A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 43, ISSUE 20 - OCTOBER 07, 2005

02 AERODYNAMICS
Includes aerodynamics of flight vehicles, test bodies, airframe components and combinations, wings, and control surfaces.

Also includes aerodynamics of rotors, stators, fans, and other elements of turbomachinery.

For related information see also 34 Fluid Mechanics and Thermodynamics.

20050215111 NASA Langley Research Center, Hampton, VA, USA

Numerical Study of High-Temperature Jet Flow Using RANS/LES and PANS Formulations

Abdol-Hamid, Khaled S.; Elmiligui, Alaa; [2005]; 17 pp.; In English; 23rd AIAAApplied Aerodynamics Conference, 6-9 Jun. 2005, Toronto, Ontario, Canada Contract(s)/Grant(s): WBS 23-781-10-12 Report No.(s): AIAA Paper 2005-5092; Copyright; Avail.: CASI: A03, Hardcopy

Two multi-scale-type turbulence models are implemented in the PAB3D solver. The models are based on modifying the Reynolds Averaged Navier-Stokes (RANS) equations. The first scheme is a hybrid RANS/LES model utilizing the two-equation (k(epsilon)) model with a RANS/LES transition function dependent on grid spacing and the computed turbulence length scale. The second scheme is a modified version of the Partially Averaged Navier-Stokes (PANS) model, where the unresolved kinetic energy parameter (f(sub k)) is allowed to vary as a function of grid spacing and the turbulence length scale. This parameter is estimated based on a novel two-stage procedure to efficiently estimate the level of scale resolution possible for a given flow on a given grid for Partial Averaged Navier-Stokes (PANS). It has been found that the prescribed scale resolution can play a major role in obtaining accurate flow solutions. The parameter f(sub k) varies between zero and one and equal to one in the viscous sub layer, and when the RANS turbulent viscosity becomes smaller than the LES viscosity. The formulation, usage methodology, and validation examples are presented to demonstrate the enhancement of PAB3D's time-accurate and turbulence modeling capabilities. The accurate simulations of flow and turbulent quantities will provide valuable tool for accurate jet noise predictions. Solutions from these models are compared to RANS results and experimental data for high-temperature jet flows. The current results show promise for the capability of hybrid RANS/LES and PANS in simulating such flow phenomena. Author

Turbulence Models; Reynolds Averaging; Navier-Stokes Equation; Kinetic Energy; Gas Jets; High Temperature Gases

20050215169 NASA Glenn Research Center, Cleveland, OH, USA

Single-Stage, 3.4:1-Pressure-Ratio Aspirated Fan Developed and Demonstrated

Braunscheidel, Edward P.; Research and Technology 2003; May 2004; 4 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy

Researchers are constantly pursuing technologies that will increase the performance of gas turbine engines. The aspirated compressor concept discussed here would allow the compression system to perform its task with about one-half of the compressor blades. To accomplish this, the researchers applied boundary layer control to the blades, casing, and hub. This method of boundary layer control consisted of removing small amounts of air from the main flow path at critical areas of the compressor. This bleed air could be used by other systems such as engine cooling or could be re-injected into lower pressure areas that require air for enhanced performance.

This effort was initiated by the Massachusetts Institute of Technology (MIT) in response to a solicitation from the Defense Advanced Research Projects Agency (DARPA) who sought to advance research in flow control technology. The NASA Glenn Research Center partnered with MIT (principal investigator), Honeywell Aircraft Engines (cycle analysis, structural analysis, and mechanical design), and Pratt & Whitney (cycle analysis and aero-analysis) to conceptualize, design, analyze, build, and test the aspirated fan stage. The aero-design and aero-analysis of this fan stage were jointly executed by MIT and Glenn to minimize the amount of bleed flow needed and to maintain the highest efficiency possible (ref. 1). Mechanical design issues were complicated by the need to have a shrouded rotor with hollow blades, with rotor stress levels beyond the capabilities of titanium.

The high stress issues were addressed by designing a shroud that was filament wound with a carbon fiber/epoxy matrix, resulting in an assembly that was strong enough to handle the high stresses. Both the rotor (preceding photographs) and stator (following photograph) were fabricated in two halves and then bolted together at the hub and tip, permitting the bleed passages to be machined into each half before assembly. Author

Gas Turbine Engines; Compressor Blades; Boundary Layer Control; Air Flow; Design Analysis; Pressure Ratio

20050215181 NASA Glenn Research Center, Cleveland, OH, USA

Glow Discharge Plasma Demonstrated for Separation Control in the Low-Pressure Turbine

Ashpis, David e.; Hultgren, Lennart S.; Research and Technology 2003; May 2004; 4 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy

Flow separation in the low-pressure turbine (LPT) is a major barrier that limits further improvements of aerodynamic designs of turbine airfoils. The separation is responsible for performance degradation, and it prevents the design of highly loaded airfoils. The separation can be delayed, reduced, or eliminated completely if flow control techniques are used.

Successful flow control technology will enable breakthrough improvements in gas turbine performance and design. The focus of this research project was the development and experimental demonstration of active separation control using glow discharge plasma (GDP) actuators in flow conditions simulating the LPT. The separation delay was shown to be successful, laying the foundation for further development of the technologies to practical application in the LPT. In a fluid mechanics context, the term ‘flow control' means a technology by which a very small input results in a very large effect on the flow.

In this project, the interest is to eliminate or delay flow separation on LPT airfoils by using an active flow control approach, in which disturbances are dynamically inserted into the flow, they interact with the flow, and they delay separation. The disturbances can be inserted using a localized, externally powered, actuating device, examples are acoustic, pneumatic, or mechanical devices that generate vibrations, flow oscillations, or pulses. A variety of flow control devices have been demonstrated in recent years in the context of the external aerodynamics of aircraft wings and airframes, where the incoming flow is quiescent or of a very low turbulence level. However, the flow conditions in the LPT are significantly different because there are high levels of disturbances in the incoming flow that are characterized by high free-stream turbulence intensity. In addition, the Reynolds number, which characterizes the viscous forces in the flow and is related to the flow speed, is very low in the LPT passages. Author

Low Pressure; Gas Turbine Engines; Free Flow; Control Equipment; Active Control; Viscous Flow; Separated Flow; Magnetohydrodynamic Flow; Glow Discharges

20050215258 NASA Glenn Research Center, Cleveland, OH, USA

Flutter Stability of the Efficient Low Noise Fan Calculated

Bakhle, Milind A.; Srivastava, Rakesh; Research and Technology 2003; May 2004; 2 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy

The TURBO-AE aeroelastic code has been used to verify the flutter stability of the Efficient Low Noise Fan (ELNF), which is also referred to as the trailing-edge blowing fan. The ELNF is a unique technology demonstrator being designed and fabricated at the NASA Glenn Research Center for testing in Glenn's 9-by-15-Foot Low-Speed Wind Tunnel. In the ELNF, air can be blown out of slots near the trailing edges of the fan blades to fill in the wakes downstream of the rotating blades. This filling of the wakes leads to a reduction of the rotor-stator interaction (tone) noise that results from the interaction of wakes with the downstream stators.

The ELNF will demonstrate a 1.6-EPNdB1 reduction in tone noise through wake filling, without increasing the broadband noise. Furthermore, the reduced blade row interaction will decrease the possibility of forced response and enable closer spacing of blade rows, thus reducing engine length and weight.

During the design of the ELNF, the rotor blades were checked for flutter stability using the detailed aeroelastic analysis capability of the three-dimensional Navier-Stokes TURBOAE code. The aeroelastic calculations were preceded by steady calculations in which the blades were not allowed to vibrate. For each rotational speed, as the back-pressure was increased, the mass flow rate decreased, and the operating point moved along the constant speed characteristic (speed-line) from choke to stall as shown on the fan map.

The TURBO-AE aeroelastic analyses were performed separately for the first two vibration modes (bending and torsion) and covered the complete range of interblade phase angles or nodal diameters at which flutter can occur. The results indicated that the ELNF blades would not encounter flutter at takeoff conditions. The calculations were then repeated for a part-speed condition (70-percent rotational speed), and the results again showed no flutter in the operating region. On the fan map (shown), the predicted flutter point at part speed condition was located beyond the stall line, which means that the ELNF will not encounter flutter since it will never operate beyond the stall line. All the calculations done so far have been for the nonblowing case, and selected calculations will be repeated with air blowing from the trailing edge of the fan. Author

Flutter; Stability; Low Noise; Fan Blades; Rotor Aerodynamics; Interactional Aerodynamics; Trailing Edges; Blowing; Aerodynamic Noise

20050216557 Army Research Lab., Aberdeen Proving Ground, MD USA

Application of Computational Fluid Dynamics to a Monoplane Fixed-Wing Missile With Elliptic Cross Sections

Heavey, Karen; Sahu, Jubaraj; Jul. 1, 2005; 39 pp.; In English; Original contains color illustrations Report No.(s): AD-A437107; ARL-TR-3549; No Copyright; Avail.: CASI: A03, Hardcopy

This report describes a computational study undertaken to investigate the performance of the CFD++ flow solver for prediction of nonlinear aerodynamics of a complex finned missile using structured hexahedral and unstructured tetrahedral grids. A monoplane fixed-wing missile with elliptic cross sections provided a geometrically complex model. Numerical solutions were obtained for this configuration at supersonic speed for various roll orientations, angles of attack, and jaw angles. Steady-state solutions were obtained using a three-dimensional Reynolds-Averaged Navier-Stokes solver with a two-equation turbulence model. Numerical results show the qualitative features of the flow fields at various cross-sectional and streamwise positions along the computational model of the missile. Aerodynamic coefficients were extracted from the computed solutions and found to match well with the available experimental data for these configurations. These numerical results show the effectiveness of using computational fluid dynamics techniques to produce an accurate prediction of the aerodynamics of geometrically complete configurations. DTIC

Aerodynamic Characteristics; Computational Fluid Dynamics; Fixed Wings; Missiles; Monoplanes

Source: NASA.