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This Defense Intelligence Reference Document, dated 1 November 2010, was produced by the Defense Intelligence Agency under its Advanced Aerospace Weapon System Applications (AAWSA) Program. It surveys propulsion technologies that include chemical, ion, and nuclear fission rockets, fusion schemes, aneutronic fusion, and antimatter propulsion. It also covers radiation shielding and speculates on research needs over the next 30 years for missions from low Earth orbit to Mars, Jupiter, Saturn, and Alpha Centauri. The document concludes that aneutronic fusion promises to be an important mechanism for future space propulsion.
UNCLASSIFIED/ ,'F811. 8Ffllil-,le lal!i! 8111!¥ COMPARISON OF SPECIFIC IMPULSE FOR VARIOUS ROCKET DESIGNS Chemical Rockets Chemical rockets burn solid or liquid propellant. The exhaust gas exits the rocket through a Laval nozzle and generates thrust. Figure 1 provided the outline of a liquid- fueled rocket using liquid hydrogen as the fuel and liquid oxygen (LOX) to combust the fuel. Unlike turbofan and RAM engines, rockets are not airbreathing and must carry their own oxidizer. The Laval nozzle is a principle component of chemical rockets; its design is based on compressible fluid flow theory. 4 In general, the nozzle is made up of contoured convergent-divergent cross sections. Conical cross sections are also sometimes used. Its purpose is to transform pressure energy into kinetic energy. Nuclear fusion rockets may also make use of Laval nozzles by heating up a liquid propellant and ejecting it as a supersonic gas. In subsonic flow, fluid can only be accelerated by decreasing the cross-sectional area of the duct section that it is traveling through, as in a Venturi tube. Once the velocity in a fluid reaches the speed of sound (Mach 1), the fluid can continue to accelerate only if it is expanded. The Laval nozzle combines a converging section where the flow is subsonic, a throat where the flow is accelerated to sonic speed (Mach 1), and a diverging cone where the flow is accelerated to supersonic speed. The performance of a rocket is based on its thrust, where T = (dm/dt) x Vexhaust, and by maximizing the exit velocity, the thrust reaches a maximum. The pressure of the combusting gases in the combustion chamber directly affect the amount of thrust that the rocket can achieve. Whether a rocket is propelled by gases from combusting propellant or by gases heated through a nuclear fission or fusion reaction, two equations determine the thrust of the Laval nozzle. The maximum mass flow rate through the nozzle can be computed in terms of the nozzle area (Athroat) the combustion or heated gas pressure (Po), and the heated gas temperature (To). ( d '"J = p,, A,,,,,,,,, d t llllX 11/,U/H Jr: Y + I ;(r!1]' (1.15) l ,'j2vRT (p J"V - y-1" I- P",',''.nluw\/ - (1.16) In these equations, R is the gas constant of the propellant gas and v is a thermodynamic property of the gas called the ratio of specific heats. These equations make it possible to compute the maximum thrust generated by a propellant gas through a Laval nozzle based on the pressure and temperature of the gas in the combustion chamber or heating tank. 5 UNCLASSIFIED/ ,1f81il 8fFI&I.«1k WEliii Qtll Y
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Report, from the dia collection. The PDF is mirrored here; the original link is above. 50 pages are in the text index: search them above, or from the library's search.