Influência da razão de equivalência na combustão e no empuxo gerado por scramjet em voo atmosférico a Mach 5,8 e 20 km de altitude
Researches in scramjet technology (supersonic combustion ramjet) have demonstrated the viability of this propulsive system in promoting access flights to space. The non-necessity to carry the oxidizer on board and the hypersonic flight capability are outstanding features. Scramjets also exhibit grea...
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| Formato: | Dissertação |
| Idioma: | pt_BR |
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Universidade Federal do Rio Grande do Norte
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| Endereço do item: | https://repositorio.ufrn.br/handle/123456789/45135 |
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| Resumo: | Researches in scramjet technology (supersonic combustion ramjet) have demonstrated the viability of this propulsive system in promoting access flights to space. The non-necessity to carry the oxidizer on board and the hypersonic flight capability are outstanding features. Scramjets also exhibit greater specific impulse when compared to rockets for speeds corresponding up to Mach 15. In the present work, the thrust generation capacity of a scramjet integrated aerospace vehicle was evaluated from the combustion of hydrogen and atmospheric air at supersonic speed, for the flight of the scramjet vehicle at 20 km of altitude and speed of 1709.6 m/s, corresponding to Mach number of 5.79. A Theoretical-analytical method was used, considering the steady-state. The First Law of Thermodynamics, without and with reaction based on enthalpy of formation, was used to determine the thermodynamics properties of flow. Correlations between the equivalence ratio with the mixture temperature, mixture speed, and exhausted gases temperature after combustion were obtained. Two cases for combustion were considered: the case at constant pressure and the case with the constant cross-sectional area for the combustor. The temperature and velocity after the expansion process of the gases from the combustion were also determined for each of the cases. A mixture velocity of 1152 m/s was assumed for both cases, resulting, for the constant pressure case, in an equivalence ratio in the range of 0.648 to 0.774 and a mixture temperature of 876 K to 856 K by prefixing the combustion temperature ranging from 2400 K to 2600 K. Considering this range of equivalence ratio, we obtained average specific impulse equal to 4049.7 s and positive specific thrust varying between 760.8 and 879.3 N/ kg air. For the case with the constant area, the equivalence ratio was limited to the maximum value of 0.139 due to the thermal throttling effect, obtaining maximum specific thrust and specific impulse equal to 195.5 N/ kg air and 4914 s, respectively. The two cases were compared considering equivalence ratio range varying from 0 to 0.139, obtaining specific thrust and specific impulse about 3.3 % higher than the case with the constant area with the disadvantage that the temperature reached value 19.2 % higher than the case with constant pressure combustion. It was concluded that the scramjet can generate thrust at the established flight conditions for both cases, provided that for the case with the constant area the maximum limit for the equivalence ratio is respected. |
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