Presentation on theme: "MAE 5391: Rocket Propulsion Overview of Propulsion Systems."— Presentation transcript:
MAE 5391: Rocket Propulsion Overview of Propulsion Systems
2 Rocket Technologies
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 Chemical Limitations Why we have thermo! V exit = nozzle exit velocity (m/s) R u = universal gas constant ( 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 Cold Gas GasMolecular Weight Specific Impulse (sec) Air Argon CO Helium Hydrogen Nitrogen Methane Cold Gas: Expand a pressurized gas through a nozzle
6 Liquid Monopropellant ParameterValue CatalystLCH 227/202 Steady-state thrust (N) Isp (sec) 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) Chamber pressure (bar) Nozzle expansion ratio61:1 Mass flow rate (gm/sec) Valve power27 W 28 VDC Thruster mass (kg) N 2 H 4 4 NH 3 + N ,280 joules MonoProp: Decompose a single propellant and expand the exhaust through a nozzle
7 Liquid Bi-Propellant StorableIsp sec finert= Cryogenic Isp 320 – 452 sec finert= BiProp: Combust (burn) two propellants (fuel + oxidizer) in a combustion chamber and expand exhaust through a nozzle Finert = Finert=
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 = sec Finert= , 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 When solids go bad!
10 Hybrids Isp= 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 Nuclear Thermal Propulsion NERVA Program Thrust = 890,000N Isp = 838 sec Working fluid = Hydrogen Test time = 30 minutes Stopped in 1972 Finert= (shielding)
12 Electrothermal-Resistojets Working Fluid Thrust (mN)Isp (sec)Power (W)Cp (kJ/kg K)Tc (K) hydrogen water nitrous oxide 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 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 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 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 Pulsed Plasma Thrusters Isp = sec P = 1 – 100 W Thrust = 5 N/W
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