Supersonic air-breathing engines are the future of high-speed transportation vehicles. Recent studies have witnessed significant progress in the research on scramjets since it is the most suitable and promising engine for the air-breathing hypersonic propulsion [1
]: the greater specific impulse makes the scramjet appealing for civil and military applications [3
]. However, the issues employing supersonic combustion are still unsolved: fuel and oxidizer must mix, ignite, and burn in a very short time, i.e., in a reasonable length avoiding strong shocks within the engine [4
]. The crucial and complex task in the design of a scramjet is achieving proper and fast fuel–air mixing, good combustion efficiency, and optimum total pressure losses to make the energy provided by the combustor overcomes the drag forces experienced at high velocities. An effective fuel/air mixing and combustion are hard to achieve in scramjets since the residence time is of the order of milliseconds. Previous research efforts have mainly concentrated on injection techniques to maximize the mixing characteristics [5
]. To design an efficient fuel injector, mixing, combustion, as well as total pressure losses, must be investigated: in fact, there is always a trade-off between the penetration of fuel jet and losses in stagnation pressure since high injection angles lead to strong bow shock which enhances vorticity and mixing inside the flow but also strongly affects the total pressure los, Deep understanding is, therefore, necessary to answer these key issues.
Several injection strategies have come up with a special interest in generating streamlined vorticity. This is achieved by establishing pressure and density gradients, e.g., employing strategic injector design configurations, such as strut, wall, or swept ramp injectors, transverse, swirl, and cavities [7
], which allow the baroclinic waves to arise. A common and simple approach is normal injection through a wall orifice. Before the fuel jet, a bow shock arises due to the obstruction of the free stream by the propellant jet. A barrel shock occurs if the jet is under-expanded and will be terminated by a Mach disk. The high pressure of the fuel jet causes an unfavorable pressure gradient to exist within the boundary layer upstream of the injector causing a separation zone upstream of the injector [11
]: this region helps to stabilize the flame and achieve mixing; however, this is at the cost of significant total pressure loss.
On the other end, the effect of the bow shock could be reduced by lowering the injection angle (30 and 60 rather than 90 degrees) in a supersonic flow, which results in a reduction of total pressure losses. The axial momentum of the jet also helps to maximize the net thrust. Conclusions made by researchers [14
] in the case of angled injection stated that the auto-ignition and flames stabilization problems in the scramjet engines could be reduced especially below Mach 10. In recent years, cavities in supersonic combustion chambers have been proposed as a new concept to hold and stabilize the flame. Xing et al. [16
] examined the impact of cavity stream on the performance of scramjet combustors. Further studies [17
] showed that using a cavity after the injector significantly improves the combustion efficiency in a supersonic flow. The effect of sub-cavity in a supersonic cavity flow was studied by Panigrahi et al. [20
], who noted that the sub-cavity mitigates the shear layer reattachment at the aft wall of the cavity.
Injecting the fuel upstream of the cavity reports better penetration and mixing and distributes the fuel not only into the main stream but also the cavity. Eventually, most of the oxygen is consumed which improves the efficiency of the combustion [21
]. The cavity draws the fuel and air into it to improve the mixing by creating recirculation regions and by holding the products of reaction to stabilize the flame. Barnes et al. [22
] stated that injection from the cavity floor makes a more uniform mixing than injecting from the front wall, which made an extremely rich fuel mixture inside the cavity, over the upper combustibility constrain. In the study of Lin et al. [23
], it is observed that the distribution of the fuel into subsonic regions of the cavity depends on the injector location upstream of the cavity. From the study, Wang et al. [24
] stated that the pressure ratio of fuel to the mainstream plays a significant role in the stabilization modes of the flame in a cavity-based combustor with upstream injection.
The inclined injection upstream of the cavity reports lower total pressure losses compared to 90 degrees injection and consists of the flame inside of the cavity [25
]. The mixing of the fuel–air and the efficiency of the combustion is affected by certain factors in case of the injection upstream of the cavity like fuel jet parameters, fuel jet-cavity interaction, and penetration extent of fuel jet [27
]. Jeong et al. [30
] performed three hydrogen injection strategies in a cavity-based scramjet combustor. They reported that upstream injection enhances the fuel diffusion and shorter ignition delay compared to the parallel injection.
From the open literature, it is perceived that injection integrated with cavity encounters the critical characteristics of uniform mixing, flame holding, and stabilization mostly in 2D combustors. The works of literature related to the performance of circular combustor using axisymmetric cavities in a reacting supersonic flow field is scant. This paper investigates the effect of different injection strategies in an axisymmetric cavity-based circular scramjet combustor in a reacting flow field. Three injection schemes are incorporated within the cavity, and the performance is compared with the upstream injection. The performance of the scramjet combustor has been evaluated in terms of wall static pressure, hydrogen and water mass fractions, and combustion efficiency, as well as total pressure loss.