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A computational fluid dynamics-based analysis has been performed to assess propulsive efficiency benefits of a boundary layer ingesting propulsion system. This analysis is based on a full-scale con...

Computational studies have been carried out to assess engine operability for an Ultra High Bypass (UHB) engine installation on a Hybrid Wing Body (HWB) aircraft. In this installation the engines are mounted on the upper aft fuselage. At certain aircraft operating conditions at the boundary of the flight envelope, such as 0.2 flight Mach with very high angle of attack or takeoff with 30 knot crosswind, the installed engine flowfield and inlet distortion may differ from what has been typical for conventional tube+wing aircraft with underwing podded engines. Definition of inlet distortion at these key operating conditions for this configuration was provided by the airframer. The engine response to the inlet distortion was computed based on a coupled inlet-fan CFD simulation. Separate analyses were performed to assess performance, fan operability, core engine operability, and fan blade vibratory stress.

An analytical model was developed to investigate the dynamics of nonpremixed flames in a shear layer established between a mainstream flow of fuel-rich combustion products and a porous surface with an angled injection of air. In the model, a one-step overall chemical reaction was employed, together with boundary layer conservation equations solved using similarity solutions. Parametric studies were performed to understand the effects of equivalence ratio, temperature, and mass flow rate of the fuel and air streams on the flame standoff distance, surface temperature, and heat flux at the surface. The analytical model predictions were compared with computational fluid dynamics results obtained using the FLUENT commercial code for both the laminar and the turbulent flow models. Qualitative agreement in surface temperature was observed. Finally, the flame stability limits predicted by the model were compared with available experimental data and found to agree qualitatively, as well.

A Computational Fluid Dynamics (CFD) based analysis has been performed to assess propulsive efficiency benefits of a Boundary Layer Ingesting (BLI) propulsion system. This analysis is based on a full-scale configuration at relevant flight conditions defined in previous vehicle-level system studies and component designs, and includes the full propulsor geometry. A method to calculate the net thrust, equivalent power, and propulsive efficiency of BLI propulsion systems is presented, and applied to the CFD solutions. It was found that the BLI system results in a propulsive efficiency benefit of 4.2-4.5% over that of a conventional propulsion system for the case of a fan designed primarily for aerodynamic performance. These results compare favorably to those predicted by engine cycle-based system studies and BLI theory, and do not account for anticipated additional nacelle drag reduction benefits. It is recognized that future design work will be required to impart aeromechanics capability to the fan, and thus that the subject results provide an estimate of the upper benefit bound for a full-scale BLI propulsor in a relevant environment.

An adoptively refined quadtree grid method for the numerical solution of the incompressible Navier-Stokes equations is presented. A pressure-based scheme with allocated primitive variables is used as the solution algorithm. A process of grid refinement and flow solution is repeated until a sufficiently resolved solution is obtained. The present method has been applied to a variety of test cases. The results show that an adoptively refined quadtree grid can yield a better grid distribution over the flow, therefore yielding a more accurate solution as well as an improved convergence rate than a structured grid with a similar number of grid points.

A combine d computational and experimental study has been carried out to assess a novel concept for airfoil separation control. This concept involves the use of a thin, counter flowing wall jet, injected on the pressure surface of an airfoil in the counter -streamwi se direction. The wall jet provides several boundary layer energization benefits beyond what can be accomplished using conventional streamwise blowing on the airfoil suction surface. The computational portion of the study was carried out using the Comput ational Fluid Dynamics (CFD) CFL3D Navier -Stokes code for a two -dimensional airfoil operating over a Mach number range of 0.2 to 0.4. The calculations were used to study the effects of the counter -flow injection slot geometry and placement, blowing rate, and details of the flowfield with and without separation control. The CFD calculations indicated such promising results that, at the conclusion of this portion of the work, a validation experiment was carried out in the United Technologies Research Center Large Subsonic Wind Tunnel using a two dimensional section of a full -scale Sikorsky SSCA09 airfoil. The CFD results were utilized to guid e the set -up of the experiment, including slot size, location, and plenum pressure . Results of this study indicate t hat the counter -flowing wall jet concept provides dramatic separation control for an airfoil operating at normally stalled, high -angle -of -attack conditions. This flow control was achieved using relatively low blowing momentum coefficients, on the order of 0.01 and lower. Maximum lift coefficient was increased by approximately 25%, and stall angle was increased on the order of 4 – 5 degrees relative to the baseline airfoil operating without flow control. Excellent agreement was obtained between predicted and measured airfoil lift performance in the presence of the counter flowing wall jet injection, and CFD results indicate that this approach requires substantially reduced amounts of blowing air when compared to conventional, streamwise upper -surface slot injection. CFD calculations were also carried out to perform an initial assessment of the ability to extend the counter -flow fluid injection concept to three -dimensional flows relevant to gas turbine engine inlet geometries and operating conditions .

Ceramic matrix composite (CMC) turbine vanes, due to their high temperature capability, allow significantly higher firing temperatures with minimal cooling. Turbine vanes were designed for a gas turbine engine with special attention to attachment methods that minimize thermal stresses due to large differences in coefficients of thermal expansion between the CMC airfoil and metal platforms. Detailed aerodynamic, thermal and structural analyses were performed to ensure component reliability. The paper describes the component design, analysis, fabrication, and rig testing of a silicon carbide fiber reinforced silicon carbide matrix (SiC/SiC) turbine vane.Copyright © 2005 by ASME