Sibtosh Pal

Atomization characteristics of impinging liquid jets

The atomization characteristics of sheets formed by both laminar and turbulent impinging jets were experimentally studied as a function of flow and injector geometric parameters. In particular, sheet breakup length along the sheet centerline, distance between adjacent waves apparent on the sheet, and drop-size distributions were measured over a Weber number range between 350-6600 and a Reynolds number range between 2.8 x 103 to 2.6 x 10 4. A linear stability-based model was used to determine the most unstable wave number and the corresponding growth rate factor on two-dimension al thinning inviscid and viscous sheets. These wave characteristics were used to predict both the sheet breakup length and the resulting drop sizes. A second model, applicable for a low Weber number regime, in which sheet disintegration is controlled by stationary antisymmetric waves, was used to predict the shape of the sheet formed by two impinging liquid jets. The linear stabilitybased theory predictions of breakup length did not agree in trend or magnitude with experimental measurements. However, for Weber numbers less than 350, the measured breakup length for laminar impinging jets was within 50% of that predicted by the stationary antisymmetric wave-based model. Finally, drop-size predictions based on linear stability theory agreed in trend, but not in magnitude, with the measured drop sizes. The contrast between the sheet atomization characteristics of laminar vs turbulent impinging jets suggest that the initial conditions of the impinging jets significantly influence the sheet breakup mechanism. Also, the comparison between experimental results and theoretical predictions indicates that the impact wave generation process at the jet impingement point needs to be incorporated in the theoretical models for sheet atomization. Nomenclature d = diameter F = thickness distribution h = sheet thickness k = wave number L = length of injection element / = length r = radial distance from impingement point Re = Reynolds number, Ujdjv/, based on liquid properties, jet velocity, and orifice diameter Rex = Reynolds number, Ushlvi, based on liquid properties, sheet velocity, and sheet thickness 5 = ratio of gas density to liquid density t = time U = velocity W = maximum width of sheet We = Weber number, piUjdJcr, based on liquid properties, jet velocity, and orifice diameter Wes = Weber number, p,£/;/z/cr, based on liquid properties, sheet velocity, and sheet thickness x = axial distance from impingement point y = coordinate perpendicular to x in the plane of the sheet a = fan inclination angle ft = complex growth rate factor, pr + //3, 77 = disturbance amplitude 6 = impingement half-angle A = wavelength ju = dynamic viscosity

AN EXPERIMENTAL ESTIMATION OF MEAN REACTION RATE AND FLAME STRUCTURE DURING COMBUSTION INSTABILITY IN A LEAN PREMIXED GAS TURBINE COMBUSTOR

The interactions between the local flame structure produced by a premixed swirl-stabilized injector with combustion instabilities were experimentally studied for a model gas turbine combustor operating at high pressure and temperature. The model gas turbine combustor studied utilizes a sudden-expansion dump combustor with a single swirler and bluff body for enhancing mixing rate and flame stabilization, respectively. Laser-based measurements were made for both stable and unstable operating conditions. The local flame front structure was visualized using planar laser-induced fluorescence (PLIF) from the OH⊙, and the global heat release zone was interpreted from flame emission measurements. For stable combustion conditions, the mean reaction rate estimated independently from both OH-PLIF and OH * chemiluminescence measurements showed good agreement, thereby indicating confidence in the use of OH-PLIF measurements for extracting the local mean reaction rate. For unstable combustion conditions, the flamefront characteristics, including flame surface density and mean reaction rate, were evaluated together with the information from the OH * chemiluminescence measurements to identify the boundary of the heat release region at discrete phases of the unstable flame. Analysis of the flame structures during combustion instability indicated significant variations during different phases of the instability. The heat release flow field, particularly in the recirculation regions appearing at the corner and inner face of the dump plane, varied substantially. Rayleigh index information indicated that the recirculation zones play an important part in driving the instability. In contrast, the high shear layer formed along the interface between reactants and hot products produced a region where the instability was damped due to a lowering of the heat release.

Deflagration to detonation transition processes by turbulence-generating obstacles in pulse detonation engines

The results from a series of detonation experiments conducted to characterize the deflagration-to-detonation transition (DDT) process for ethylene-air mixtures in a 44-mm-square, 1.65-m-long tube are described. Experiments were conducted for both single-shot detonations involving quiescent mixtures as well as multicycle detonations involving dynamic fill. For the experiments, high-frequency pressure and flame emission measurements were made to obtain the compression wave and flame speeds, respectively. In addition, schlieren and hydroxyl-radical/planar-laser-induced-fluorescence (OH-PLIF) imaging were applied to investigate the interactions between the shock-wave and combustion phenomena during both deflagration and detonation. For ethylene-air mixtures, strategically placed obstacles were necessary to achieve DDT. The effect of the presence of obstacles on flame acceleration was systematically investigated by changing the obstacle configuration. The parametric study of obstacle blockage ratio, spacing between obstacles, and length of the obstacle configuration indicated that for successful detonations the obstacle needs to accelerate the flame to a minimum flame speed of roughly half the Chapman-Jouguet detonation velocity. Differences in the flame and compression wave velocities demonstrated the development of a coupled feedback mechanism as the wave propagated along the tube. A series of simultaneous schlieren and OH-PLIF images showed that the obstacle plays a major role in generating small/large-scale turbulence that enhances flame acceleration. Localized explosions of pockets of unburned mixture further enhanced the shock-wave strength to continuously increase the flame speed. The results of this experimental study support the importance of obstacles as a means to enhance DDT and provide a potential solution for practical pulse-detonation-engine applications.

Benchmark Wall Heat Flux Data for a GO2/GH2 Single Element Combustor

Wall heat flux measurements in a 1.5 in. diameter circular cross-section rocket chamber for a uni-element shear coaxial injector element operating on gaseous oxygen (GOz)/gaseous hydrogen (GH,) propellants are presented. The wall heat flux measurements were made using arrays of Gardon type heat flux gauges and coaxial thermocouple instrumentation. Wall heat flux measurements were made for two cases. For the first case, GOZ/GHz oxidizer-rich (O/F=l65) and fuel-rich preburners (O/F=1.09) integrated with the main chamber were utilized to provide vitiated hot fuel and oxidizer to the study shear coaxial injector element. For the second case, the preburners were removed and ambient temperature gaseous oxygen/gaseous hydrogen propellants were supplied to the study injector. Experiments were conducted at four chamber pressures of 750, 600, 450 and 300psia for each case. The overall mixture ratio for the preburner case was 6.6, whereas for the ambient propellant case, the mixture ratio was 6.0. Total propellant flow was nominally 0.27-0.29 Ibm/s for the 750 psia case with flowrates scaled down linearly for lower chamber pressures. The axial heat flux profile results for both the preburner and ambient propellant cases show peak heat flux levels a t axial locations between 2.0 and 3.0 in. from the injector face. The maximum heat flux level was about two times greater for the preburner case. This is attributed to the higher injector fuel-to-oxidizer momentum flux ratio that promotes mixing and higher initial propellant temperature for the preburner case which results in a shorter reaction zone. The axial heat flux profiles were also scaled with respect to the chamber pressure to the power 0.8. The results at the four chamber pressures for both cases collapsed to a single profile indicating that at least to first approximation, the basic fluid dynamic structures in the flow field are pressure independent as long as the chamber/njector/nozzle geometry and injection velocities remain the same.

Experimental and analytical characterization of a shear coaxial combusting GO2/GH2 flowfield

Detailed measurements of the flow velocities and major species distributions downstream of a uni-element shear coaxial injector in a gaseous oxygen/hydrogen (GO2/GH2) rocket chamber are presented and compared with CFD predictions from a reacting Navier-Stokes solution. The velocity profiles are obtained by means of laser Doppler velocimetry, while the species (H2, O2, and H2O) are measured by Raman spectroscopy. The CFD predictions are in good agreement with the measurements, although they moderately overpredict the spreading rate of the reaction zone at the downstream location. Both experimental and numerical results indicate that the shear coaxial injector mixes slowly but that combustion is completed on a length scale representative of that tested in actual combustors. The experiments show high c* efficiency, and the computations show the flame combusts all the GO2, both suggesting complete combustion. Detailed CFD resolution shows that the base region behind the GO2 post tip acts as a flameholder for the GO2/GH2 flame. Overall, the results demonstrate that advanced diagnostics and CFD predictions can be used in concert as tools for gas/gas injector design. (Author)

Shear and Swirl Coaxial Injector Studies of LOX/GCH4 Rocket Combustion Using Non-Intrusive Laser Diagnostics

Uni-element liquid rocket experiments were completed for two different swirl coaxial injector configurations and a shear coaxial injector configuration using liquid oxygen/gaseous methane propellants. Several different non-intrusive optical techniques such as OH planar laser induced fluorescence, OH* chemiluminescence, laser light scattering and shadowgraph imaging, were employed to observe flame position and liquid core structure. All measurements were made in the near injector region of the chamber for steady state chamber pressures of approximately 4.14 MPa and propellant mass flowrates of 0.118 kg/s and 0.039 kg/s for liquid oxygen (LOX) and gaseous methane respectively. The two swirl injectors studied varied in the size of the methane annulus, maintaining the LOX post diameter constant. The shear coaxial injector was identical in size to the swirl injector with the smallest fuel annulus but the liquid flow was not swirled. The application of optical diagnostics, particularly those involving simultaneous measurements using two different techniques, provided detail information on the spray and flame structure for each injector. OH radicals were observed to be located in fuel rich areas of the flow while OH* chemiluminescence resided very close to the edge of LOX core as determined by light scattering measurements. Differences in the OH-PLIF signal location in this study and previous studies using hydrogen (H2) and LOX were explained by differences in injector geometry and mixing processes. The flow structures, atomization process and mixing were found to have a major effect on the location of OH-PLIF, OH* chemiluminescence and the combustion efficiency.

LOX/GH2 Shear Coaxial Injector Atomization Studies at Large Momentum Flux Ratios

The general view of the atomization process for gas/liquid shear-coaxial rocket engine injectors envisions a relatively short intact liquid core from which ligaments or drops are continually shed due to surface instabilities. At some point in the process, the presence of the intact liquid core ceases and only the drop field remains. This classical phenomenological breakup model indicates that the progress of liquid atomization depends primarily on the momentum flux and/or velocity ratios between the gas and liquid streams. Recent rocket combustion research experiments at gas-to-liquid momentum flux ratios (J) of about one to five had cast doubt on this breakup/atomization model for some typical injection conditions. Some rocket engines, such as the Ariane 5 Vulcain, run injectors at higher momentum fluxes (J ≈ 10-11). To evaluate the applicability of the core-stripping spray model at even extreme conditions for rocket applications, liquid oxygen (LOX) / gaseous hydrogen (GH2) combustion experiments were carried out at much higher momentum flux values (J ~ 22 and 50). A shadowgraph imaging technique was applied to record the LOX jet structure at various axial locations from which LOX dense-core length measurements were made. LOX core lengths were found to scale with the inverse of momentum flux; however, the flowfield even at J > 20, exhibits a long sinuous LOX core region, eventually breaking up into large LOX structures that gasify in the core wake. This LOX core fragmentation process seems to dominate the primary atomization process even at very high momentum flux ratios. Several existing core length correlations, none of which were developed under combustion conditions, were examined for applicability to realistic rocket engine conditions.

EXPERIMENTAL STUDY OF A PULSE DETONATION ENGINE DRIVEN EJECTOR

Results of an experimental effort on pulse detonation driven ejectors are presented and discussed. The experiments were conducted using a pulse detonation engine (PDE)/ejector setup that was specifically designed for the study. The results of various experiments designed to probe different aspects of th e PDE/ejector setup are reported. The baseline PDE was operated using ethylene (C2H4) as the fuel and an oxygen/nitrogen (O2 + N2) mixture at an equivalence ratio of one. The PDE only experiments included propellant mixture characterization using a laser absorption technique, high fidelity thrust measurements using an integrated spring-damper system, and shadowgraph imaging of the detonation/shock wave structure emanating from the tube. The baseline PDE thrust measurement results are in excellent agreement with experimental and modeling results reported in the literature. These PDE setup results were then used as a basis for quantifying thrust augmentation for various PDE/ejector setups with constant diameter ejector tubes and various detonation tube/ejector tube overlap distances. The results show that for the geometries studied here, a maximum thrust augmentation of 24% is achieved. Further increases are possible by tailoring the ejector geometry based on CFD predictions conducted elsewhere. The thrust augmentation results are complemented by shadowgraph imaging of the flowfield in the ejector tube inlet area and high frequency pressure transducer measurements along the length of the ejector tube.

Experimental Study of Major Species and Temperature Profiles of Liquid Oxygen/Gaseous Hydrogen Rocket Combustion

Understanding thefundamentalsofhigh-pressure, high-temperaturecombustion of practical rocketpropellants is key to improving rocket engine performance and efe ciency. The application of Raman spectroscopy for making detailed species concentration measurements of oxygen, hydrogen, and water vapor in the elevated-pressure, combusting e owe eld downstream of a liquid-oxygen/gaseous-hydrogen swirl-coaxial injector element at fuel-rich conditionsispresented.Detailsregardingmeasurementerroranalysisandrecommendationsforfurtherree nement ofthediagnostictechniquearealsopresented.To theauthors’ knowledge,theseexperimentsrepresentthee rsttime that instantaneous temperature and major species proe les have been successfully made in such an environment. Theseresults,whichincludebothinstantaneousand time-averagedtemperatureand speciesconcentrationproe les, impact both rocket fuel preburner design and computational e uid dynamic code validation activities. Nomenclature f .T/ = temperature dependent function that relates the spectral bandwidth strength to the signal strength = focal length =# = focal length divided by the diameter of the lens K = constant that accounts for the laser pulse energy, species Raman cross section, optical collection efe ciency, and optical solid angle N = measurement uncertainty n = number density S = Raman signal intensity aeB = background shot noise due to e ame luminosity aeD = dark current noise of the intensie er aeR = Readout noise aeS = signal shot noise due to Raman signal

Experimental and Computational Investiagtion of Combustor Acoustics and Instabilities, Part II: Transverse Modes

A combined experimental-computational study of transverse acoustic modes and combustion instabilities in a rectangular liquid rocket chamber is presented. Experimental results show that transverse modes can be spontaneously excited in the rectangular chamber. The amplitudes of the acoustic response are governed by the number and location of the injector elements. In general, stronger response of the 1W mode is observed when the injector element is positioned near a pressure anti-nodal location. Companion CFD solutions of the Euler and Navier-Stokes solutions are also performed and compared with the experimental measurements. Good qualitative agreement of the acoustic chamber response is obtained. Further, the computational studies are utilized to perform parametric studies of the eects of non-linear forcing and viscous eects due to the presence of side-wall boundary layers.