13th Brazilian Congress of Thermal Sciences and Engineering December 05-10, 2010, Uberlandia, MG, Brazil
SINGLE-SHOT PULSED DETONATION DEVICE FOR PDE COMBUSTION SIMULATION
Carla S. T. Marques, email@example.com Antonio C. de Oliveira, firstname.lastname@example.org
Aerothermodynamic and Hypersonic Division, Institute of Advanced Studies – Departmentof Aerospatial Science and Technology, Rodovia dos Tamoios, km 5.5, 12228-001 São José dos Campos – SP, Brazil
Fernando B. Dovichi Filho William C. Ferraz
Aerothermodynamic and Hypersonic Division, Institute of Advanced Studies – Department of Aerospatial Science and Technology, Rodovia dos Tamoios, km 5.5, 12228-001 São José dos Campos – SP, Brazil
José B. Chanes Jr., email@example.comAerothermodynamic and Hypersonic Division, Institute of Advanced Studies – Department of Aerospatial Science and Technology, Rodovia dos Tamoios, km 5.5, 12228-001 São José dos Campos – SP, Brazil
Abstract. In this work, a single-shot pulsed detonation device for experimental simulation of the PDE combustion conditions in supersonic and hypersonic flight regimes (M∞ ≤ 6) is proposed. Thecombustion device is composed of: ignition system, detonation tube without obstacles, divergent nozzle and test chamber. A nanosecond spark discharge has been developed to promote a direct detonation initiation or a DDT in the shortest distance possible. Pulses of approximately 25, 65 and 80 ns (FHWM) were acquired. Emission and Schlieren images were taken to characterize the spark discharge. The pulseddetonation device was planned to achieve typical specific impulse (Isp ≈ 200s) for a single-shot PDE. 1-D calculations were performed to establish the combustion conditions for experimental tests. Flight conditions between 0 and 20,000 m of altitude fueled with H2/air and C2H4/air mixtures were calculated within 0.5-1.5 atm range of initial pressure at different equivalence ratios for two nozzlearea ratios. The nozzle exit velocities are notably higher if mean pressure of PDE cycle (Pmean) is considered. Assuming mean pressures of PDE cycle, experimental conditions with Mach number (Mx) from 3.0 to 3.7 can be simulated for these explosive mixtures. The results show PDE flight regimes from transonic to supersonic. However, it is promising to reach higher pressures and velocities throughthe ignition system proposed and, consequently hypersonic flight regimes. Keywords: nanosecond discharge, Thyratron, detonation, PDE, scramjet. 1. INTRODUCTION Pulsed detonation engines (PDEs), in recent decades, have been developed due to their high potential application as an aerospace propulsion system across subsonic, supersonic and hypersonic flight regimes (Tangilara et al, 2005). There arewide applications for these devices, but the most promising is the scramjet development (Povinelli, 2002; Falepim et al, 2001; Cambier et al, 1995), which will allow easy access to space. The thermodynamic advantage of detonations in front of deflagrations processes, which provide high specific impulse in a broad range of Mach number, the reduced complexity and the lower operational cost becamePDEs very attractive (Kailasanath, 2003) and nowadays they are a real possibility as aircraft engines (FlightGlobal, 2008). However, reliable and repeatable detonations in the shortest distance possible are required to achieve a practical PDE; and the key to make this feasible is the detonation initiation (Lee et al, 2005; Zhukov et al, 2006). Detonations can be initiated through two different ways:a slow mode where there is a deflagration-to-detonation transition (DDT) and a fast mode generated by a powerful ignition or a strong shock wave. Sometimes, they are designated as auto-ignition (or thermal initiation) and direct ignition, respectively (Marques, 1996). Direct initiation of hydrocarbon propellant detonation demands high-energy input of kilojoules magnitude (Lee et al, 2000; Zhukov...