High Enthalpy Supersonic Flow Research Articles

Dynamic response characteristics of flame during condition transition in a cavity-based scramjet combustor

Experimental study is performed to investigate the dynamic response characteristics of the flame when global equivalence ratio increases suddenly during the condition transition in a cavity-based scramjet combustor. High-enthalpy supersonic flow enters the combustor on conditions that Mach number, stagnation temperature, and total pressure are 2.52, 1486 K, and 1.6 MPa, respectively. In this paper, more attention is paid to the characteristics of flame evolution after sudden increase of fuel injection pressure and combustion characteristic when the injection pressure gets stabilized again. Three flame modes after condition transition have been observed and analyzed according to the transient chemiluminescence image captured by high-speed photography. There are two kinds of stable flame modes after transition mission at a low injection pressure difference, which are mainly reflected in the distribution of flame and intensity of visible light. Whereas, once the injection pressure difference exceeds a critical value, intense periodic flame flashback phenomenon begins to appear in the combustor. It can clearly capture the back and forth motion of the flame in the combustor, along with the local blowout and reignition phenomenon. Furthermore, the dominant frequency of flame flashback decreases and contrast of combustion intensity becomes more obvious as the injection pressure difference increases.

Combustion Oscillations in Scramjet Combustor with Different Fuel Injection Schemes

This work experimentally studied combustion oscillations in a cavity-based combustor operating under high-enthalpy supersonic flow. Three fuel-injection distances and the corresponding combined fuel injection schemes were tested and compared with the global equivalence ratio unchanged. The results indicate that the combustions of longer and medium distances (35 and 21 mm, respectively) present stronger oscillations, which are mainly attributed to the strengthened oscillations of the mainstream flame (MF). It is also found that combined fueling schemes can help weaken the oscillations. The predominant frequencies of the combustion oscillations in different fueling schemes are all distributed at approximately 400 Hz, which are proven to be induced by the MF. The low-frequency range (0–50 Hz) is more dominant for the shorter injection distance (7 mm) than in the other cases, which is because the flames are strongly influenced by low-frequency disturbances with fluctuations in the fueling injection and inflow. The analysis reveals that these disturbances act primarily on the cavity flame.

Behaviors of the reacting flowfield during the spontaneous formation of ramjet mode under a supersonic inflow

This work experimentally studied the formation process of the ramjet mode occurring in a cavity-based combustor operating at a high-enthalpy supersonic flow. The ramjet mode is featured by the phenomenon that the incoming supersonic inflow is decelerated to be subsonic before it enters the combustor, which is caused by the strong heat release under a high equivalence ratio. In the experiments, the ignition is performed after a steady fuel mass flow rate has been achieved. According to the flame behavior and the flowfield structure, the formation process of the ramjet mode can be divided into three stages, among which stage 1 (from ignition to the cavity shear-layer mode) is shortest, while stage 3 (from the lifted shear-layer mode to the ramjet mode) consumes the longest time. In stage 2, flashback occurs and shock–shock interactions are found to be strongly coupled with the local combustion which have an influence on the propagation velocity of the backpressure. A thickening boundary layer upstream of the separation shock is observed when the separation shock has interwoven with the jet-induced bow-shock. The thickening process could be extremely short (in 100 μs) before the thickened boundary layer separates, during which the propagation velocity of the backpressure can be apparently decelerated. The same phenomena shown in the supplementary experiments confirm that the thickening boundary layer and its deceleration effect on the propagation of the backpressure are not accidental but more likely to be inherent to the flashback occurring under a supersonic flow.

A Mixed Radiative-Convective Technique for the Calibration of Heat Flux Sensors in Hypersonic Flow

The ability to measure the very high heat fluxes that typically occur during the hypersonic re-entry phase of space vehicles is generally considered a subject of great importance in the aerospace field. Most of the sensors used for these measurements need to be checked periodically and re-calibrated accordingly. Another bottleneck relates to the need to procure thermal sources that are able to generate reliable reference heat fluxes in the range between 100 and 1000 kW/m2 (as order of magnitude). In the present study, a method is presented by which, starting from a calibration system with a capacity of approximately 500 kW/m2 only, heat fluxes in the range of interest for hypersonic applications are generated. The related procedure takes advantage of established standards for the characterization of a radiative heat flux. It also builds on the hybrid radiative-convective nature of typical hypersonic heat fluxes and the yet poorly explored possibility to use convective sources of heat to produce high-intensity fluxes. The reliability of such a strategy has been tested using a high enthalpy supersonic flow facility relying on an electric arc-heater and pure Nitrogen as work gas. Stagnation-point heat fluxes have been successfully measured (with reasonable accuracy) in the range between 600 and 1500 kW/m2 for values of the centerline enthalpy spanning the interval from to 6 to 24 MJ/kg.

Accuracy of Direct Skin-Friction Measurements in High-Enthalpy Supersonic Flows

Measurement of skin friction and a survey of airflow momentum were conducted in a rectangular straight duct with airflow Mach number, unit Reynolds number, and stagnation enthalpy at the nozzle exit of 2.5, 2.7 x 10 7 m -1 , and 1.1 MJ/kg, respectively. Skin fiction was measured by a direct skin-friction sensor at the test section that was connected to the duct exit, with the state of airflow there also being surveyed. From the result of the measurement and survey at two streamwise locations with different lengths, mean skin friction was calculated by two different and independent methods. The mean skin friction calculated by the directly measured skin friction was compared with that calculated by the momentum conservation law in order to examine the accuracy of direct skin-friction measurement. The accuracy of skin-friction measurement using the direct skin-friction sensor in the test was found to be ±5% (root-sum-square value).

Dynamic oxidation of ultra-high temperature ZrB 2–SiC under high enthalpy supersonic flows

The dynamic response to oxidation of a hot-pressed ZrB 2 + 15 vol%. SiC ceramic composite was studied under aero-thermal heating using a high enthalpy supersonic flow of a N 2/O 2 gas mixture in plasma wind tunnel. Microstructural features of the reaction scale developed upon oxidation were analyzed and correlated to test conditions through Computational Fluid Dynamics simulations. The significant heat flux and temperature gradients of the sample’s surface exposed to the highly energized N 2/O 2 gas stream led to the formation and evolution of distinct layered oxide sub-scales. The diffusion of oxidants through silica-rich 3D glassy network was proposed as the rate governing factor for oxidation.

The application of coherent anti-Stokes Raman scattering to temperature measurements in a pulsed high enthalpy supersonic flow

Broadband single pulse coherent anti-Stokes Raman scattering (CARS) experiments employing a folded box phase matching geometry in a shock tunnel flow are presented. Rovibrational spectra of molecular nitrogen, produced at the exit of a pulsed supersonic nozzle for a range of flow enthalpies, are examined. Difficulties peculiar to the application of the optical technique to a high enthalpy pulsed flow facility are discussed and measurements of flow temperatures are presented. Theoretically calculated values for temperatures based upon algorithms used to determine shock tunnel flow conditions agree well with experimental measurements using the CARS technique.

高速作動バルブ方式衝撃波風洞および高速マスサンプリングプローブを用いた超音速気体混合過程に関する実験

In order to examine the supersonic mixing process in a high enthalpy supersonic flow, the shock tunnel using the quick-action valves has been developed and confirmed their operation. The present shock tunnel enables to make a high enthalpy flow without contamination in high reproducibility by easy operation. The nozzle is designed to make parallel supersonic flows which are comprised of high-enthalpy air stream and helium stream. Mach number of air and helium flow are 3.4 and 4.2, respectively. The structure of mixing region of parallel supersonic flows thus obtained was examined by means of the mass-sampling probe which enables one to examine directly the concentration ration of the mixture gases. The measurement shows that the mixing region almost agrees with the region observed by Schlieren photograph.

Laminar heat transfer to a two-dimensional backward facing step from the high-enthalpy supersonic flow in the shock tube

Laminar heat transfer to a two-dimensional backward facing step from the high-enthalpy supersonic flow in the shock tube