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  • 1
    Electronic Resource
    Electronic Resource
    NUMERICAL UNCERTAINTIES IN TRANSONIC FLOW CALCULATIONS FOR AEROFOILS WITH BLUNT TRAILING EDGES (1997)
    Chichester : Wiley-Blackwell
    International Journal for Numerical Methods in Fluids 24 (1997), S. 355-373 
    Details
    ISSN: 0271-2091
    Keywords: computational fluid dynamics ; transonic airfoils ; numerical uncertainty ; Engineering ; Numerical Methods and Modeling
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Mechanical Engineering, Materials Science, Production Engineering, Mining and Metallurgy, Traffic Engineering, Precision Mechanics
    Notes: Numerical uncertainties are quantified for calculations of transonic flow around a divergent trailing edge (DTE) supercritical aerofoil. The Reynolds-averaged Navier-Stokes equations are solved using a linearized block implicit solution procedure and mixing-length turbulence model. This procedure has reproduced measurements around supercritical aerofoils with blunt trailing edges that have shock, boundary layer and separated regions. The present effort quantifies numerical uncertainty in these calculations using grid convergence indices which are calculated from aerodynamic coefficients, shock location, dimensions of the recirculating region in the wake of the blunt trailing edge and distributions of surface pressure coefficients. The grid convergence index is almost uniform around the aerofoil, except in the shock region and at the point where turbulence transition was fixed. The grid convergence index indicates good convergence for lift but only fair convergence for moment and drag and also confirms that drag calculations are more sensitive to numerical error. © 1997 by John Wiley and Sons, Ltd.
    Additional Material: 15 Ill.
    Type of Medium: Electronic Resource
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    Paper (German National Licenses)
  • 2
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    Numerical simulation of shock waves and their interaction in a supersonic rocket engine operating at different conditions (2004)
    Springer
    Details
    Publication Date: 2004-06-01
    Print ISSN: 1343-8875
    Electronic ISSN: 1875-8975
    Topics: Computer Science
    Published by Springer
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    PAPER CURRENT
  • 3
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    Computational fluid dynamics simulation of a supersonic rocket thruster flow compared with experimental data (2004)
    Springer
    Details
    Publication Date: 2004-09-01
    Print ISSN: 1343-8875
    Electronic ISSN: 1875-8975
    Topics: Computer Science
    Published by Springer
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    PAPER CURRENT
  • 4
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    Supersonic Rocket Thruster Flow Predicted by Numerical Simulation (2004)
    In:  CASI
    Details
    Publication Date: 2018-06-02
    Description: Despite efforts in the search for alternative means of energy, combustion still remains the key source. Most propulsion systems primarily use combustion for their needed thrust. Associated with these propulsion systems are the high-velocity hot exhaust gases produced as the byproducts of combustion. These exhaust products often apply uneven high temperature and pressure over the surfaces of the appended structures exposed to them. If the applied pressure and temperature exceed the design criteria of the surfaces of these structures, they will not be able to protect the underlying structures, resulting in the failure of the vehicle mission. An understanding of the flow field associated with hot exhaust jets and the interactions of these jets with the structures in their path is critical not only from the design point of view but for the validation of the materials and manufacturing processes involved in constructing the materials from which the structures in the path of these jets are made. The hot exhaust gases often flow at supersonic speeds, and as a result, various incident and reflected shock features are present. These shock structures induce abrupt changes in the pressure and temperature distribution that need to be considered. In addition, the jet flow creates a gaseous plume that can easily be traced from large distances. To study the flow field associated with the supersonic gases induced by a rocket engine, its interaction with the surrounding surfaces, and its effects on the strength and durability of the materials exposed to it, NASA Glenn Research Center s Combustion Branch teamed with the Ceramics Branch to provide testing and analytical support. The experimental work included the full range of heat flux environments that the rocket engine can produce over a flat specimen. Chamber pressures were varied from 130 to 500 psia and oxidizer-to-fuel ratios (o/f) were varied from 1.3 to 7.5.
    Keywords: Aircraft Propulsion and Power
    Type: Research and Technology 2003; NASA/TM-2004-212729
    Format: application/pdf
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  • 5
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    Numerical Simulation of the RTA Combustion Rig (2005)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The Revolutionary Turbine Accelerator (RTA)/Turbine Based Combined Cycle (TBCC) project is investigating turbine-based propulsion systems for access to space. NASA Glenn Research Center and GE Aircraft Engines (GEAE) planned to develop a ground demonstrator engine for validation testing. The demonstrator (RTA-1) is a variable cycle, turbofan ramjet designed to transition from an augmented turbofan to a ramjet that produces the thrust required to accelerate the vehicle from Sea Level Static (SLS) to Mach 4. The RTA-1 is designed to accommodate a large variation in bypass ratios from sea level static to Mach 4 conditions. Key components of this engine are new, such as a nickel alloy fan, advanced trapped vortex combustor, a Variable Area Bypass Injector (VABI), radial flameholders, and multiple fueling zones. A means to mitigate risks to the RTA development program was the use of extensive component rig tests and computational fluid dynamics (CFD) analysis.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-2005-213899 , E-15270 , 40th Combustion, 28th Airbreathing Propulsion, 22nd Propulsion Systems Hazards, 4th Modeling and Simulations Joint Subcommittees Meetings; Jun 13, 2005 - Jun 17, 2005; Charleston, SC; United States
    Format: application/pdf
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  • 6
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    Simulation of the Flow Field Associated with a Rocket Thruster Having an Attached Panel (2003)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Two-dimensional inviscid and viscous numerical simulations are performed to predict the flow field induced by a H2-O2 rocket thruster and to provide insight into the heat load on the articles placed in the hot gas exhaust of the thruster under a variety of operating conditions, using the National Combustion Code (NCC). The simulations have captured physical details of the flow field, such as the plume formation and expansion, formation of the shock waves and their effects on the temperature and pressure distributions on the walls of the apparatus and the flat panel. Comparison between the computed results for 2-D and adiabatic walls and the related experimental measurements for 3-D and cooled walls shows that the results of the simulations are consistent with those obtained from the related rig tests.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NASA/TM-2003-212347 , E-13935 , NAS 1.15:212347 , 2003 Fluids Engineering Division Summer Meeting; Jul 06, 2003 - Jul 10, 2003; Honolulu, HI; United States
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  • 7
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    Three-dimensional blade vortex interactions (1991)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: A three-dimensional time dependent Navier-Stokes analysis was applied to the rotor blade vortex interaction problem. The numerical procedure is an iterative implicit procedure using three point central differences to represent spatial derivatives. A series of calculations were made to determine the time steps, pseudo-time steps, iterations, artificial dissipation level, etc. required to maintain a nondissipative vortex. Results show the chosen method to have excellent non-dissipative properties provided the correct parameters are chosen. This study was used to set parameters for both two- and three-dimensional blade vortex interaction studies. The case considered was the interaction between a vortex and a NACA0012 airfoil. The results showed the detailed physics during the interaction including the pressure pulse propagating from the blade. The simulated flow physics was qualitatively similar to that experimentally observed. The BVI phenomena is the result of the buildup and violent collapse of the shock waves and local supersonic pockets on the blade surfaces. The resulting pressure pulse build-up appears to be centered at the blade leading edge.
    Keywords: AERODYNAMICS
    Type: AHS International Specialists'' Meeting on Rotorcraft Basic Research; Mar 25, 1991 - Mar 27, 1991; Atlanta, GA; United States
    Format: text
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  • 8
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    The Effect of Spray Initial Conditions on Heat Release and Emissions in LDI CFD Calculations (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: The mass and velocity distribution of liquid spray has a primary effect on the combustion heat release process. This heat release process then affects emissions like nitrogen oxides (NOx) and carbon monoxide (CO). Computational Fluid Dynamics gives the engineer insight into these processes, but various setup options exist (number of droplet groups, and initial droplet temperature) for spray initial conditions. This paper studies these spray initial condition options using the National Combustion Code (NCC) on a single swirler lean direct injection (LDI) flame tube. Using laminar finite rate chemistry, comparisons are made against experimental data for velocity measurements, temperature, and emissions (NOx, CO).
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NASA/TM-2008-215422 , E-16591
    Format: application/pdf
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  • 9
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    NASA National Combustion Code Simulations (2001)
    In:  Other Sources
    Details
    Publication Date: 2019-07-18
    Description: A systematic effort is in progress to further validate the National Combustion Code (NCC) that has been developed at NASA Glenn Research Center (GRC) for comprehensive modeling and simulation of aerospace combustion systems. The validation efforts include numerical simulation of the gas-phase combustor experiments conducted at the Center for Turbulence Research (CTR), Stanford University, followed by comparison and evaluation of the computed results with the experimental data. Presently, at GRC, a numerical model of the experimental gaseous combustor is built to simulate the experimental model. The constructed numerical geometry includes the flow development sections for air annulus and fuel pipe, 24 channel air and fuel swirlers, hub, combustor, and tail pipe. Furthermore, a three-dimensional multi-block, multi-grid grid (1.6 million grid points, 3-levels of multi-grid) is generated. Computational simulation of the gaseous combustor flow field operating on methane fuel has started. The computational domain includes the whole flow regime starting from the fuel pipe and the air annulus, through the 12 air and 12 fuel channels, in the combustion region and through the tail pipe.
    Keywords: Aircraft Propulsion and Power
    Type: NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; 11
    Format: text
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  • 10
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    Numerical Prediction of Non-Reacting and Reacting Flow in a Model Gas Turbine Combustor (2005)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The three-dimensional, viscous, turbulent, reacting and non-reacting flow characteristics of a model gas turbine combustor operating on air/methane are simulated via an unstructured and massively parallel Reynolds-Averaged Navier-Stokes (RANS) code. This serves to demonstrate the capabilities of the code for design and analysis of real combustor engines. The effects of some design features of combustors are examined. In addition, the computed results are validated against experimental data.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NASA/TM-2005-213898 , GT2004-53496 , E-15269 , Turbo Expo 2004; Jun 14, 2004 - Jun 17, 2004; Vienna; Austria
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