Mirko Gamba

Ignition and Near-Wall Burning in Transverse Hydrogen Jets in Supersonic Crossflow

The present experimental work investigates flame structure and near-wall ignition observed in transverse hydrogen jets in supersonic crossflows. An underexpanded sonic hydrogen jet issued from a flat plate perpendicularly to an incoming high-enthalpy crossflow is considered. The cross-flow approximates aerothermal conditions of Mach 8 flight at 30 km altitude. Transverse jets at different jet-to-crossflow momentum flux ratios J in the range 0.3 – 5.0 have been considered. The freestream (crossflow) conditions are maintained at a Mach number of 2.4 with a static temperature and pressure of 1400 K and 40 kPa, respectively. Experiments are carried out at the Stanford Expansion Tube Facility. Imaging techniques, such as Schlieren and OH* chemiluminescence imaging, are applied to characterize the global features of the system, such as ignition and flame penetration. The identification and investigation of the instantaneous reaction zone structure is carried out with planar laser-induced fluorescence imaging of the hydroxyl radical. Several orthogonal imaging planes have been considered to map the spatial features of the reacting regions. The importance of the unsteady bow shock, separation and recirculation regions on near-wall ignition, combustion and transport processes at sufficiently large values of J are identified and discussed.

Supersonic Combustion of Pylon-Injected Hydrogen in High-Enthalpy Flow with Imposed Vortex Dynamics

The ignition and combustion characteristics of the hydrogen plume issued from two pylon-type injectors in a Mach 2.4, high-enthalpy airflow are presented. Specifically, the focus of the study is on the effects of the imposed interaction and subsequent dynamics of a system of selected supersonic streamwise vortices on the reacting plume morphology and its evolution. The design phase of the experimental campaign was carried out with a reduced-order model, with the goal of identifying peculiar interactions among streamwise vortical structures introduced in the flow of interest. Two vortex interaction modes have been selected and later implemented using ramp-type vortex generators positioned symmetrically and asymmetrically on the pylon injectors, as prescribed by the results of the simulations reported here. Hydrogen/air combustion experiments, aimed at investigating the selected cases, were conducted in the expansion tube facility of the High-Temperature Gas Dynamics Laboratory at Stanford University. Stagn...

Volumetric PIV and 2D OH PLIF imaging in the far-field of a low Reynolds number nonpremixed jet flame

Cinematographic stereoscopic PIV with temporal and spatial resolution ranging from 2.6 to 5.5 Kolmogorov scales, which is sufficient to accurately represent most of the dissipation structures, is used in conjunction with Taylor’s frozen flow hypothesis to generate quasi-instantaneous pseudo-volumes of the three-component velocity field in the far-field of a nonpremixed jet flame at the jet exit Reynolds number (Red) of 8000. The 3D data enable the computation of the nine components of the velocity gradient tensor and other important kinematic quantities. The volumetric PIV is combined with single-shot simultaneous OH PLIF imaging to mark the instantaneous reaction zone at one plane in the reconstructed volume. The combined datasets enable the investigation of the relationship between the reaction zone and the fully-3D representations of strain, vorticity, kinetic energy dissipation and dilatation, and of the impact of heat release on the structure of turbulence. In this Red = 8000 flame, it is observed that sheet-like layers of vorticity and dissipation tend to coincide and are aligned with the OH layers, an effect that is believed to be due to the stabilizing effect of heat release on this relatively low Reynolds number jet flame. Furthermore, the spatial organization of the strain field is predominantly driven by the presence of the flame rather than turbulence. Finally, intense dissipation is mostly due to the laminar shear caused by the presence of the flame rather than to the strain generated by vortical structures as typically observed in nonreacting jets.

Secondary Diaphragm Thickness Effects and Improved Pressure Measurements in an Expansion Tube

T HE aim of this Technical Note is to present practical considerations for measuring pressure in an expansion-tube flow facility; some aspects of this work can be applied to pressure measurement strategies in other impulse (e.g., reflected shock tunnel) or continuous flow facilities that may suffer from similar issues. We assume that the reader is familiar with the terminology and general operating principles regarding expansion tubes; formore information regarding expansion tubes, please refer to the seminal work of Trimpi [1]. In many studies relevant to aerospace engineering, propulsion, and fluid mechanics, pressure measurement is often a diagnostic of primary interest [2–7], but obtaining high-quality (i.e., high signalto-noise, or SNR) pressure measurements can be a challenge in expansion-tube facilities (or other impulse facilities). Because of the short test times in impulse facilities, high-bandwidth pressure transducers are necessary to capture transients of interest, but as we show in this work, certain transducer architectures are more susceptible to noise than others. Analog or digital signal processing can be used to filter some noise from the measurements, but at the expense of measurement bandwidth or accuracy; thus, the engineer must compromise between bandwidth and SNR to acquire reliable pressure measurements. Practical information regarding the acquisition of optimal (i.e., high SNR) pressure measurements in impulse facilities is sparse; however, some articles do provide useful details of implementation, for example Dufrene et al. [8] present specifics of pressure instrumentation in an expansion-tube facility, and Beresh et al. [9] report on the response characteristics of a variety of pressure transducers. Expansion-tube facilities also produce a contaminated freestream flow due to the presence of the secondary diaphragm; the primary shock wave may reflect off the secondary diaphragm [10], and secondary diaphragm fragments of a broad range of sizes are entrained in the test gas [11]. We have found that diaphragm fragments impacting the structure in which pressure transducers are mounted can have a significant, adverse effect on pressure measurement quality. An example of this effect is presented in Fig. 1, which shows a pressure time history alongwith a schlieren image taken at the instant in time when a diaphragm fragment impacts the model that houses the pressure transducers. The shadowed shape in the schlieren image is a flat platewith a pitot-pressure transducer (PCB113A26)mounted above the plate in a conical housing; directly below the pitot opening, a wall-mounted transducer (PCB 112A22) simultaneously measures the wall static pressure. (We will interpret these pressure traces in Sec. III.A, but for now, we only demonstrate the level of noise in the two traces.) The displayed still schlieren image is taken from a high-speed schlieren video that captures the full event during a test. A diaphragm fragment can be seen impacting the top of the model, at which time both pressure traces become extremely noisy. We have also observed the spontaneous eruption of noise in pressure time histories in many other tests, and generally, the time at which the pressure traces become noisy varies from shot to shot, and the onset of noise never occurs before the arrival of the test gas. These introductory observations suggest that diaphragm fragments impacting the model are a major source of noise in these pressure measurements. To test the hypothesis that diaphragm fragment impacts are a source of noise in pressure measurements, we make pressure measurements as a function of secondary diaphragm thickness, expecting thinner diaphragms (and less massive fragments) to cause less noise. Two types of pressure transducers are also tested, two different piezoelectric sensors and two different piezoresistive sensors, and the noise characteristics of each sensor are compared. As a secondary objective, we also test whether the secondary diaphragm thickness affects test-gas conditions as inferred from shock speed measurements. Last, a practical solution for minimizing noise in piezoelectric transducers is provided.

Supersonic Combustion and Flame-Holding Characteristics of Pylon Injected Hydrogen in a Mach 2.4 High Enthalpy Flow

This work investigates the structure of the reaction zones downstream of a pylon injecting hydrogen fuel into a Mach 2.4 high enthalpy flow, with the aim of characterizing the behavior and evolution of the expected vortical structures. The pylon injector incorporates expansion ramps that serve as vortex generators. Fuel is injected through a thin slit along the entire length of the base of the ramps. Tests were conducted in the Stanford University Expansion Tube Facility at the High Temperature Gasdynamics Laboratory. All the experiments were conducted at a stagnation enthalpy of 2.8MJ/kg, a static temperature of 1400K, a static pressure of 40kPa, and Mach number of 2.4. The total temperature of the injected hydrogen was 300K. The ignition and flame holding characteristics were investigated using schlieren, time-integrated OH* chemiluminescence, and instantaneous OH planar laser-induced fluorescence (OH PLIF) imaging. OH PLIF was applied to capture the evolution of the reactive vortical system in planes normal to the freestream flow direction at a distance of 1.8cm, 4.3cm, 7.6cm and 10.7cm from the fuel exit plane. Resulting images show the distribution of OH radicals in the plume; these images also show a peculiar geometric pattern in the plume shape, suggesting that the vortices generated by the pylon configuration, and the consequent vortex dynamics, played a dominant role in the mixing and combustion processes.

Unsteadiness characteristics and three-dimensional leading shock structure of a Mach 2.0 shock train

The structure and unsteadiness characteristics of a shock train in a Mach 2.0 ducted flow are studied. The shock train is generated and stabilized by a back-valve that imposes a desired back pressure. Despite nominally constant boundary conditions (i.e., inflow and outflow), the shock train fluctuates about its time-average position. High-speed schlieren imaging is used to quantify the amount of unsteadiness. The results show that the shock fluctuation speed and the magnitude of shock displacement are independent of back pressure. However, the angles of the leading shock lambda foot and the length of the leading shock Mach stem change as back pressure is increased indicating the shock structure transitions from oblique to normal as the shock train moves upstream. The structural changes of the shock train are reflected in time-averaged pressure profiles. As the shock train moves upstream the length of the shock train decreases and the pressure rise across the shock train increases. Finally, stereo particle image velocimetry is used to study the structure of the leading shock in a normal shock train (i.e., when the back pressure is high and pushes the shocks upstream). The velocity fields show a large separation region on the side-wall of the duct and a degree of axisymmetry that indicates a nominally conical shape. The overall structure of the leading shock is modeled as two truncated cones with the small ends coinciding to form the Mach stem.

Periodic forcing of a shock train in Mach 2.0 flow

High-speed schlieren movies and pressure measurements are collected to analyze the response of a shock train due to downstream forcing. The shock train is generated in a Mach 2.0 ducted flow and controlled by a downstream butterfly valve. Cyclic opening and closing of the valve (at rates up to 10 Hz) leads to oscillations in back pressure measured at the end of the duct. Subsequently, the shock train oscillates between two locations in the duct, traveling at speeds up to 3.5 m/s. Different cases with varied forcing frequency and magnitude of back pressure change are studied. For each case we evaluate the response of the back pressure and shock location by quantifying the magnitude of change, rise time, delay time, and maximum rate of change. Some of the key results include: 1) there is a linear relationship between the magnitude of the shock displacement and the back pressure change that is independent of the forcing parameters; 2) the leading shock in the shock train responds to forcing ≈6 ms after the back pressure starts to change indicating an upstream propagating disturbance; 3) the rise time of the back pressure and leading shock location are approximately the same.

Relationship between intermittent separation and vortex structure in a three-dimensional shock/boundary-layer interaction

The relationship between the three-dimensional vortex structures and flow-separation zones generated by a shock wave/boundary-layer interaction within a low-aspect-ratio duct was studied using stereoscopic particle imaging velocimetry measurements. In this configuration, the interaction of the incident shock with all walls was important in controlling the flowfield; the three interactions coupled to produce a strongly distorted flowfield. Conditional sampling was used to construct the local probability of reverse flow maps, and thus quantify the distribution of regions of intermittent separation on both bottom-walls and side-walls. The latter regions were found to be significantly larger and more likely to separate than the former. Thus, it was concluded that the sidewall and corner flow interactions dominate in this configuration. A triple decomposition of motion was used to construct a three-dimensional representation of the vortex features generated by the interaction. The results indicated that the fl...

Design and characterization of the Michigan hypersonic expansion tube facility (MHExT)

The development of the new hypersonic impulse facility at the University of Michigan capable of generating aerothermal flow conditions representative of flight Mach numbers ranging between 4 and 11 is discussed. Three conditions were designed and run extensively for the purposes of validating and characterizing the generated flows. The conditions were chosen to span a range of accessible flow properties while maintaining relevance to the studies of supersonic combustion. The performance of the facility was assessed based on its ability to repeatedly and accurately generate flow conditions of interest. The temporal and spatial uniformity of the flow were determined using simultaneous pitotstatic pressure measurements, pressure measurements with a pitot rack, and schlieren imaging. Flow uniformity, useful test time, and core-flow sizes are established. A shot-to-shot statistical analysis of the resulting shock speeds reveal their variability to amount to a fraction of a percent, attributing to the ability of the facility to consistently generate the desired flow conditions.

Experimental study of supersonic turbulent corner flow evolution in a low aspect ratio rectangular channel

Supersonic turbulent flows in a rectangular duct are characterized by secondary flows in each of the four corners of the duct. These secondary flows comprise of two counter rotating vortices forcing the fluid towards the corner. The vortices are developed as a result of the Reynolds stress gradients which exist in each of the corners. Previous studies 1,2 have primarily used intrusive techniques such as Preston/pitot tubes and hot wire in order to study the secondary flows in the corners. However use of these techniques would substantially modify the flow in low aspect ratio rectangular channels. Thus, there is a need for non-intrusive measurements to study the evolution of corner vortices in such flows. In this study we present the experimental results of stereo particle image velocimetry conducted on a Mach 2.75 flow in a low aspect ratio rectangular channel. Data is recorded at multiple cross-section planes and analyzed to study the formation of corner vortices and their effects on the structure of the mean flow field. Results show that the corner vortices are formed in the corner and then convect away from it. There is a significantly strong secondary flow associated with the corner vortices possibly causing open flow separation. The vortices significantly modify the local variation of the skin friction coefficient.