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Luo Z.,University of Connecticut | Lu T.,University of Connecticut | Liu J.,Taitech Inc.
Combustion and Flame | Year: 2011

A detailed mechanism for methane-ethylene mixtures enriched with excessive amount of NO was systematically reduced for efficient numerical simulations of flames in arc-heated co-flowing air. Methane and ethylene were selected as the surrogate fuel in the present study due to their drastically different features of ignition and extinction properties and flame propagation speeds, such that the mixtures of them may be utilized to mimic practical hydrocarbon fuels with various kinetic properties in experiments. The recently released USC Mech-II for C1-C4 was grafted with the NOx sub-mechanism in GRI-Mech 3.0 with updated reaction parameters for prompt NO formation. The resulting detailed mechanism with 129 species and 900 reactions was first validated against experiments involving NOx enrichment and reasonably good agreements were observed. The detailed mechanism was then employed as the starting mechanism for the reduction. A skeletal mechanism with 44 species and 269 reactions was derived using the methods of directed relation graph (DRG) and DRG-aided sensitivity analysis (DRGASA); a 39-species reduced mechanism with 35 semi-global reaction steps was further obtained using the linearized quasi steady state approximations (LQSSA). Five species related to prompt NO were retained in the reduced mechanism because of their significant impacts on the fuel oxidation. The reduced mechanism closely agrees with the detailed mechanism for ignition and extinction of homogenous mixtures, as well as selected 1-D flames over a wide range of parameters with NO concentrations between 0% and 3%. The observed worst-case relative error of the reduction is approximately 20%. The reduced mechanism was further validated with experiments involving excessive NOx enrichment. © 2010 The Combustion Institute.

Hassan E.,University of Michigan | Hassan E.,Air Force Research Lab | Boles J.,Taitech Inc. | Aono H.,Japan Aerospace Exploration Agency | And 3 more authors.
Progress in Aerospace Sciences | Year: 2013

The supersonic jet-in-crossflow problem which involves shocks, turbulent mixing, and large-scale vortical structures, requires special treatment for turbulence to obtain accurate solutions. Different turbulence modeling techniques are reviewed and compared in terms of their performance in predicting results consistent with the experimental data. Reynolds-averaged Navier-Stokes (RANS) models are limited in prediction of fuel structure due to their inability to accurately capture unsteadiness in the flow. Large eddy simulation (LES) is not yet practical due to prohibitively large grid requirement near the wall. Hybrid RANS/LES can offer reasonable compromise between accuracy and efficiency. The hybrid models are based on various approaches such as explicit blending of RANS and LES, detached eddy simulation (DES), and filter-based multi-scale models. In particular, they can be used to evaluate the turbulent Schmidt number modeling techniques used in jet-in-crossflow simulations. Specifically, an adaptive approach can be devised by utilizing the information obtained from the resolved field to help assign the value of turbulent Schmidt number in the sub-filter field. The adaptive approach combined with the multi-scale model improves the results especially when highly refined grids are needed to resolve small structures involved in the mixing process. © 2012 Elsevier Ltd.

Storch A.M.,Alliant Techsystems | Bynum M.,Alliant Techsystems | Liu J.,Taitech Inc | Gruber M.,Air Force Research Lab
17th AIAA International Space Planes and Hypersonic Systems and Technologies Conference 2011 | Year: 2011

As part of the Hypersonic International Flight Research Experimentation (HIFiRE) Direct-Connect Rig (HDCR) test and analysis activity, three-dimensional computational fluid dynamics (CFD) simulations were performed using two Reynolds-Averaged Navier Stokes solvers. Measurements obtained from ground testing in the NASA Langley Arc-Heated Scramjet Test Facility (AHSTF) were used to specify inflow conditions for the simulations and combustor data from four representative tests were used as benchmarks. Test cases at simulated flight enthalpies of Mach 5.84, 6.5, 7.5, and 8.0 were analyzed. Modeling parameters (e.g., turbulent Schmidt number and compressibility treatment) were tuned such that the CFD results closely matched the experimental results. The tuned modeling parameters were used to establish a standard practice in HIFiRE combustor analysis. Combustor performance and operating mode were examined and were found to meet or exceed the objectives of the HIFiRE Flight 2 experiment. In addition, the calibrated CFD tools were then applied to make predictions of combustor operation and performance for the flight configuration and to aid in understanding the impacts of ground and flight uncertainties on combustor operation. © 2011 by the American Institute of Aeronautics and Astronautics, Inc.

Lin K.-C.,Taitech Inc. | Rajnicek C.,Air Force Research Lab | McCall J.,Air Force Research Lab | Carter C.,Air Force Research Lab | Fezzaa K.,Argonne National Laboratory
Nuclear Instruments and Methods in Physics Research, Section A: Accelerators, Spectrometers, Detectors and Associated Equipment | Year: 2011

Pure- and aerated-liquid jets were observed using the ultra-fast X-ray phase contrast imaging technique. Highly convoluted wrinkle structures were seen on the column surface of a turbulent pure-liquid jet, gas bubbles were discovered inside droplets and ligaments of aerated-liquid sprays, and apparently homogenous two-phase mixtures were observed inside the aerated-liquid injector. The major limitation of this X-ray technique lies in its line-of-sight nature, which can create overlapped objects/interfaces on the X-ray images. © 2011 Elsevier B.V. All rights reserved.

Liu J.,Taitech Inc | Gruber M.,Air Force Research Lab
17th AIAA International Space Planes and Hypersonic Systems and Technologies Conference 2011 | Year: 2011

The HIFiRE Flight 2 experiment was designed to study mode transition and supersonic combustion performance using a surrogate hydrocarbon fuel over a Mach number range from 5.5 to 8. This paper describes a CFD effort to evaluate the operability and combustion performance of the HF2 flowpath configuration at different flight Mach numbers. The simulation results suggest that the present combustor operates in dual mode at Mach 6 with the primary injection equivalence ratio equal to or less than 0.7, and it transitions to scramjet mode between Mach 7 and 8. At Mach 8, the combustion performance exceeds the project goal with substantial margin. To increase confidence in the CFD predictions, two baseline cases were chosen for a sensitivity study on the total equivalence ratio, simulation geometry, near-wall turbulence treatment, and turbulence model. These parameters were found to have significant impact on combustion operability and heat transfer at some conditions. The results from the present sensitivity study may improve the understanding of some uncertainties in combustor operation during the ground and flight tests.

Lin K.-C.,Taitech Inc. | Ryan M.,Universal Technology Corporation | Ryan M.,Aerojet Rocketdyne | Carter C.,U.S. Air force | And 2 more authors.
Journal of Propulsion and Power | Year: 2010

The structures of sonic ethylene jets delivered from orifices of three different diameters and two injection angles (30 and 90 deg) into a Mach 2 supersonic crossflow were studied experimentally. The ratio of the cross-sectional areas of the largest and smallest injectors is 25:1. Time-averaged spontaneous vibrational Raman scattering was used to quantify injectant concentrations by constructing two-dimensional spanwise concentration images from the onedimensional linewise Raman scattering images. Based on the present data set, new penetration height correlations were developed to treat cases with injection angles of both 30 and 90 deg. Excluding the influence of wall boundary layer, the present measurements show that the properties of fuel plume structures, such as shape, size, and concentration profiles, are scalable with the injector size. The measured ethylene concentrations were also compared with predictions from the revised jet penetration code, which was calibrated primarily with hydrogen and helium. Discrepancies were observed between the measurements and the jet penetration code predictions for the structures of ethylene fuel plumes. The experimental data generated from the present study can be used to validate the numerical simulations. © 2010 by the American Institute of Aeronautics and Astronautics, Inc.

Edwards J.R.,North Carolina State University | Boles J.A.,Taitech Inc. | Baurle R.A.,NASA
Combustion and Flame | Year: 2012

This work presents results from large-eddy/Reynolds-averaged Navier-Stokes (LES/RANS) simulations of the well-known Burrows-Kurkov supersonic reacting wall-jet experiment. Generally good agreement with experimental mole fraction, stagnation temperature, and Pitot pressure profiles is obtained for non-reactive mixing of the hydrogen jet with a non-vitiated air stream. A lifted flame, stabilized between 15 and 20. cm downstream of the hydrogen jet, is formed for hydrogen injected into a vitiated air stream. Flame stabilization occurs closer to the hydrogen injection location when a three-dimensional combustor geometry (with boundary layer development resolved on all walls) is considered. Volumetric expansion of the reactive shear layer is accompanied by the formation of large eddies which interact strongly with the reaction zone. Time averaged predictions of the reaction zone structure show an under-prediction of the peak water concentration and stagnation temperature, relative to experimental data, but display generally good agreement with the extent of the reaction zone. Reactive scalar scatter plots indicate that the flame exhibits a transition from a partially-premixed flame structure, characterized by intermittent heat release, to a diffusion-flame structure that could probably be described by a strained laminar flamelet model. © 2011 The Combustion Institute.

Lee J.,NASA | Lin K.-C.,Taitech Inc. | Eklund D.,U.S. Air force
AIAA Journal | Year: 2015

Some of the challenges in fuel injection for high-speed propulsion systems are discussed. Flush-wall circular injection is commonly used in high-speed propulsion systems, as implementation is mechanically simple and conceptually scalable. Many design parameters for flush-wall injectors, including orifice shape, orifice size, injection angle, injector placement, and the range of operating conditions, have been studied, and the resulting thermal, fluid, and mechanical properties have been documented for a broad range of injector types and conditions. Injector characterization efforts are needed to confirm the limitations of existing correlations as predictive tools for the design of larger fuel injectors.

Loh C.Y.,NASA | Loh C.Y.,Taitech Inc. | Jorgenson P.C.E.,NASA
AIAA Journal | Year: 2010

A time-accurate, upwind, finite volume method for computing compressible flows on unstructured grids is presented. The method is second-order-accurate in space and time and yields high resolution in the presence of discontinuities. In the basic Euler and Navier-Stokes upwind scheme, many concepts of high-order upwind schemes are adopted: the surface flux integrals are carefully treated, a Cauchy-Kowalewski time-stepping scheme is sed in the time-marching stage, and a multidimensional limiter is applied in the reconstruction stage. However, even with these up-to-date improvements, the basic upwind scheme is still plagued by the so-called pathological behaviors (for example, the carbuncle, the expansion shock, etc.), which are mostly triggered due to some undesirable local numerical instability.Asimple multidimensional dissipation model is used to systematically suppress such behaviors and stabilize the scheme for flows at high Mach numbers, whereas for flows at very low Mach number (for example, M = 0:02), it is found that computation can be directly carried out without invoking preconditioning. The modified, stabilized scheme is referred to as the enhanced time-accurate upwind scheme (Loh, C. Y., and Jorgenson, P. C. E., "A Time Accurate Upwind Unstructured Finite Volume Method for Compressible Flow with Cure of Pathological Behaviors," AIAA Paper 2007-4463, 2007.) in this paper. The unstructured grid capability renders flexibility for use in complex geometry, and the present enhanced time-accurate upwind Euler and Navier-Stokes scheme is capable of handling a broad spectrum of flow regimes from high supersonic to subsonic at very low Mach number, appropriate for both computational fluid dynamics and computational aeroacoustics. Numerous examples are included to demonstrate the robustness of the scheme. Copyright Clearance Center, Inc.

Tam C.-J.,Taitech Inc. | Hsu K.-Y.,Innovative Scientific Solutions, Inc. | Hagenmaier M.,U.S. Air force | Raffoul C.,U.S. Air force
Journal of Propulsion and Power | Year: 2013

Direct-connect wind-tunnel facilities that produce uniform flow entering the test section are generally used for scramjet-component (isolator/combustor) studies. In freejet experiments and flight tests, however, air enters the engine through an inlet, and flow entering the isolator and combustor is typically distorted. The distortion effects can include nonuniform boundary-layer thicknesses on the walls and relatively strong oblique shock waves. This research focuses on the effects of inlet distortion on a round scramjet isolator. A numerical study was performed using various distortion devices, including ramps, injector ports, and injector slots, that were placed downstream of the directconnect facility nozzle to simulate a distorted flowfield from a prescribed inlet. The computational results provided a methodology for simulating flow distortion in direct-connect testing. Based on the numerical findings, experimental testing was conducted in the supersonic wind-tunnel facility to validate the numerical results and determine the impact of flow distortion. In addition, this paper focuses on design, fabrication, and execution of flow-distortion experiments. Air injection in the distortion section was used to create the flow distortion, and shock angle increases with the injection flow. Comparisons of wall pressure, exit flow profiles, and shock-holding capability are made from the experimental and numerical results. Good agreement between these data on wall pressures and flow profiles was found. Copyright © 2013 by the von Karman Institute for Fluid.

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