It has been reported in literature that the use of protrusion in the combustion chamber of a hybrid rocket motor enhances the regression rate. This study reports careful experiments conducted to show that the improvement in burn rate with protrusions is only up to a certain fraction of the overall burn time. From the results obtained, it is seen that an X/L of 0.5 is the best location for a graphite protrusion. The protrusions are also shown to increase the combustion efficiency by as much as 45 % when placed at an X/L of 0.5. This, more than the improvement in the regression rate is very useful, especially for small scale motors whose combustion efficiencies are otherwise very low.
Nomenclature
Al= aluminum AP = ammonium perchlorate A t = throat area, m 2 C p = specific heat of propellant, kJ/kg-K C * exp = experimental characteristic velocity,m/s C * theo = theoritical characteristic velocity, m/s d i = intial port diameter, m d f = final port diameter, m G ox = oxidizer mass flux, g/cm 2 s I sp = specific impulse, N-s/Kg L g , L = length of fuel grain, m L = characteristic length, m m f = mass of fuel burnt, gṁ f = mass flow rate of fuel, g/ṡ m ox = mass flow rate of oxidizer, g/ṡ m tot = total mass flow rate, g/s n = flux exponent O/F = oxidizer to fuel ratio P c = average chamber pressure, bar P E = polyethylene t b = burn time, s X = axial distance of protrusion, m ρ f , ρ p = density of fuel, kg/m 3 η c * = combustion efficiency λ p = thermal conductivity of propellant, W/m-K τ p = thermal response time, s
