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" Average specific gravity of solid propellant. b Conditions for I, and c* :

Combustion chamber pressure: 1000 psia Nozzle exit pressure: 14.7 psia Optimum nozzle expansion ratio Frozen equilibrium

" Average specific gravity of solid propellant. b Conditions for I, and c* :

Combustion chamber pressure: 1000 psia Nozzle exit pressure: 14.7 psia Optimum nozzle expansion ratio Frozen equilibrium

In gas generators and preburners (see Section 10.5), for staged combustion cycle rocket engines (explained in Section 6.5) the gas temperatures are much lower, to avoid damage to the turbine blades. Typically, the combustion reaction gases are at 900 to 1200 K, which is lower than the gas in the thrust chamber (2900 to 3600 K). The thermochemical analysis of this chapter can also be applied to gas generators; the results (such as gas temperature T\, the specific heat cp, specific heat ratio k, or composition) are used for estimating turbine inlet conditions or turbine power. Examples are listed in Table 5-10 for a chamber pressure of 1000 psia. Some species in the gases will not be present (such as atomic oxygen or hydroxyl), and often real gas properties will need to be used because some of these gases do not behave as a perfect gas at these temperatures.

TABLE 5-7. Variation of Calculated Performance Parameters for an Aluminized Ammonium Perchlorate Propellant as a Function of Chamber Pressure for Expansion to Sea Level (1 atm) with Shifting Equilibrium

Chamber pressure (psia)

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