Oblique shock

About: Oblique shock is a research topic. Over the lifetime, 6551 publications have been published within this topic receiving 119823 citations.

Particle Image Velocimetry in Mach 3.5 and 4.5 Shock-Tunnel Flows

Abstract: For the e rst time, a particle image velocimetry (PIV) system was used to study high-Mach-number e ows in a shock-tunnel facility with velocities of more than 1.5 km/s and measuring times in the millisecond range. An application of PIV to such a transient high-speed e ow is considerably more dife cult than to a continuous e ow because no online adjustments of the optics and the particle seeding can be done. Additionally, a proper seeding and timingofthefacility iscrucial.Firstwewilldiscussthemeasured velocity e eldbehind acontoured Lavalnozzle (design Mach number4.5 ). The measurement data show that thee owe eld at the nozzleexit is parallel to the nozzle axis and homogeneous as expected from supersonic nozzle theory. The average measured velocity corresponds very well to the calculated e ow velocity. The results are compared to measurements made with a conical Mach 3.5 nozzle that exhibits a diverging e owe eld. A wedge was further introduced into the parallel Mach 4.5 nozzle e ow to study the seed particle performance downstream of an oblique shock. The measured results are also in good agreement with calculated velocities from oblique shock theory. PIV has, therefore, proven to be an efe cient measurement method for high-speed and short-duration simulation facilities.

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.

Shock front instability associated with reflected ions at the perpendicular shock

Abstract: Two-dimensional hybrid simulations of perpendicular, supercritical collisionless shocks are carried out in a geometry with the magnetic field perpendicular to the simulation plane so that parallel propagating fluctuations, such as Alfven ion cyclotron waves, are suppressed. In terms of average profile and large downstream ion temperature anisotropy, the results resemble those from earlier one-dimensional hybrid simulations, and differ markedly from the results of two-dimensional simulations in which field-parallel propagating fluctuations are included. In addition, we find an instability at the shock front, in which a pattern of magnetic field and density enhancements propagates along the shock surface in the direction of gyration and at the average speed of the ions reflected at the shock. The instability mechanism depends on a spatio-temporal modulation of the fraction of reflected ions over the shock surface. The instability has a threshold that depends on the Mach number and the upstream ion plasma beta, being stabilized by an increased beta and decreased Mach number. In a realistic three-dimensional planar shock, this instability will be only one of several mechanisms contributing to shock front nonstationarity. However, at a three-dimensional curved shock, there is a region where the instability mechanism described may dominate.