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Propulsion o o o o Basic propulsion Propulsion performance Propulsion types Launch vehicles Basic Propulsion Propulsion Basics Three primary measures of propulsion system effectiveness and efficiency are: 1. Thrust - the force available for launch or flight 2. Thrust duration - the time available for acceleration 3. Thrust efficiency - measured by the specific impulse = Isp Propulsion Basics Three primary measures of propulsion system effectiveness and efficiency are: 1. Thrust - the force available for launch or flight 2. Thrust duration - the time available for acceleration 3. Thrust efficiency - measured by the specific impulse = Isp Propulsion Basics 1. Thrust - propulsion thrust used for launch, orbital change, orbital maintenance, station keeping operations, and attitude control Launch - very high thrust required (106 -107 Newtons (1 Newton = 0.2248 lb) Examples are the Atlas, Delta, Proton, Ariane 5, Long March Apogee and orbit boost - moderate thrust needed (103 -105N) Examples are Inertial Upper Stage (IUS), Payload Assist Module (PAM), Centaur upper stage Attitude control - low thrust needed (1-100 N or less typical) - short duration bursts and frequent operational cycles Propulsion Basics 2. Thrust duration - a measure of the total energy available in the propulsion system Short duration (.01 - 10 sec) Primarily used for attitude control Intermediate duration 10 – 1,000 sec Used for launch and boost Long duration 103 - 107 sec Used for deep space propulsion (ion, nuclear) Propulsion Basics 3. Thrust efficiency - measured by specific impulse, Isp (units are seconds) Low (1-100 s) Used in attitude control on smaller spacecraft Moderate (100-500 s) Launch & boost; attitude control on larger spacecraft High (500 – 5,000 s) Interplanetary propulsion, station keeping Propulsion Basics Momentum = p = mass x velocity Measured by combustion exhaust mass flow rate Force = F = dp/dt = d(mv)/dt = vdm/dt + mdv/dt F = mdv/dt = ma for constant mass Thrust = T = Ve dm/dt (mass rate flow times exhaust velocity) Thrust comes from exhaust mass reaction force and the pressure difference between the rocket nozzle pressure and the ambient (surrounding) pressure Propulsion Basics Thrust from internal (nozzle) and external (ambient) pressure difference T = PeAe - PaAe = Ae(Pe - Pa) where Pe = exhaust pressure, Ae = exhaust chamber area, and Pa= ambient pressure The greater difference between internal pressure, Pe, and external pressure, Pa, the greater the force Total thrust = Ve dm/dt + Ae(Pe - Pa) Propulsion Basics Exhaust velocity - Ve The chemical rocket motor performance is determined primarily by the exhaust velocity - higher exhaust velocity improves thrust performance Ve is often expressed as: where Ve = exhaust velocity, Ro = universal gas constant Tc = combustion chamber temp m = molecular weight of exhaust gas Pe = exhaust pressure, Pc = chamber pressure γ = gamma = specific heat ratio (function of fuel chemistry) Propulsion Basics The equation shows that increased exhaust velocity (and performance) is available by: Increasing chamber-to-exhaust pressure ratio (vacuum of space makes the chemical rocket more efficient in space than in the atmosphere Increasing combustion chamber temperature (limitation on this temperature is based on the chamber and nozzle design and materials) Decreasing molecular weight - Hydrogen the best fuel but needs oxidizer (H2 + O2 is one of the best propellant combinations) Lower specific heat ratio - This is a property of the fuel chemistry and has a limited range available Propulsion Basics Specific impulse Impulse = I = Thrust (force) x time increment Δt, or I = TΔt Specific Impulse = Isp = I/mg where m = propellant unit mass, g = gravitational acceleration at the Earth's surface Isp = T/mg for time increment Δt of 1 sec, a constant thrust and a constant mass flow Propulsion Basics Specific impulse Isp is a measure of efficiency either for propellants or the motor Defined as the thrust per unit weight flow rate of propellant High Isp = high thrust efficiency per unit mass important especially in mass-critical designs Propulsion Basics Propulsion Basics Nozzle design An important design criteria for the chemical rocket motor is the exhaust nozzle Exhaust velocity is maximized and pressure differences are optimized within the nozzle parameters, from curvature to internal diameter ratios This is accomplished by the nozzle first constricting the exhaust gas flow to increase the exhaust velocity, then provide a divergent flow to accelerate the gas to even greater velocity Propulsion Basics Nozzle design Propulsion Basics Nozzle design A number of other parameters must be optimized for the nozzle design, including the shape of the throat area and the angles of convergence and divergence Another important element in exhaust nozzle design is the area ratio and resulting gas expansion, which also represents the maximum exhaust velocity This ratio is also important since the nozzle is matched to the combustion chamber output and the ambient pressure(s) at the nozzle exit Propulsion Basics Nozzle design The area ratio is defined as the exit area divided by the throat area, or e = Aexit/Athroat Ideally, the nozzle expansion curve would produce the same pressure at the exit point along the wall as the ambient pressure at the nozzle exit While this suggests an extended nozzle, weight limitations and aerodynamic pressures and loads restrict the length of the nozzle, as does the cooling mechanism Propulsion System Performance Propulsion Basics Propulsion system performance The performance of a rocket motor or engine can be measured in several ways, but the most common are Isp ΔV Ve ΔV is the velocity change, or acceleration times the thrust duration = a Δt Propulsion Basics ΔV ΔV = acceleration x Δt = F Δt/m = Isp g This a dimensional relationship only - the more accurate representation is: ΔV ideal performance = Isp g ln[mi/(mi - mp)] were ln = natural log, mi = initial vehicle mass, mp = mass of consumed propellant, mi - mp = mf = empty (final) vehicle weight (less is better) Propulsion Basics ΔVideal performance = Isp g ln[mi/(mi - mp)] is the basic relationship between ΔV (performance) and Isp for a single stage propulsion system. This is not the same as the ΔV required to attain proper orbit, or for orbit changes. This equation allows us to calculate the initial-to-final mass ratio for the launch vehicle or booster stage (plus payload) by simple algebraic rearrangement. mi/mf = exp[ ΔVrequired/(g Isp)] (50-200 typical for mi/mf ) To calculate the propellant mass use the same equation, but solve for mp mp = mi[1- exp[ ΔVrequired/(g Isp)]] Propulsion Basics Note that the actual ΔV required needs to include the gravitational force when departing to or approaching a planet, and the drag force when passing through an atmosphere, both reducing the ΔVideal available The extra ΔV for gravity and drag usually requires 1500 to 2000 m/s for departure from Earth Propulsion Basics ΔVactual required = ΔVideal + ΔVg + ΔVD ΔVg = ΔV loss due to gravity = g x Δt were, g = acceleration of gravity at height r (9.89 m/s2 at the surface of the Earth), Δt is the time variable which is to be integrated over the time of flight ΔVd = ΔV loss due to aerodynamic drag Orbit/Trajectory Launch to LEO (400 km) (excluding gravity, atmospheric drag) ). Gravity & drag during ascent through atmosphere (typical) Flight path correction from vertical to horizontal Launch to LEO (total) ΔV (m/s) Percent Payload 7,750 100% (reference) 1,400 350 9,500 100% (reference) Launch to GEO Lunar impact Lunar landing (soft) Launch to lunar orbit Circumlunar mission with LEO return 10,200 12,500 14-15,000 13,500 16,000 10-25% 35-45% 20-30% 20-30% 25-35% Lunar landing and return 16-18,000 1-4% Mars Mars - landing and return 11,390 23-27,000 20-30% 0.1-1% Venus Venus - soft landing 11,450 23-25,000 Mercury Jupiter Sun Earth Escape Escape Solar System 12,500 13,930 30,450 12,700 17,500 Chart of maximum vehicle velocity versus vehicle initial-to-final mass ratio (minitia/lmfinal, or 1/MR) in an idealized drag -free, gravitationless flight environment Plotted are five Isp values which show the maximum ΔV limitations of propulsion efficiency (Isp) Single stage rockets are typically limited to a 1/MR of approximately 20, while staged vehicles are capable of mass ratios of 200 or more (Sutton) ). Propulsion Basics From the above chart, the limitation of staged chemical rocket motors with a maximum Isp of roughly 450 s is approximately 20,000 m/s with a initial-to-final mass ratio of 150 Included in the final mass fraction, mf, are the vehicle inert mass (structure, tanks, engine, residual fuel, etc.), as well as the payload mass. The propellant mass fraction, expressed as 1 - mf/mi, would ideally be 1, although completely impractical since the entire vehicle would be propellant. The ideal propellant mass fraction of 1 would also mean that the final-to-initial mass ratio be zero, or the final mass is zero. Also, this is an obviously impractical design. Propellant mass ratios beyond 0.85 require careful design, with a practical limit near 0.95. Propulsion Basics Limitations on mass ratio For deep-space or interplanetary missions, a high payload capacity is desirable Ideally a small final-to-initial mass fraction with a minimal remaining vehicle mass The propellant mass fraction would be modest to accommodate a payload, on the order of 0.70 to 0.80 Propulsion Basics Limitations on mass ratio Final-to-initial mass ratio (MR) would be 0.1 to 0.2 (1/MR = 10-20) Another way to look at the ΔV limitation is to use the final-to-initial mass ratio mf/mi, = e -ΔV/g Isp = e -(ΔV/Ve) if we substitute the exhaust velocity Ve for Isp times g Propulsion Basics Limitations on mass ratio For an interplanetary payload, the vehicle's exhaust velocity needs to be comparable to the mission velocity requirement (ΔV) A vehicle with a 400 s Isp would have difficulty delivering a sizable fraction of the total remaining vehicle mass as payload beyond Mars since g Isp is of the order of 4,000 m/s Needs to be 12,000-14,000 m/s for the ΔV requirement beyond Mars • Available payload mass for chemical rockets are nearly insignificant beyond Jupiter • Higher Isp electric boosters are capable of reaching our Galaxy's interstellar environment • Missions beyond Jupiter are possible for chemical rockets using gravity assists ). Propulsion Basics Propulsion calculations Calculate the empty weight of the vehicle from the known Isp and initial weight Step 1 Calculate the initial-to-final mass ratio for a lunar flyby mission with a ΔV of 12,500 m/s and a LOX/LH2 propulsion system with an Isp of 435 sec and a final mass of 1.811x106 kg (g = 9.80 m/s) mi/mf = e[ΔVrequired/(g Isp)] mi/mf = e[12,500 m/s/(9.80 m/s x 435 sec)] = e2.93 = 18.77 Propulsion Basics Propulsion calculations Step 2 Working backwards and knowing to final mass of the vehicle (empty weight of vehicle plus payload weight), find the final mass of the vehicle mf = mi/e[ΔVrequired/(g Isp)] but its easier than this since mf = 1/(mi/mf) x mi mf = 1/18.77 x 1.811x16 kg = 96,480 kg Propulsion Basics Propulsion calculations Step 3 Find the empty mass of the vehicle knowing the payload mass, which is 8,600 kg mempty weight = mf - mpayload = 96,480 - 8,600 kg = 87,880 kg Propulsion Basics Propulsion calculations Step 4 Find the propellant mass required for the mission using mp = mi {1-e[-ΔVrequired/(g Isp)]} = 1.811x106 kg {1 - e[-12,500 m/s/(9.80 m/s x 435 sec]} = 1.811x106 kg {1 - 0.0533) = 1.715x106 kg Propulsion Basics Propulsion calculations Step 5 From the previous calculation, the propellant mass fraction can be found with a simple ratio mp/mi = 1.715x106 kg/1.811x106 kg = 0.947 (94.7% fuel) For comparison, the payload mass fraction is mpayload/mi = 8,600 kg/1.811x106 kg = 0.0047 = 0.47% payload with 99.53% vehicle and propellant Propulsion Types Chemical Rockets 1. Chemical (solid, liquid, hybrid) Basic requirements for a chemical, high thrust, high performance system • High temperature chemical reaction • Moderate to high Isp • Low exhaust gas molecular weight Chemical Rockets Liquid bipropellant - separate fuel and oxidizer Chemical Rockets Chemical liquid – Monopropellant Single fuel with a self-contained oxidizer Most common fuel is hydrazine N2H4 (goes to NH3, N2, H2) Variations are monomethyl hydrazine, unsymmetrical-dimethyl hydrazine Hypergolic when passed over a catalyst surface. The simplicity and moderate performance of this system make this useful for many spacecraft propulsion applications Chemical Rockets Liquid monopropellant Oxidizer Specific gravity Liquid oxygen 1.14 (LOX) Hydrogen peroxide (H2O2) Nitric acid (HNO3) Boiling point (1 atm) Characteristics 90 K (-183oC, -298oF) •Not hypergolic but can combust spontaneously with many materials at elevated pressures •Most commonly used rocket fuel oxidizer •Non-toxic and non-corrosive 423 K (150oC, 302oF) •Oxygen and heat are released by the decomposition of hydrogen peroxide into H20 + O2 •Decomposition is spontaneous with exposure to a catalyst such as platinum or iron oxide •H2O2 was used to generate gas to drive turbopumps in the V-2, X-1 and X-15 o o 1.26-1.41 356 K (83 C, 181 F) •Nitric acid and its variants are highly corrosive •Red fuming nitric acid is nitric acid + 5-20% nitrogen dioxide; more stable, less corrosive than pure nitric acid •Addition of <1% fluorine ion (HF) reduces corrosion (inhibited red fuming nitric oxide) •Used as propellant oxidizer with gasoline, amines, and hydrazine Nitric acid is hypergolic when combined with hydrazine and amines 1.19 •Mildly corrosive unless mixed with water •Spontaneous combustion occurs when exposed to many materials •NTO is hypergolic when combined with most fuels •High vapor pressure requires relatively heavy tank •Used in numerous Russian rockets, the Titan booster series, and the Space Shuttle attitude control •Highly toxic - exposure limit < 5 ppm Nitrogen tetroxide (N2O4, NTO) 1.44 291 K (18oC, 64oF) Fluorine 1.11 83 K (-190oC, -310oF) •Fluorine and fluorides have been proposed in various fuel combinations which are highly corrosive, difficult to handle, and toxic •No fluorine oxidizers have been used for production rocket engines Fuel RP-1 Specific gravity Boiling point (1 atm) Characteristics 0.80-0.815 420 K (147oC, 297oF) •Highly-refined kerosene • Developed as a fuel that could also be used for cooling high-temperature nozzles and combustion chambers • Sulfur, aromatics, and unwanted isomers removed to permit use at high temps • Greater stability, lower toxicity, less residue, higher performance than other hydrocarbons •High flash point 336 K • Safer, less explosive than many hydrocarbon fuels including gasoline •Used in Russian R-7 booster and its derivatives, Soviet N-1, Atlas, Thor, Delta I-III, Titan I, Saturn I, IB, V (1st stage) Liquid hydrogen (H2, LH2) 0.07 (requires large fuel tanks) 20 K (requires •High specific impulse extensive insulation for tank and feed •Highly flammable when hydrogen gas is mixed with air lines) •Increased density possible with supercooled solid or slush hydrogen (not yet used) •Non-toxic (breathable gas; can replace nitrogen in an artificial atmosphere) •Non toxic exhaust gas when reacted with oxygen •Hydrogen britalizes most metals, making turbopump design more challenging than with other fuels Fuel Methane (CH4) Specific gravity 0.47 Boiling point (1 atm) Characteristics 90 K (-183oC, -298oF) •Hydrocarbon fuel •Stored cryogenically •Low cost; freely available from gas wells, biomass decomposition •Potentially useful for Mars return missions •Under research but not used in production liquid fuel engines •Possible fuel for arc-jet or resistojet thrusters Hydrazine (N2H4) 0.80 Monomethyl 0.88 hydrazine (CH3NHNH2, MMH) 387 K (186oC, 303oF) •Used as both a monopropellant and a bipropellant fuel •Hypergolic fuel as monopropellant •Hypergolic when mixed with nitrogen tetroxide and with nitric acid •Spontaneous ignition also possible if it is spilled as a liquid in air, and in contact with many materials •Used in the production of stainless steel, nickel and some aluminum alloys •Not used in iron, copper, or some aluminum alloys •Very long storage life •Highly toxic • Exposure limit < 0.1 ppm • Known carcinogen •Has been used as a monopropellant for gas generators (e.g. Space Shuttle hydraulic system), and spacecraft attitude control 360 K (87oC, 189oF) •Better liquid temperature range than hydrazine •Lower reaction threshold to shock waves than hydrazine •Slightly lower Isp than for hydrazine •Highly toxic • • Exposure limit < 0.2 ppm Suspected carcinogen Oxidizer Liquid oxygen (LOX) Fuel Isp (theoretical, Isp (theoretical, vacuum) 1 atm) Liquid hydrogen (lLH2) 477 s 390 S LOX Kerosene (RP-1) 370 s 300 s LOX Monomethyl hydrazine 365 s 301 s LOX Methane (CH4) 368 s 296 s Liquid ozone (O3) Hydrogen 580 s Nitrogen tetroxide (N2O4) Hydrazine (N2H4) 334 s Red fuming nitric acid RP-1 Hydrogen peroxide (H2O2) Monopropellant H2O2 Fluorine (Fl) Fl 292 s 269 s 154 s (90% H2O2) RP-1 279 s Lithium 542 s Hydrogen 580 s 410 s Propulsion Basics Liquid propellant complications A. Zero gravity fuel feed Liquid fuels are not confined to any specific region within the tank in low or micro gravity but there are several ways to allow for positive fuel feed during low gravity conditions Capillary devices - use surface tension to keep gas and liquid separated in the tank (requires pressurization). Used on Shuttle and Viking Diaphragms and bladders - physically separate gas an liquid with flexible lining made of elastomer or Teflon (requires pressurization). Used on Magellan and Voyager Bellows - an expandable metal device to separate gas from liquid (requires pressurization) Propulsion Basics Liquid propellant complications B. Temperature extremes Operating temperature range of liquid propellants is limited and is a function of pressure Cryogenic liquids must be kept at low temperatures at moderate pressures Storage system, transfer and pumping system, and the combustion components must be able to reliably withstand high temperature extremes (20K [storage] to 6,000K [combustion]) Hydrazine and its variations and nitrogen tetroxide do not require cryogenic systems but may require electric strip heaters Propulsion Basics Liquid propellant complications C. Oscillations Liquid propellants can oscillate or slosh while in the tanks, especially during high thrust during launch Potential for a serious effect on the tank structure as the fuel surges up and down Baffles, reinforcements and overall tank design are used to reduce this effect Chemical Rockets Chemical - Solid Solid rocket fuel is typically identified by the type of chemical binder used HTPB (hydroxy-terminator polybutadiene) is a stronger binder, more flexible, and faster curing, but suffers from a slightly lower Isp than PBAN (used on Delta II, Delta III, Delta IV, Titan IVB and Ariane PBAN (polybutadiene acrylic acid acrylonitrile) and uses fast-curing, toxic isocynates with a slightly higher Isp, is less costly, less toxic, and used in the larger boosters (Titan III, the Space Shuttle SRBs, and the new SLS) HTPB is or has been Electric Propulsion Electric Propulsion Electric propulsion entails either 1. Accelerating charged particles; or 2. Heating cold gas Both processes use electrical power Electric and/or magnetic fields are used to accelerate charged (ionized) particles Electrothermal, microwave solar, nuclear, or arc current methods are used to heat cold gas Electric Propulsion Advantages High Isp Reduced propellant mass (or increased payload mass) Increased mass savings with increased duration/distance Disadvantages High power requirement High system mass possible Ion plume can degrade surfaces and create charge buildup High cost Electric Propulsion Common electric engines Electrothermal Resistojet thruster Arcjet thruster Solar thermal thruster RF-heated (microwave) thruster Electrostatic Ion thruster Electrodynamic Magnetoplasmadynamic thruster (MPD) Hall-effect thruster Pulsed-plasma thruster Variable specific impulse thruster Nuclear Nuclear thermal ion Electric Propulsion Electromagnetic field ion engine Ion production Electrical (resistive) heating Electron excitation Microwave excitation Nuclear heat source Ion acceleration Electrostatic field Magnetic field Current flow in electric/magnetic field Electric Propulsion Electrostatic ion engine Electric field accelerates oppositely charged ions in opposite directions Force is proportional to field strength and ion charge F = qE where E is the electric field strength and q is the ion’s charge Vexit = 2qE/m where Vexit is the electric charge exit speed and m is the ion mass High charge/mass ratio is desired Xenon common today Krypton, cesium, mercury also used Thrust proportional to power/current T = Pη/(gIsp) where T is thrust, P is thrust power, and η is thrust efficiency Electric Propulsion Electrostatic ion engine requirements High electrostatic field (voltage) High charge/mass ratio propellant Highly-ionized high atomic mass gas Vacuum conditions Neutralized ion beam to prevent charge separation Electron beam injected into ion beam as it exits engine Electric Propulsion Electrostatic ion engine example (this example includes magnetic field confinement) Electric Propulsion Electrodynamic ion engine A. Magnetoplasmadynamic ion engine Also uses electrostatic field but indirectly Most applications are the Includes magnetic field acceleration from a radial current flow Current flow between cathode and anode produces an induced magnetic field Current of ions in self-generated magnetic field produces axial acceleration F = qV X B where the accelerating force is equal to the ion charge times the ion’s velocity as a cross product with the induced magnetic field strength Additional magnetic field around device increases acceleration force Electric Propulsion Magnetoplasmadynamic ion engine Electrodynamic ion engine example Electric Propulsion Electrodynamic ion engine B. Hall effect (plasma) thruster Ions accelerated in an axial electrostatic field Radial magnetic field accelerates ions axially Lower efficiency than electrostatic engine but greater thrust range Electrons forward towards anode Positive ions rearward towards exhaust Divergent exhaust plume deceases net exit velocity Common applications include station-keeping on geostationary satellites Highest power requirement and highest thrust of typical electrical power (EP) thrusters Electric Propulsion Hall effect (plasma) thruster Electric Propulsion Hall effect (plasma) thruster Photograph of experimental Hall-effect thruster with dual (inner & outer) magnetic rings and electron injector Electric Propulsion Electrodynamic ion engine C. Variable specific impulse magnetoplasma engine Ions accelerated by electromagnetic field and microwave (RF) energy Intended to bridge the gap between high-thrust, low-specific impulse propulsion systems and low-thrust, high-specific impulse systems Can operate in either mode Developed by astronaut Franklin Chang-Diaz Also called the variable specific impulse magnetoplasma rocket (VASIMR) Electric Propulsion Electrothermal (heating cold gas) Two types currently used Resistojet Arcjet Heating mechanisms Electrical (resistive or arc) heating Microwave excitation Nuclear heat source Solar heating Modest performance Isp modest but greater than cold gas Modest thrust but greater than cold gas Electric Propulsion Electrothermal engine - resistojet Cold gas injected into an expansion chamber is accelerated by heating the gas Heating possible with several methods Non-reactive and reactive gases used Non-reactive – more efficient Electrical (resistive) heating Microwave excitation Nuclear heat source Solar heating Nitrogen, hydrogen examples Reactive gases – less efficient Hydrazine, methane are examples Electric Propulsion Electrothermal engine resistojet Electric Propulsion Electrothermal engine - arcjet Gas heating from high current arc between interior cathode and anode shell Higher temperatures than resistojet possible, hence higher Isp Higher power requirement than resistojet Non-reactive gases – more efficient Reactive gases – less efficient Electric Propulsion Electrothermal engine - arcjet Electric Propulsion Electrothermal engine - other Solar heating also possible by directing focused light on gas chamber Microwave heating also possible Electric Propulsion Nuclear (not electric except for nuclear ion) Nuclear heating of cold gas can reach extremely high Isp Hot radioactive core can reach much higher temps than resistojet and arcjet Nuclear thermal Additional acceleration of ionized gas with electrostatic/magnetic fields also possible Nuclear ion Electric Propulsion Nuclear thermal (non-electric) Cold gas heated by reactor core to high temperature (5,000 K) and high Isp Electric Propulsion NERVA nuclear thermal reactor model evolution Electric Propulsion NERVA nuclear thermal reactor Electric Propulsion Electric propulsion - general Power proportional to thrust Power supply output proportional to power supply mass Acceleration inversely proportional to spacecraft mass Balance required between thrust power and the mass required for thrust Electric Propulsion Electric propulsion optimization Questions?