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Comparison ofLiąuid Pr opel fant Rocket Engine Feed Systems - 1 - 13

. gg

-RT_in

(2.9.4)

where, Rgg: Turbinę drive gas constant (J/kgK). y_RI!: Turbinę drive gas specific heat ratio.

T,i„: Turbinę drive gas temperaturę (K).

Pdtu- Turbinę discharge pressure (Pa).

Pm,: Turbinę inlet pressure (Pa).

Thus, from eąuations 2.9.2 to 2.9.4 and 2.1.4, the required propellant pump power can be found as function of the gas generator hot gas parameters (This expression can be verified both in [1] and [2]):

P_otu

^{= r}lu

k “ U.J

<2.9.5)

being, M^ the turbinę impels gasses molar mass (kg/kmol).

Finally, the propellant total mass that impels the turbinę can be estimated combining the previous expression with 2.9.1 and 2.3.1, giving:

"V = 0 + *„—

Pcmp	“/ + «. 1	^M0,(7a, “')
1.	vJ	R v II ' gg

'	f
TIIU	1-	[eĄ”
		i Pm)
\		J)

(2.9.6)

III. DATA ESTIMATION

To tracę the results curves it is necessary to assume some data values. In this section the methods and the sources of such estimations are detailed.

3.1. Combustion parameters

First, a propellant combination for the rocket engine is adopted. In this case, the comparisons will be performed using:

•    Fuel: Mono-methyl Hydrazine (MMH)

•    Oxidizer: Nitrogen Tetroxide (NTO)

From which the following data are transcribed [2]:

Table 1: Characteristics of chosen propellants.
Propellant	Composition	Density [kg/m^3]	Materiał compatibility
MMH	CH3NH-NH2	878	Al. SS, Teflon, Kel-F. Polyethylene
NTO	n,o4	1440	Al, SS, Ni. Teflon