Ivett A. Leyva

Dark Core Analysis of Coaxial Injectors at Sub-, Near-, and Supercritical Conditions in a Transverse Acoustic Field

An experimental study on the effects of an externally-imposed transverse acoustic field in a N2 shear coaxial jet at sub-, near-, and supercritical pressures is presented. Such fields and their interaction with the jet (i.e., breakup, mixing, etc.) are believed to play a critical role during combustion instabilities in liquid rocket engines. The shear coaxial injector used here is similar to those used in cryogenic liquid rockets. By using N2 as the working fluid, the chemistry effects on combustion instability are separated from the effects of a transverse acoustic field on coaxial jets. Furthermore, through this choice, ambiguities associated with composition dependence on mixture critical properties are eliminated. The acoustic field is generated by a piezo-siren and the resonant frequency studied is ~3kHz. The pressures in the chamber range from 1.5-4.9 MPa to span subcritical to supercritical pressures. The outer to inner jet velocity ratio varies from ~1.2 to 23 and the momentum flux ratio (MR) varies from ~0.2 to 23. These ratios are mainly varied by changing the temperature and flow rates of the outer jet. At least 2000 backlit images were taken at 41kHz for each run. The main metric investigated is the length of the dark, or inner jet, core. This length is related to the mixing processes in a coaxial jet. The shorter the core length the faster the mixing occurs. Both the axial and the total, or curved, dark core lengths are studied. For momentum flux ratios ~1<MR<~4 the differences in the axial and curved dark core lengths between acoustics on and off are statistically significant, which means acoustics do shorten the core for this range. For subcritical pressures the MR range where the jet is shortened is larger. Preliminary results on the frequency analysis of the dark core lengths and width is also presented.

Experimental and computational study of high enthalpy double-wedge flows

A series of experiments designed to study reacting nitrogen flow over double-wedge geometries was conducted in the T5 shock tunnel at the California Institute of Technology. These experiments were designed using computational fluid dynamics to test nonequilibrium chemistry models. Surface heat transfer rate measurements were made, and holographic Mach-Zehnder interferometry was used to visualize the flow. Analysis of the data shows that computations using standard thermochemical models cannot reproduce the experimental results. The computed separation zones are smaller than the experiments indicate. However, the computed heat transfer values match the experimental data in the separation zone, and on the second wedge the computed heat transfer distribution matches the shape and heights of the experimental distribution but is shifted due to the difference in the size of the separation zones

Transition delay in hypervelocity boundary layers by means of CO 2 /acoustic instability interactions

A novel method to delay transition in hypervelocity flows over slender bodies by injecting CO2 into the boundary layer of interest is investigated. The results presented here consist of both experimental and computational data. The experimental data was obtained at Caltech’s T5 reflected shock tunnel, while the computational data was obtained at the University of Minnesota. The experimental model was a 5 degree sharp cone, chosen because of its relevance to axisymmetric hypersonic vehicle designs and the wealth of experimental and numerical data available for this geometry. The model was instrumented with thermocouples, providing heat transfer measurements from which transition locations were determined and the efficacy of adding CO2 in delaying transition was gauged. For CO2/N2 freestream blends without injection, the transition Reynolds number more than doubled for mixtures with 40% CO2 mole fraction compared to the case of 100% N2. For the cases with injection, shadowgraph visualizations were obtained, allowing verification of the injection timing. The computations provide encouraging results that for the injection schemes proposed CO2 is reaching high enough temperatures to excite vibrational modes and thus delay transition.

Effect of Gas Injection on Transition in Hypervelocity Boundary Layers

A novel method to delay transition in hypervelocity flows in air over slender bodies by injecting CO2 into the boundary layer is presented. The dominant transition mechanism in hypersonic flow is the inviscid second (Mack) mode, which is associated with acoustic disturbanceswhich are trapped and amplified inside the boundary layer [8]. In dissociated CO2-rich flows, nonequilibrium molecular vibration damps the acoustic instability, and for the high-temperature, high-pressure conditions associated with hypervelocity flows, the effect is most pronounced in the frequency bands amplified by the second mode [3].

On the Impact of Injection Schemes on Transition in Hypersonic Boundary Layers

Abstract : Three geometries are explored for injecting CO2 into the boundary layer of a sharp five degree half-angle cone. The impact of the injection geometry, namely discrete injection holes or a porous conical section, on tripping the boundary layer is examined, both with and without injected flow. The experiments are conducted at Caltechs T5 reflected shock tunnel. Two different air free-stream conditions are explored. For the discrete-hole injectors, the diameter for the injection holes is 0.75 mm nominally and the length to diameter ratio is about 30. One injector has a single row of holes and the other has four rows. With the 4-row geometry fully turbulent heat transfer values are measured within four centimeters of the last injection row for both free-stream conditions. The 1-row injector results on a reduction of 50% in the transition Reynolds number. The porous injector does not move the transition Reynolds number upstream by itself with no injection flow.

Preliminary Results on Coaxial Jet Spread Angles and the Effects of Variable Phase Transverse Acoustic Fields (Postprint)

Abstract : An experimental study on the jet spreading angle of N2 shear coaxial jets at sub-, near-, and supercritical pressures is presented. The jet spreading angle is an important parameter which characterizes the mixing between two flows forming a shear layer. The present results are compared with previous experimental data, CFD results, and theoretical predictions. The angle measurements are made directly from at least 20 backlit images. The shear coaxial injector used here is similar to those used in cryogenic liquid rockets. The chamber pressure ranges from 1.5 to 5.0 MPa to span subcritical to supercritical pressures. The chamber to outer jet density ratio varies from 0.17-4.8 and the momentum flux ratio between the outer and the inner jet varies from 0.37 to 30. These ratios are mainly varied by changing the temperature and flow rates of the outer jet. For the ranges of conditions studied it is found that the tangent of the jet spreading angle is roughly constant and approximately 0.19 with std. dev. of 0.02. The value is lower than those predicted by different theories for planar mixing layers of variable density for gaseous flows. The second part of the paper focuses on the initial results obtained by combining two piezo-sirens which generate a transverse acoustic field to excite the coaxial jet. The resonant frequency studied is approximately 3kHz and delta P/P varies from 1-1.6%. These two acoustic sources can have an arbitrary phase between them so the position of the jet with respect to the pressure and velocity field can be adjusted. The main parameter investigated is the length of the dark inner jet core. The initial results indicate an effect of the phase angle on the dark core length but the differences are statistically significant only in the extreme cases.

Carbon Dioxide Injection for Hypervelocity Boundary Layer Stability

An approach for introducing carbon dioxide as a means of stabilizing a hypervelocity boundary layer over a slender bodied vehicle is investigated through the use of numerical simulations. In the current study, two different test bodies are examined. The first is a fivedegree-half-angle cone currently under research at the GALCIT T5 Shock Tunnel with a 4 cm porous wall insert used to transpire gas into the boundary layer. The second test body is a similar cone with a porous wall over a majority of cone surface. Computationally, the transpiration is performed using an axi-symmetric flow simulation with wall-normal blowing. The effect of the injection and the transition location are gauged by solving the parabolized stability equations and using the semi-empirical e N method. The results show transition due to the injection for the first test body and a delay in the transition location for the second test body as compared to a cone without injection under the same flight conditions. The mechanism for the stabilizing effect of carbon dioxide is also explored through selectively applying non-equilibrium processes to the stability analysis. The results show that vibrational non-equilibrium plays a role in reducing disturbance amplification; however, other factors also contribute.

Transition within a hypervelocity boundary layer on a 5-degree half-angle cone in air/CO2 mixtures

Laminar to turbulent transition on a smooth 5-degree half angle cone at zero angle of attack is investigated computationally and experimentally in hypervelocity flows of air, carbon dioxide, and a mixture of 50% air and carbon dioxide by mass. Transition N factors above 10 are observed for air flows. At comparable reservoir enthalpy and pressure, flows containing carbon dioxide are found to transition up to 30% further downstream on the cone than flows in pure air in terms of x-displacement, and up to 38% and 140%, respectively, in terms of the Reynolds numbers calculated at edge and reference conditions.

Turbulent Spot Observations within a Hypervelocity Boundary Layer on a 5-degree Half-Angle Cone

Laminar to turbulent transition is a critically important process in hypersonic vehicle design. Higher thermal loads, by half an order of magnitude or more, result from the increased heat transfer due to turbulent flow. Drag, skin friction, and other flow properties are also significantly impacted. Transition to turbulence in initially laminar boundary layers can occur along many paths. In low-speed flow under ideal conditions (quiet freestream, nominally smooth surfaces with favorable or zero pressure gradient and minimal crossflow) transition occurs over a finite distance and is associated with the creation and growth of propagating patches of turbulent flow, known as turbulent spots. Spots may be due to the breakdown of linear instabilities or induced by “bypass mechanisms” associated with nonideal effects in the flow or model. H.W. Emmons (1951) was the first to propose that laminar boundary layers break down through the convergence of spots, after observations of a water-table analogy to air flow. Spot formation has been studied extensively in subsonic flows, a recent review of past and current work on spots in incompressible flows is given by Strand and Goldstein (2011).

On the Effect of a Transverse Acoustic Field on a Flush Shear Coaxial Injector

An experimental study on the effects of an externally-imposed transverse acoustic field in a flush shear coaxial jet is presented. In this case the inner jet recess is zero and both the inner and outer jet exit planes coincide. Since recess is a design variable used when designing new injectors, this study complements previous studies from this group where the injector geometries included a recess. The shear coaxial injector used here is similar to those used in cryogenic liquid rockets. By using N2 as the working fluid, the chemistry effects are separated from the fluid mechanic effects of a transverse acoustic field on coaxial jets. The acoustic field is generated by two piezo-sirens whose resonant frequency is ~3kHz. The acoustic pressures generated are about 0.2-0.8% the value of the chamber pressure. The phase angle between these two sources is varied at 45 degree intervals. Two values of pressures are studied, 1.5 MPa (Reduced Pressure, Pr=0.44) where the flow is subcritical and 3.6 MPa (Pr=1.06) where the pressure is nearcritical. The outer to inner jet velocity ratio varies from ~1.5 to 17 and the outer to inner jet momentum flux ratio (J) varies from ~0.09 to 20. These ratios are mainly varied by changing the temperature and flow rates of the outer jet. At least 3000 backlit images were taken at 20 kHz for each run. These images are the main analysis tool to study the jet behavior. The most dramatic effects resulting in about 90% reduction of the length of the inner jet core were obtained at nearcritical conditions for the J=1.7 and 3.5 cases.