Burning Rate and Combustion Wave Structure

Ammonium perchlorate (AP) is the crystalline oxidizer most widely used to formulate composite propellants for rockets. Unlike double-base propellants, AP composite propellants are not used as gun propellants because the combustion products of AP composite propellants generate a high mass fraction of hydrogen chloride (HC1), which acts to oxidize and erode the interior surface of gun barrels. 

The temperature in the condensed phase increases from the initial propellant temperature T0 to the burning surface temperature Ts by the conductive heat feedback from the burning surface. Then, the temperature increases in the gas phase because of the exothermic reaction above the burning surface and reaches the final combustion temperature Tg. Since the physical structure of AP composite propellants is highly heterogeneous, the temperature fluctuates from time to time and also from location to location. The temperature profile in Fig. 7-1 is shown as a time-averaged profile. This is in clear contrast with double-base propellants shown in 

The heat flux transferred back from the gas phase to the burning surface is dependent on the temperature gradient in the gas phase, since the temperature gradient 

Propellant Binder HTPB AP particle size (pm) 400 200 20 3 Catalyst BEFP 

The relationship between temperature sensitivity and burning rate is shown in Fig. 7-8 as a function of AP particle size and burning rate catalyst (BEFP)1131. The 

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Measurements of Burning Rate and Combustion Wave Structure... 

In order to clarify the combustion wave structure of AP composite propellants, photographic observations of the gas phase at low pressure are very informative. The reaction rate is lowered and the thickness of the reaction zone is increased at low pressure. Fig. 7.3 shows the reduced burning rates of three AP-HTPB composite propellants at low pressures below 0.1 MPa.FI The chemical compositions of the propellants are shown in Table 7.1. The burning rate of the propellant with the composition ap(0-86) is higher than that of the one with ap(0-80) at constant pressure. However, the pressure exponents are 0.62 and 0.65 for the ap(0-86) and Iap(0.80) propellants, respectively that is, the burning rate is represented by r for the p(0.86) propellant and by r for the p(0.80) propellant. 

Though the physicochemical properties of HTPE and HTPS are different, both are subject to a similar super-rate burning effect. However, the magnitude of the effect is dependent on the type of binder used. As in the case of double-base propellants, the combustion wave structures of the respective propellants are homogeneous, even though the propellant structures are heterogeneous and the luminous flames are produced above the burning surfaces. 

Fig. 12.11 shows the structure of a rocket plume generated downstream of a rocket nozzle. The plume consists of a primary flame and a secondary flame.Fil The primary flame is generated by the exhaust combustion gas from the rocket motor without any effect of the ambient atmosphere. The primary flame is composed of oblique shock waves and expansion waves as a result of interaction with the ambient pressure. The structure is dependent on the expansion ratio of the nozzle, as described in Appendix C. Therefore, no diffusional mixing with ambient air occurs in the primary flame. The secondary flame is generated by mixing of the exhaust gas from the nozzle with the ambient air. The dimensions of the secondary flame are dependent not only on the combustion gas expelled from the exhaust nozzle, but also on the expansion ratio of the nozzle. A nitropolymer propellant composed of nc(0-466), ng(0-369), dep(0104), ec(0 029), and pbst(0.032) is used as a reference propellant to determine the effect of plume suppression. The burning rate characteristics of the propellants are shown in Fig. 6-31. Since the nitropolymer propellant is fuel-rich, the exhaust gas forms a combustible gaseous mixture with the ambient air. This gaseous mixture is ignited and afterburning occurs somewhat downstream of the nozzle exit. The major combustion products in the combustion chamber are CO, Hj, CO2, N2, and HjO. The fuel components are CO and H2, the mole fractions of which at the nozzle throat are co(0.47) and iH2(0.24).

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