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  • 1
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    Measurements on NASA Langley Durable Combustor Rig by TDLAT: Preliminary Results (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Detailed knowledge of the internal structure of high-enthalpy flows can provide valuable insight to the performance of scramjet combustors. Tunable Diode Laser Absorption Spectroscopy (TDLAS) is often employed to measure temperature and species concentration. However, TDLAS is a path-integrated line-of-sight (LOS) measurement, and thus does not produce spatially resolved distributions. Tunable Diode Laser Absorption Tomography (TDLAT) is a non-intrusive measurement technique for determining two-dimensional spatially resolved distributions of temperature and species concentration in high enthalpy flows. TDLAT combines TDLAS with tomographic image reconstruction. Several separate line-of-sight TDLAS measurements are analyzed in order to produce highly resolved temperature and species concentration distributions. Measurements have been collected at the University of Virginia's Supersonic Combustion Facility (UVaSCF) as well as at the NASA Langley Direct-Connect Supersonic Combustion Test Facility (DCSCTF). Measurements collected at the DCSCTF required significant modifications to system hardware and software designs due to its larger measurement area and shorter test duration. Initial LOS measurements from the NASA Langley DCSCTF operating at an equivalence ratio of 0.5 are presented. Results show the capability of TDLAT to adapt to several experimental setups and test parameters.
    Keywords: Optics
    Type: AIAA Paper 2013-0696 , NF1676L-15902 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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  • 2
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    The Extraction of One-Dimensional Flow Properties from Multi-Dimensional Data Sets (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: The engineering design and analysis of air-breathing propulsion systems relies heavily on zero- or one-dimensional properties (e.g. thrust, total pressure recovery, mixing and combustion efficiency, etc.) for figures of merit. The extraction of these parameters from experimental data sets and/or multi-dimensional computational data sets is therefore an important aspect of the design process. A variety of methods exist for extracting performance measures from multi-dimensional data sets. Some of the information contained in the multi-dimensional flow is inevitably lost when any one-dimensionalization technique is applied. Hence, the unique assumptions associated with a given approach may result in one-dimensional properties that are significantly different than those extracted using alternative approaches. The purpose of this effort is to examine some of the more popular methods used for the extraction of performance measures from multi-dimensional data sets, reveal the strengths and weaknesses of each approach, and highlight various numerical issues that result when mapping data from a multi-dimensional space to a space of one dimension.
    Keywords: Spacecraft Propulsion and Power
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  • 3
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    Numerical Simulations of a Co-Axial Supersonic-Combusting Free-Jet Experiment (2008)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: CFD calculations using the Reynolds-averaged Navier-Stokes equations coupled with species continuity equations have been made for a supersonic coaxial-jet CFD-validation experiment to determine the sensitivity of the external flowfield to the main-nozzle exit profile. Four different nozzle exit profiles were used in the study: a uniform profile, one computed using only the nozzle geometry, one computed using the nozzle geometry and part of the upstream facility combustor, and one using the nozzle and the full facility combustor. Two cases were examined using the four profiles: a non-reacting case without coflow and a reacting case with hydrogen coflow. Results show that the nozzle exit profile has a significant effect on the external flowfield. The uniform profile produced the longest jet while the profile created with the full combustor produced the shortest jet. The nozzle-only and part-combustor profiles fell between the other two profiles. The reacting flow was found to be more sensitive to the nozzle exit profile since it affects the downstream mixing and combustion. These calculations indicate the importance of properly setting the nozzle-exit profile for this type of calculation.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: 55th JANNAF/42nd combustion/30th Airbreathing Propulsion/30th Exhaust Plume Technology/24th Propulsion Systems Hazards/12th SPIRITS User Group Joint Subcommittee Meeting; May 12, 2008 - May 16, 2008; Newton, MA; United States
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  • 4
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    CFD Modeling Needs and What Makes a Good Supersonic Combustion Validation Experiment (2005)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: If a CFD code/model developer is asked what experimental data he wants to validate his code or numerical model, his answer will be: "Everything, everywhere, at all times." Since this is not possible, practical, or even reasonable, the developer must understand what can be measured within the limits imposed by the test article, the test location, the test environment and the available diagnostic equipment. At the same time, it is important for the expermentalist/diagnostician to understand what the CFD developer needs (as opposed to wants) in order to conduct a useful CFD validation experiment. If these needs are not known, it is possible to neglect easily measured quantities at locations needed by the developer, rendering the data set useless for validation purposes. It is also important for the experimentalist/diagnostician to understand what the developer is trying to validate so that the experiment can be designed to isolate (as much as possible) the effects of a particular physical phenomena that is associated with the model to be validated. The probability of a successful validation experiment can be greatly increased if the two groups work together, each understanding the needs and limitations of the other.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: JANNAF 40th Combustion/28th Airbreathing Propulsion/22nd Propulsion Systems Hazards/4th Modeling and Simulation Joint Subcommittee Meeting; Jun 13, 2005 - Jun 17, 2005; Charleston, SC; United States
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  • 5
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    Analysis of the Effects of Vitiates on Surface Heat Flux in Ground Tests of Hypersonic Vehicles (2008)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: To achieve the high enthalpy conditions associated with hypersonic flight, many ground test facilities burn fuel in the air upstream of the test chamber. Unfortunately, the products of combustion contaminate the test gas and alter gas properties and the heat fluxes associated with aerodynamic heating. The difference in the heating rates between clean air and a vitiated test medium needs to be understood so that the thermal management system for hypersonic vehicles can be properly designed. This is particularly important for advanced hypersonic vehicle concepts powered by air-breathing propulsion systems that couple cooling requirements, fuel flow rates, and combustor performance by flowing fuel through sub-surface cooling passages to cool engine components and preheat the fuel prior to combustion. An analytical investigation was performed comparing clean air to a gas vitiated with methane/oxygen combustion products to determine if variations in gas properties contributed to changes in predicted heat flux. This investigation started with simple relationships, evolved into writing an engineering-level code, and ended with running a series of CFD cases. It was noted that it is not possible to simultaneously match all of the gas properties between clean and vitiated test gases. A study was then conducted selecting various combinations of freestream properties for a vitiated test gas that matched clean air values to determine which combination of parameters affected the computed heat transfer the least. The best combination of properties to match was the free-stream total sensible enthalpy, dynamic pressure, and either the velocity or Mach number. This combination yielded only a 2% difference in heating. Other combinations showed departures of up to 10% in the heat flux estimate.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: 55th JANNAF Propulsion Meeting/42nd Combustion/30th Airbreathing Propulsion/30th Exhaust Plume Technology/24th Propulsion Systems Hazards/12th SPIRITS User Group Joint Subcommittee Meeting; May 12, 2008 - May 16, 2008; Newton, MA; United States
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  • 6
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    Analysis and Design of Rectangular-Cross-Section Nozzles for Scramjet Engine Testing (2004)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The flow in the square-cross-section Mach-6 nozzle used in the NASA Langley Research Center Arc-Heated Scramjet Test Facility has been analyzed using three-dimensional viscous CFD methods. The primary cause of the non-uniform flow exiting the nozzle is identified as cross-flow pressure gradients imposed on wall boundary layers. The cross-flow pressure gradients cause the boundary layer to roll up into counter-rotating vortex pairs on each of the four sides of the nozzle. These four vortex pairs produce significant non-uniformity in the nozzle-exit flow. In order to improve the quality of the test flow in the facility, two alternative nozzle designs (one axisymmetric and one rectangular with a 2-D contour) have been investigated. While the axisymmetric design produced the most uniform flow, the 2-D design also produced very good flow. The 2-D design was selected for further refinement, resulting in a new nozzle design which has been constructed and awaits calibration.
    Keywords: Aircraft Propulsion and Power
    Type: AIAA Paper 2004-1137 , 42nd AIAA Aerospace Science Meeting and Exhibit; Jan 05, 2004 - Jan 08, 2004; Reno, NV; United States
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  • 7
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    Design of a Mach-15 Total-Enthalpy Nozzle With Non-uniform Inflow Using Rotational MOC (2004)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A new computer program to design nozzles with non-uniform inflow has been developed using the rotational method of characteristics (MOC). This program has been used to design a nozzle for the NASA's HYPULSE shock-expansion tunnel for use in scramjet engine tests at a Mach-15 flight-enthalpy condition. The nozzle has an area ratio of 9.5:1 that expands the inflow from Mach 6 along the centerline to Mach 8.7. Although the density and Mach number vary radially at the exit due to the non-uniformities of the inflow, the MOC procedure produces exit flow that is parallel and has uniform static pressure. The design has been verified with CFD which compares favorably with the MOC solution.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: AIAA Paper 2005-0691 , 43rd AIAA Aerospace Sciences Meeting and Exhibit; Jan 10, 2005 - Jan 13, 2005; Reno, NV; United States
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  • 8
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    Design of a Mach-15 Total-Enthalpy Nozzle with Non-Uniform Inflow Using Rotational MOC (2005)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A new computer program to design nozzles with non-uniform idow has been developed using the rotational method of characteristics (MOC). This program has been used to design a nozzle for the NASA's HYPULSE shock-expansion tunnel for use in scramjet engine tests at a Mach-15 flight-enthalpy condition. The nozzle has an area ratio of 9.5:l that expands the inflow from Mach 6 along the centerline to Mach 8.7. Although the density and Mach number vary radially at the exit due to the non-uniformities of the inflow, the MOC procedure produces exit flow that is parallel and has uniform static pressure. The design has been verified with CFD which compares favorably with the MOC solution.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: AIAA Paper 2005-0691 , 43rd AIAA Aerospace Sciences Meeting and Exhibit; Jan 10, 2005 - Jan 13, 2005; Reno, NV; United States
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  • 9
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    Design of a Mach-3 Nozzle for TBCC Testing in the NASA LaRC 8-ft High Temperature Tunnel (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A new nozzle is being constructed for the NASA Langley Research Center 8-Foot High Temperature Tunnel. The axisymmetric nozzle was designed with a Mach-3 exit flow for testing Turbine-Based Combined-Cycle engines at a Mach number in the vicinity of the transition from turbojet to ramjet operation. The nozzle contour was designed using the NASA Langley IMOCND computer program which solves the potential equation using the classical method of characteristics. To include viscous effects, the design procedure iterated the MOC contour generation with CFD Navier-Stokes calculations, adjusting MOC input parameters until target nozzle-exit conditions were achieved in the Navier-Stokes calculations. The design process was complicated by a requirement to use the final 29.5 inches of an existing 54.5-inch exit-diameter Mach-5 nozzle contour. This was accomplished by generating a Mach-3 contour that matched the radius of the Mach-5 contour at the match point and using a 3rd order polynomial to create a smooth transition between the two contours. During the final evaluation of the design it was realized that the throat diameter is more than half that of the upstream mixing chamber. This led to the concern that large vortical structures generated in the mixer would persist downstream, affecting nozzle-exit flow. This concern was addressed by analyzing the results of three-dimensional, viscous, numerical simulations of the entire flowfield, from the exit of the facility combustor to the nozzle exit. An analysis of the solution indicated that large scale structures do not pass through the throat and that both the total temperature and species (CO2) are well mixed in the mixer, providing uniform flow to the nozzle and subsequently the test cabin.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: AIAA Paper-2008-3703 , 26th AIAA Aerodynamic Measurement Technology and Ground Testing Conference; Jun 23, 2008 - Jun 26, 2008; Seattle, WA; United States
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  • 10
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    Modeling Combustion in Supersonic Flows (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper discusses the progress of work to model high-speed supersonic reacting flow. The purpose of the work is to improve the state of the art of CFD capabilities for predicting the flow in high-speed propulsion systems, particularly combustor flow-paths. The program has several components including the development of advanced algorithms and models for simulating engine flowpaths as well as a fundamental experimental and diagnostic development effort to support the formulation and validation of the mathematical models. The paper will provide details of current work on experiments that will provide data for the modeling efforts along with with the associated nonintrusive diagnostics used to collect the data from the experimental flowfield. Simulation of a recent experiment to partially validate the accuracy of a combustion code is also described.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: 3rd International Symposium on Non-Equilibrium Processes, Plasma, Combustion, and Atmospheric Phenomena (NEPCAP 2007); Jun 25, 2007 - Jun 29, 2007; Sochi; Russia
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