c* = o c Γ 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 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 • Isp = 850 – 1,000 sec • T / W ~ 7:1 50 50 25 2/12/20 Project Rover, 1955-1972 NERVA Rocket, Isp ~ 900 sec NERVA-Powered Mars Mission Kiwi-B4-A Reactor/Rocket 51 51 52 52 26 2/12/20 Liquid/Particle-Core Nuclear Rocket • Nuclear
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  • Enhancement of Volumetric Specific Impulse in HTPB/Ammonium Nitrate Mixed

    Enhancement of Volumetric Specific Impulse in HTPB/Ammonium Nitrate Mixed

    Utah State University DigitalCommons@USU All Graduate Plan B and other Reports Graduate Studies 12-2016 Enhancement of Volumetric Specific Impulse in TPB/H Ammonium Nitrate Mixed Hybrid Rocket Systems Jacob Ward Forsyth Utah State University Follow this and additional works at: https://digitalcommons.usu.edu/gradreports Part of the Propulsion and Power Commons Recommended Citation Forsyth, Jacob Ward, "Enhancement of Volumetric Specific Impulse in TPB/AmmoniumH Nitrate Mixed Hybrid Rocket Systems" (2016). All Graduate Plan B and other Reports. 876. https://digitalcommons.usu.edu/gradreports/876 This Report is brought to you for free and open access by the Graduate Studies at DigitalCommons@USU. It has been accepted for inclusion in All Graduate Plan B and other Reports by an authorized administrator of DigitalCommons@USU. For more information, please contact [email protected]. ENHANCEMENT OF VOLUMETRIC SPECIFIC IMPULSE IN HTPB/AMMONIUM NITRATE MIXED HYBRID ROCKET SYSTEMS by Jacob W. Forsyth A report submitted in partial fulfillment of the requirements for the degree of MASTER OF SCIENCE in Aerospace Engineering Approved: ______________________ ____________________ Stephen A. Whitmore Ph.D. David Geller Ph.D. Major Professor Committee Member ______________________ Rees Fullmer Ph.D. Committee Member UTAH STATE UNIVERSITY Logan, Utah 2016 ii Copyright © Jacob W. Forsyth 2016 All Rights Reserved iii ABSTRACT Enhancement of Volumetric Specific Impulse in HTPB/Ammonium Nitrate Mixed Hybrid Rocket Systems by Jacob W. Forsyth, Master of Science Utah State University, 2016 Major Professor: Dr. Stephen A. Whitmore Department: Mechanical and Aerospace Engineering Hybrid rocket systems are safer and have higher specific impulse than solid rockets. However, due to large oxidizer tanks and low regression rates, hybrid rockets have low volumetric efficiency and very long longitudinal profiles, which limit many of the applications for which hybrids can be used.