Juntao Chang

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

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.

Novel Oscillatory Patterns of Hypersonic Inlet Buzz

Unstart flows of a hypersonic inlet were experimentally studied at freestream Mach number of 4.5–6.0. With the aid of high-speed schlieren and time-accurate pressuremeasurements, the unsteady flow processes of the entire inlet, including the shock systemmotions and the surface pressure fluctuations, were recorded and discussed. The started flowfield analysis was conducted first, and then the unstarted flowfield was analyzed by using the unsteady pressure signals and schlieren pictures. Results indicate that two novel oscillatory patterns were observed in comparison with the past reported inlet buzz patterns. One is a mixed oscillatory pattern that mixes the “big buzz”with “little buzz,” and the other is a nonoscillatory violent pattern. These novel findings on oscillatory patterns of hypersonic inlet can provide more insight on inlet buzz mechanism, prediction, and control.

Effects of boundary-layers bleeding on unstart/restart characteristics of hypersonic inlets

A series of mixed-compression hypersonic inlets at different bleeding rates were simulated at different freestream conditions in this paper. The unstart/restart characteristics of hypersonic inlets were analysed and the reasons why the unstart/restart phenomenon is in existence is presented. The unstart/restart characteristics of hypersonic inlets at different bleeding rates were given. The effects of boundary-layer bleeding on the performance parameter (mass-captured coefficient, total-pressure recovery coefficient), starting and restarting Mach number of hypersonic inlets were discussed. In conclusion, boundary-layer bleeding can improve the performance parameter of hypersonic inlets, and can reduce the starting and restarting Mach number, and can broad the operation range of the hypersonic inlet.

Dimensionless analysis of the unstart boundary for 2D mixed hypersonic inlets

Inlet unstart boundary is one of the most important issues of the hypersonic inlet and is also the foundation of the protection control of a scramjet. To solve this problem, the 2D internal steady flow of a 2D mixed internal/external compression hypersonic inlet was numerically simulated at different freestream conditions and backpressures with a RANS (Reynolds-Averaged Navier-Stokes) solver using a RNG (Renormalisation Group) k -e turbulence model, and two different inlet unstart phenomena were analysed. The dimensional analysis method was introduced to find the essence variables describing the inlet unstart boundary based on “numerical experimental” data in this paper. The dimensionless pressure ratios of the forebody and isolator were analysed respectively. The results show that the unstart boundary of the 2D mixed hypersonic inlet is determined by M 0 , α and Re 0 . Pressure ratio π increases with M 0 increasing, and it increases firstly and decreases then with α increasing. Pressure ratio π increases with Re 0 increasing. Re 0 (Re 0 7 ) has a major effect on π and Re 0 (Re 0 > 2×10 7 ) has little effect on π.

Mechanism Study on Local Unstart of Hypersonic Inlet at High Mach Number

A viscous numerical simulation was performed to study the forebody shock reflection at the cowl under the condition of inflow Mach number higher than the design value for different forebody/cowl configurations. With the help of shock reflection theory and numerical simulation, it was found that the shock reflection changed from overall regular reflection wave configuration to overall Mach reflection wave configuration as the cowl angle decreased or the ramp angle increased. Mach reflection in this situation would be harmful; it could result in the shock propagating forward to cause a shock detachment at the cowl lip (the so-called local unstart of inlet in this paper). And the local unstart of inlet essentially resulted from the transition of forebody shock reflection from overall regular reflection to overall Mach reflection at the cowl. It was also found through comparison with inviscid analysis that the viscous effect had to be considered in the local unstart of inlet for practical applications. The ob...

Unstart Margin Characterization Method of Scramjet Considering Isolator–Combustor Interactions

Characterizing the unstart margin is very important for the hypersonic inlet to achieve maximum performance within safety range. A novel approach for characterizing the unstart margin when combustor–isolator interactions are taken into account is presented. First, experiments with four Mach numbers were conducted in a direct-connect supersonic isolator–combustor. The start-to-unstart process and the unstart-to-restart process were analyzed, and it is found that they should be discussed separately due to the hysteresis and different conditions of the isolator. A good way of indicating the unstart margin should be that the unstart margin maintains a linear relationship with the equivalence ratio. For the shock-train leading edge, it is concluded that, when combustion is considered, its behavior is rather complicated, which makes it inappropriate for characterizing the unstart margin. Then, the backpressure method is discussed, and its nonlinearity with equivalence ratio makes it not an ideal way to indicate...

A CFD assessment of classifications for hypersonic inlet start/unstart phenomena

Inlet start/unstart detection is one of the most important issues of hypersonic inlets and is also the foundation of protection controls of scramjets. In ground and flight tests, it is inevitably to introduce the sensor noises to the measurement system. How to overcome or weaken the influence of the sensor noises and the outer disturbances is an important issue to the control system of the engine. To solve this problem, the 2D inner steady flow of hypersonic inlets was numerically simulated in different freestream conditions and backpressures, and two different inlet unstart phenomena were analysed. The membership function for hypersonic inlet start/unstart can be obtained by using probabilistic output support vector machine, and the algorithm of multiple classifiers fusion is introduced. The variations of the classification accuracy with the intensity of the sensor noises and the number of the classifier were discussed respectively. In conclusion, it is useful to introduce the algorithm of support vector machine and multiple classifiers fusion to overcome or weaken the influence of the sensor noises on the classification accuracy of hypersonic inlet start/unstart. The number of the practical fusion classifiers needs a tradeoff between the fusion classification accuracy and the complexity of the classification system.

Combustion characteristic using O2-pilot strut in a liquid-kerosene-fueled strut-based dual-mode scramjet

A set of exploratory experiments are conducted to test a newly designed strut for fuel injection and flame holding in a liquid-kerosene-fueled dual-mode scramjet combustor. The thickness of the strut is 8 mm and the front blockage is about 8%. To organize stable combustion in a Mach number equal to 2.6 air flow under this thin strut using room-temperature liquid kerosene in a flash wall combustor without any cavity and other flame holders, some oxygen is injected through a set of orifices at the back of the strut, based on which a stable center local flame is generated at the back of the strut and the main flow combustion can be organized around this local flame. Experimental results show that stable combustion can be achieved at the center of the combustor with a wide range of equivalence ratio from 0.19 to 1 based on this center flame strut strategy. Through the analysis of the pressure distribution along the combustor, different combustor modes appear with different equivalence ratio. The article also gives some discussions about different influence of the oxygen to the combustion process under different equivalence ratio.

Multi-objective regulating and protecting control for ducted rocket using a bumpless transfer scheme

This article deals with the problem of multi-objective regulating and protecting control for a ducted rocket, in order to get maximum thrust while avoiding extremely dangerous phenomenon like inlet buzz. First, the mathematical models involving gas flow regulating system and ducted rocket are introduced. The description of inlet buzz margin is also given and analyzed. Second, the multi-objective switching control problem of ducted rocket is proposed and discussed. Third, a robust bumpless transfer scheme for solving the problem is presented. The designed bumpless transfer compensator based on model reference adaptive sliding-mode control ensures that the switched system performs a smooth transition at the transfer moment. The stability analysis is then given utilizing the Lyapunov functional approach. Finally, the bumpless transfer strategy is applied to the ducted rocket control system, and the simulation results show its effectiveness.

Combustion characteristics of a dual-mode scramjet injecting liquid kerosene by multiple struts

In order to optimize fuel allocation between multiple injection struts in scramjet combustor, experimental investigation has been carried out to study the combustion characteristics of a dual-mode scramjet, in which liquid kerosene was injected from three staggered arranged struts, and the effect of fuel allocation between the three struts on the combustion characteristics has been analyzed. The results show strong influence of the quantity and streamwise position of struts on the wall pressure distribution and both the equivalence ratio limits of inlet unstarting and lean blowout. To obtain a wide stable operation range, just one strut is needed for injecting most of the fuel in the combustors with configuration and size as in this work, and a streamwise position closer to the downstream divergent segment is a better injection location. Lower lean blowout limit can be obtained by injecting the overall much lean fuel from just one strut with a relatively closer distance to the downstream flameholder.