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MAE 5391: Rocket Propulsion Overview of Propulsion Systems.

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Presentation on theme: "MAE 5391: Rocket Propulsion Overview of Propulsion Systems."— Presentation transcript:

1 MAE 5391: Rocket Propulsion Overview of Propulsion Systems

2 2 Rocket Technologies

3 3 Propulsion Technology Options  Thermodynamic Systems (TE KE) Cold Gas Thrusters Liquids Monopropellants Bipropellants Solids Hybrids  Nuclear (NE TE KE)  Electric Systems Electrothermal (Resistance Heating) Electrostatic (Ion with E field F=qE) Electromagnetic (plasma with B field F=JxB)  With the exception of electrostatic and electromagnetic, all use concept of gas at some temp flowing though a converging/diverging nozzle!

4 4 Chemical Limitations  Why we have thermo! V exit = nozzle exit velocity (m/s) R u = universal gas constant (8314.41 J/kmol*K) T 0 = chamber temperature (K) P e = exit pressure (Pa) P 0 = chamber pressure (Pa) M= molecular mass of gas (kg/kmol)  = ratio of specific heats (no dimensions)

5 5 Cold Gas GasMolecular Weight Specific Impulse (sec) Air28.974 Argon39.957 CO 2 44.067 Helium4.0179 Hydrogen2.0296 Nitrogen28.080 Methane16.0114 Cold Gas: Expand a pressurized gas through a nozzle

6 6 Liquid Monopropellant ParameterValue CatalystLCH 227/202 Steady-state thrust (N) 11.1 - 31.2 Isp (sec)228 - 235 Propellant specific gravity1.023 Average Density Isp ( sec)236.8 Rated total impulse (Nsec)124,700 Total pulses12,405 Minimum impulse bit (Nsec)0.56 Feed pressure (bar)6.7 - 24.1 Chamber pressure (bar)4.5 - 12.4 Nozzle expansion ratio61:1 Mass flow rate (gm/sec)5.0 - 13.1 Valve power27 W max @ 28 VDC Thruster mass (kg)0.52 3 N 2 H 4  4 NH 3 + N 2 + 336,280 joules MonoProp: Decompose a single propellant and expand the exhaust through a nozzle

7 7 Liquid Bi-Propellant StorableIsp 250-320 sec finert=0.03-.13 Cryogenic Isp 320 – 452 sec finert=0.09-0.2 BiProp: Combust (burn) two propellants (fuel + oxidizer) in a combustion chamber and expand exhaust through a nozzle Finert = 0.04-0.2 Finert=0.11-0.31

8 8 Solids  Composite propellant, consisting of an oxidizing agent, such as ammonium nitrate or ammonium perchlorate intimately mixed with an organic or metallic fuel and binder. Thrust function of burn area, Isp = 250-300 sec Finert=0.06-0.38, 2/3 of motors have fiinert below 0.2 Advantages Simple Reliable High density Isp No chamber cooling Disadvantages Cracks=disaster Can’t restart Hard to stop Modest Isp

9 9 When solids go bad!

10 10 Hybrids Isp= 290-350 sec Finert=0.2 Hybrid: Bipropellant system with liquid oxidizer (usually) and a solid fuel Catalyst Pack Combustion Chamber Nozzle Test Stand Load Cell Fuel Element H 2 O 2 /PE Hybrid Test Set-Up Polyethylene fuel rod

11 11 Nuclear Thermal Propulsion NERVA Program  Thrust = 890,000N  Isp = 838 sec  Working fluid = Hydrogen  Test time = 30 minutes  Stopped in 1972  Finert=0.5-0.7 (shielding)

12 12 Electrothermal-Resistojets Working Fluid Thrust (mN)Isp (sec)Power (W)Cp (kJ/kg K)Tc (K) hydrogen3754610014.321000 water932191002.31000 nitrous oxide1411441001.01000 Electrothermal-- electrical energy is used to directly heat a working fluid. The resulting hot gas is then expanded through a converging-diverging nozzle to achieve high exhaust velocities. These systems convert thermal energy to kinetic energy

13 13 Electrothermal-Arcjets In an arcjet, the working gas is injected in a chamber through which an electric arc is struck. The gas is heated to very high temperature (3000 – 4000 K), Arc temp =10,000K on average, and much greater in certain regions in the arc. Power = 1.8 kW, Isp = 502, Thrust = 0.2N, Propellant = hydrazine

14 14 Electrostatic-Ion Propulsion  Electrostatic-- electrical energy is directly converted into kinetic energy. Electrostatic forces are applied to charged particles to accelerate the propellant. Deep Space 1 = 4.2 kW, Thrust = 165 mN, Isp = 3800 sec 7000 hours of operation is becoming the standard!

15 15 Electromagnetic-MPD Thruster  Electromagnetic-- electromagnetic forces directly accelerate the reaction mass. This is done by the interaction of electric and magnetic fields on a highly ionised propellant plasma. NH 3 MPD, F=23 mN, Isp= 600 sec, P=430 W Stuttgart, Isp=5000sec, F=100N, P=6 MW, hydrogen

16 16 Pulsed Plasma Thrusters Isp = 500-1500 sec P = 1 – 100 W Thrust = 5  N/W

17 17 Hall Effect Thruster Power = 50W – 25kW Isp = 500 – 3000 sec Thrust = 5 mN- 1N

18 18 Propulsion System “Cost”  Performance issues Mass Volume Time (thrust) Power Safety Logistics Integration Technical Risk  The “best” (lowest “cost”) option optimizes these issues for a given set of mission requirements


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