lecture_25 - 16.512 Rocket Propulsion Prof Manuel...

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16.512, Rocket Propulsion Prof. Manuel Martinez-Sanchez Lecture 25: Basic Turbomachine Performance Turbopump Pressurization Systems 1. Cycles For higher performance, mechanical pumps must be used to feed the combustion chamber. In turn, these pumps require drive power, which is always provided by turbines using excess thermal energy in the propellants (although electrical motors have been considered for small rockets). The manner in which hot gas is provided to drive the turbines serves to distinguish among several pressurization cycles , of which the most important are summarized in Fig. 1. (From Ref. 41). The most common of these are the bi-propellant gas generator (G.G.) cycle, the expander cycle and the staged combustion cycle. (a) The Gas Generator Cycle The GG cycle was used in the F1 engine, and is also in use in the Delta II, Atlas and Titan rockets. In this cycle, a small fraction of the pressurized oxidizer and fuel are diverted to a medium-temperature burner (Gas Generator), which produces typically very fuel-rich gas to drive the turbine or turbines. These are designed with a large pressure ratio, and their exhaust is either dumped overboard, or injected at some point into the main nozzle to provide some extra thrust. Nevertheless, this cycle is inherently somewhat lossy, in that the turbine gas is not fully utilized in the main combustor. On the other hand, the power control is relatively straightforward, and there is little interaction of the feed system with the rest of the rocket. Any propellant combination can be used, all power levels are suitable, and any desired pressure level can be obtained although the I sp loss increase with pressure (1.5-4 sec per 100 atm). The mechanical power required to drive the pumps is •• =+ OP FP ox F P oP ox FP F P P Pm m η ρη ρ (1) where are the pressure rise in the oxidizer pump (OP) and fuel pump (FP), respectively which have efficiencies ∆∆ OP FP Pa n dP ηη OP FP , . Also ρρ ox F , are the liquid densities. If the gas generator mass flow rate is , the (single) turbine power is GG m = '' GG TT P Pm c T n T T ( 2 ) in which η T is the turbine isentropic efficiency (60-80%), η ' TT is its thermodynamic expansion efficiency () 1 1 =− ' ' ' TT ie n P/P γ ( 3 ) which depends on pressure ratio and GG gas specify heat ratio, ie n ' γ . Also, is the specific heat of this gas, and T ti its temperature, which is controlled through stoichiometry to values acceptable by uncooled turbines (700-1100K). ' P c Ref. 41: Turbopump System for Liquid Rocket Engines . NASA SP-8107, Aug. 1974. 16.512, Rocket Propulsion Lecture 25 Prof. M. Martinez-Sanchez Page 1 of 11
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Equating and yields the required . Suppose now the turbine exhaust is expanded to the same exit pressure as the main flow. The exhaust speed is then P P T P GG m e P () 1 21 =− ' ' GG P t.out e te cc T P / P γ (4) where ( 1 ' t.out n T TT ηη ) ( 5 ) This speed is generally lower than the main flow exit speed, 1 == pc e c T ; ( P / P ) (6) where c p and η
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lecture_25 - 16.512 Rocket Propulsion Prof Manuel...

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