Dynamic Characteristics of Combustion Mode Transitions in a Strut-Based Scramjet Combustor Model

Abstract

T HE supersonic combustion ramjet (scramjet) engine is expected to be the most efficient propulsion system in the hypersonic flight regime [1]. Given the broad range of aerothermodynamic conditions experienced during hypersonic flight, the scramjet would operate under different combustionmodes [2], andmode transition is a critical phenomenon in designing such engines because the thrust and specific impulse of the fuel in each mode varies considerably. In much of the previous work, researchers experimentally achievedmode transition and investigated the static characteristics of different combustion modes. In the open literature, Billig [3] first demonstrated mode transition in ground tests. Heiser and Pratt [4] used a one-dimensional (1-D) analysis approach to comprehend the complex aerothermodynamics of a dual-mode combustion system. The flowfield can be illustrated for threemodes: scramjet with shockfree isolator and oblique shock train, and ramjet with normal shock train. Takahashi et al. [5] and Kouchi et al. [6] observed four different combustion modes with respect to the fuel flow rate, namely, blowout,weak combustion, strong combustion, and thermal choking. As the mode transition occurred, thrust and heat-flux distribution [7] varies considerably. Sullins [8] experimentally achieved the mode transition from a scramjet with a precombustion shock system having a high pressure ratio to a scramjet with no precombustion shock system by increasing the total temperature of airflow to simulate a real acceleration process. Micka and Driscoll [9] reported two distinct reaction zones in a combustor with wall injection and a cavity flameholder corresponding to jet wake stabilization and cavity stabilization. The reaction zonewas found to only appear in the cavity stabilized mode in the scramjet mode, even for conditions where the ramjet modewas jet-wake stabilized. Also, the spreading of scramjet mode combustion is significantly less than that of the ramjet mode. Masumoto et al. [10] investigated the effect of combustor length and total temperature on combustion modes and suggested the minimum combustor length to attain supersonic or dual-mode combustion. However, there have been few studies on the dynamic characteristics of combustion mode transition, and the open literature did not fully investigate the combustor performance changes with the fuel flow rate small changes (∼1 g∕s) near the critical conditions. One interesting phenomenon, rather different from the results available in the open literature, is that the wall pressure and thrust show obvious catastrophe near the critical point of combustion mode transition. The combustion mode transition depends on the path taken (i.e. the fuel flow rate is increasedordecreased).With the sameexternal parameters, the scramjet engine may be a different combustion mode, known as hysteresis effect according to the nonlinear dynamics theory. During hypersonic flight, itmay bringgreat difficulties to the precise control of the vehicle, have a great impact on the flight safety, and even cause a flight accident [11]. Therefore, the successful development of a scramjet engine depends on further understanding and control of the nonlinear mode transition process. In this research, particular attention was focused on the dynamic characteristics of combustion mode through ground tests, especially the nonlinear catastrophic and hysteresis phenomena. As known, the transition between ramjet and scramjet mode is determined from the magnitude of ΔT0∕T0;air (either by decreasing or increasing the amount of heat release). In this paper, we linearly changed the fuel mass flow rate along two adverse paths; that is, increased and decreased the fuel equivalence ratiowhile the rate of changewas held approximately constant. In particular, to obtain performance of the model combustor around the critical conditions in detail, the heat release was changed little by little every time (corresponding to an increase in fuel equivalence ratio of 0.0125). Compared to strut injection in the center of the combustor, the transverse wall injection disturbs the boundary layer significantly. The wall injection plume forms a barrel shock, and induces a bow shock that leads to separation and the formation of a recirculation region in front of the wall injection location. These unnecessary disturbances make it difficult to determine the exact mode transition mechanism [12]. Therefore, a central strut injector has been employed, which also improved fuel mixing in the supersonic core stream and combustion performance in supersonic combustors. Because the liquid hydrocarbon fuel has greater fuel densities and endothermic cooling capabilities than hydrogen, particularly for hypersonic vehicles limited to Mach 8 flight, kerosene was used as the fuel in this research.