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  • 1
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    Design Evolution and Performance Characterization of the GTX Air-Breathing Launch Vehicle Inlet (2002)
    In:  Other Sources
    Details
    Publication Date: 2011-08-23
    Description: The design and analysis of a second version of the inlet for the GTX rocket-based combine-cycle launch vehicle is discussed. The previous design did not achieve its predicted performance levels due to excessive turning of low-momentum comer flows and local over-contraction due to asymmetric end-walls. This design attempts to remove these problems by reducing the spike half-angle to 10- from 12-degrees and by implementing true plane of symmetry end-walls. Axisymmetric Reynolds-Averaged Navier-Stokes simulations using both perfect gas and real gas, finite rate chemistry, assumptions were performed to aid in the design process and to create a comprehensive database of inlet performance. The inlet design, which operates over the entire air-breathing Mach number range from 0 to 12, and the performance database are presented. The performance database, for use in cycle analysis, includes predictions of mass capture, pressure recovery, throat Mach number, drag force, and heat load, for the entire Mach range. Results of the computations are compared with experimental data to validate the performance database.
    Keywords: Launch Vehicles and Launch Operations
    Type: 26th JANNAF Airbreathing Propulsion Subcommittee Meeting; Volume 1; 107-123; CPIA-Publ-713-Vol-1
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  • 2
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    Computational Analysis for Rocket-Based Combined-Cycle Systems During Rocket-Only Operation (2000)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: A series of Reynolds-averaged Navier-Stokes calculations were employed to study the performance of rocket-based combined-cycle systems operating in an all-rocket mode. This parametric series of calculations were executed within a statistical framework, commonly known as design of experiments. The parametric design space included four geometric and two flowfield variables set at three levels each, for a total of 729 possible combinations. A D-optimal design strategy was selected. It required that only 36 separate computational fluid dynamics (CFD) solutions be performed to develop a full response surface model, which quantified the linear, bilinear, and curvilinear effects of the six experimental variables. The axisymmetric, Reynolds-averaged Navier-Stokes simulations were executed with the NPARC v3.0 code. The response used in the statistical analysis was created from Isp efficiency data integrated from the 36 CFD simulations. The influence of turbulence modeling was analyzed by using both one- and two-equation models. Careful attention was also given to quantify the influence of mesh dependence, iterative convergence, and artificial viscosity upon the resulting statistical model. Thirteen statistically significant effects were observed to have an influence on rocket-based combined-cycle nozzle performance. It was apparent that the free-expansion process, directly downstream of the rocket nozzle, can influence the Isp efficiency. Numerical schlieren images and particle traces have been used to further understand the physical phenomena behind several of the statistically significant results.
    Keywords: Spacecraft Propulsion and Power
    Type: Journal of Propulsion and Power; Volume 16; No. 6; 1030-1039
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  • 3
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    Two equation modelling and the pseudo compressibility technique (1992)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: The primary objective of the Center for Modelling of Turbulence and Transition (CMOTT) is to further the understanding of turbulence theory for engineering applications. One important foundation is the establishment of a data base encompassing the multitude of existing models as well as newly proposed ideas. The research effort described is a precursor to an extended survey of two equation turbulence models in the presence of a separated shear layer. Recently, several authors have examined the performance of two equation models in the context of the backward facing step flow. Conflicting results, however, demand that further attention is necessary to properly understand the behavior and limitations of this popular technique, especially the low Reynolds number formulations. The objective is to validate an incompressible Navier Stokes code for use as a numerical test-bed. In turn, this code will be used for analyzing the performance of several two equation models.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: Center for Modeling of Turbulence and Transition (CMOTT): Research Briefs, 1992; p 95-100
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  • 4
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    Response Surface Modeling of Combined-Cycle Propulsion Components using Computational Fluid Dynamics (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Three examples of response surface modeling with CFD are presented for combined cycle propulsion components. The examples include a mixed-compression-inlet during hypersonic flight, a hydrogen-fueled scramjet combustor during hypersonic flight, and a ducted-rocket nozzle during all-rocket flight. Three different experimental strategies were examined, including full factorial, fractionated central-composite, and D-optimal with embedded Plackett-Burman designs. The response variables have been confined to integral data extracted from multidimensional CFD results. Careful attention to uncertainty assessment and modeling bias has been addressed. The importance of automating experimental setup and effectively communicating statistical results are emphasized.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NASA/TM-2002-211379 , NAS 1.15:211379 , E-13204 , AIAA Paper 2002-0542 , 40th Aerospace Sciences Meeting and Exhibit; Jan 14, 2002 - Jan 17, 2002; Reno, NV; United States
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  • 5
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    Numerical Analysis of the Trailblazer Inlet Flowfield for Hypersonic Mach Numbers (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A study of the Trailblazer vehicle inlet was conducted using the Global Air Sampling Program (GASP) code for flight Mach numbers ranging from 4-12. Both perfect gas and finite rate chemical analysis were performed with the intention of making detailed comparisons between the two results. Inlet performance was assessed using total pressure recovery and kinetic energy efficiency. These assessments were based upon a one-dimensional stream-thrust-average of the axisymmetric flowfield. Flow visualization utilized to examine the detailed shock structures internal to this mixed-compression inlet. Kinetic energy efficiency appeared to be the least sensitive to differences between the perfect gas and finite rate chemistry results. Total pressure recovery appeared to be the most sensitive discriminator between the perfect gas and finite rate chemistry results for flight Mach numbers above Mach 6. Adiabatic wall temperature was consistently overpredicted by the perfect gas model for flight Mach numbers above Mach 4. The predicted shock structures were noticeably different for Mach numbers from 6-12. At Mach 4, the perfect gas and finite rate chemistry models collapse to the same result.
    Keywords: Aerodynamics
    Type: NASA/TM-1999-209654 , NAS 1.15:209654 , AIAA Paper 2000-0889 , E-12010 , 38th Aerospace Sciences Meeting; Jan 10, 2000 - Jan 13, 2000; Reno, NV; United States
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  • 6
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    Computation of Three-Dimensional Compressible Flow From a Rectangular Nozzle with Delta Tabs (1999)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: A three-dimensional viscous flow analysis is performed using a time-marching Reynolds-averaged Navier-Stokes code for a 3:1 rectangular nozzle with two delta tabs located at the nozz1e exit plane to enhance mixing. Two flow configurations, a subsonic jet case and a supersonic jet case using the same rate configuration, which were previously studied experimentally, are computed and compared with the experimental data. The experimental data include streamwise velocity and vorticity distributions for the subsonic case, and Mach number distributions for the supersonic case, at various axial locations downstream of the nozzle exit. The computational results show very good agreement with the experimental data. In addition, the effect of compressibility on vorticity dynamics is examined by comparing the vorticity contours of the subsonic jet case with those of the supersonic jet case which were not measured in the experiment.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: Paper-97-GT-257 , Journal of Engineering for Gas Turbines and Power; 121; 235-242|International Gas Turbine and Aeroengine Congress and Exhibition; Jun 02, 1997 - Jun 05, 1997; Orlando, FL; United States
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  • 7
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    Inlet Development for a Rocket Based Combined Cycle, Single Stage to Orbit Vehicle Using Computational Fluid Dynamics (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Design and analysis of the inlet for a rocket based combined cycle engine is discussed. Computational fluid dynamics was used in both the design and subsequent analysis. Reynolds averaged Navier-Stokes simulations were performed using both perfect gas and real gas assumptions. An inlet design that operates over the required Mach number range from 0 to 12 was produced. Performance data for cycle analysis was post processed using a stream thrust averaging technique. A detailed performance database for cycle analysis is presented. The effect ot vehicle forebody compression on air capture is also examined.
    Keywords: Launch Vehicles and Space Vehicles
    Type: NASA/TM-1999-209279 , NAS 1.15:209279 , AIAA Paper 99-2239 , E-11744 , Joint Propulsion; Jun 20, 1999 - Jun 24, 1999; Los Angeles, CA; United States
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  • 8
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    Analysis of a Rocket Based Combined Cycle Engine during Rocket Only Operation (1998)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The all rocket mode of operation is a critical factor in the overall performance of a rocket based combined cycle (RBCC) vehicle. However, outside of performing experiments or a full three dimensional analysis, there are no first order parametric models to estimate performance. As a result, an axisymmetric RBCC engine was used to analytically determine specific impulse efficiency values based upon both full flow and gas generator configurations. Design of experiments methodology was used to construct a test matrix and statistical regression analysis was used to build parametric models. The main parameters investigated in this study were: rocket chamber pressure, rocket exit area ratio, percent of injected secondary flow, mixer-ejector inlet area, mixer-ejector area ratio, and mixer-ejector length-to-inject diameter ratio. A perfect gas computational fluid dynamics analysis was performed to obtain values of vacuum specific impulse. Statistical regression analysis was performed based on both full flow and gas generator engine cycles. Results were also found to be dependent upon the entire cycle assumptions. The statistical regression analysis determined that there were five significant linear effects, six interactions, and one second-order effect. Two parametric models were created to provide performance assessments of an RBCC engine in the all rocket mode of operation.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-1998-206639/REV1 , NAS 1.15:206639/REV1 , E-11106 , Propulsion Meeting; Jul 15, 1998 - Jul 17, 1998; Cleveland, OH; United States
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  • 9
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    Three Dimensional CFD Analysis of the GTX Combustor (2002)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The annular combustor geometry of a combined-cycle engine has been analyzed with three-dimensional computational fluid dynamics. Both subsonic combustion and supersonic combustion flowfields have been simulated. The subsonic combustion analysis was executed in conjunction with a direct-connect test rig. Two cold-flow and one hot-flow results are presented. The simulations compare favorably with the test data for the two cold flow calculations; the hot-flow data was not yet available. The hot-flow simulation Indicates that the conventional ejector-ramjet cycle would not provide adequate mixing at the conditions tested. The supersonic combustion ramjet flowfield was simulated with frozen chemistry model. A five-parameter test matrix was specified, according to statistical design-of-experiments theory. Twenty-seven separate simulations were used to assemble surrogate models for combustor mixing efficiency and total pressure recovery. Scramjet injector design parameters (injector angle, location, and fuel split) as well as mission variables (total fuel mass flow and freestream Mach number) were included in the analysis. A promising injector design has been identified that provides good mixing characteristics with low total pressure losses. The surrogate models can be used to develop performance maps of different injector designs. Several complex three-way variable interactions appear within the dataset that are not adequately resolved with the current statistical analysis.
    Keywords: Spacecraft Propulsion and Power
    Type: 26th JANNAF Airbreathing Propulsion Subcommittee Meeting; 1; 135-148; CPIA-Publ-713-Vol-1
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  • 10
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    Three Dimensional CFD Analysis of the GTX Combustor (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The annular combustor geometry of a combined-cycle engine has been analyzed with three-dimensional computational fluid dynamics. Both subsonic combustion and supersonic combustion flowfields have been simulated. The subsonic combustion analysis was executed in conjunction with a direct-connect test rig. Two cold-flow and one hot-flow results are presented. The simulations compare favorably with the test data for the two cold flow calculations; the hot-flow data was not yet available. The hot-flow simulation indicates that the conventional ejector-ramjet cycle would not provide adequate mixing at the conditions tested. The supersonic combustion ramjet flowfield was simulated with frozen chemistry model. A five-parameter test matrix was specified, according to statistical design-of-experiments theory. Twenty-seven separate simulations were used to assemble surrogate models for combustor mixing efficiency and total pressure recovery. ScramJet injector design parameters (injector angle, location, and fuel split) as well as mission variables (total fuel massflow and freestream Mach number) were included in the analysis. A promising injector design has been identified that provides good mixing characteristics with low total pressure losses. The surrogate models can be used to develop performance maps of different injector designs. Several complex three-way variable interactions appear within the dataset that are not adequately resolved with the current statistical analysis.
    Keywords: Aeronautics (General)
    Type: NASA/TM-2002-211572 , NAS 1.15:211572 , E-13357 , Combustion, Airbreathing Propulsion, Propulsion Systems Hazards, and Modelling; Apr 08, 2002 - Apr 12, 2002; Destin, FL; United States|Simulation Subcommittees Joint Meeting; Apr 08, 2002 - Apr 12, 2002; Destin, FL; United States
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