Robert S. Webster

A Generalized, Unstructured Interpolative Interface Method for Rotor-Stator Interactions

A generalized interpolative interface is developed to provide interdomain and intradomain coupling between multielement unstructured grids, including those that require highly stretched anisotropic meshes. The method centers around an extruded interpolative interface that does not require matched unstructured grids on the corresponding surfaces. In the context of rotor/stator interactions, this generalized interpolative interface is utilized between the rotor and stator sections, so that each can be solved within its own frame of reference. The interpolative interface supports the sliding of the rotor relative to the stator, such that variables are transmitted from one domain to the other accordingly. Coupling between the rotor and stator sections is accomplished via solution interpolations from mesh extrusions constructed from the interpolative interface. Further, the same interpolative interface technology is utilized to implement an axisymmetric boundary condition that also does not require matching surface grids on the periodic surfaces. Given that these axisymmetric boundary conditions intersect with the rotor/stator interface, special consideration is required in these areas, and the techniques applied are explained in this work in detail. This interpolative interface scheme is tested and applied via a performance mapping of the SDT2-R4 rotor/stator configuration as documented in Hughes, 1 simulated as a 2 rotor/5 stator passage. Results for this configuration are compared to available experimental data as well as full-wheel simulations of the same, in order to determine the eect of the axisymmetric interfaces on the overall solution. Overall agreement with experimental performance mappings is excellent.

A Generalized Interpolative Interface for Parallel Unstructured Field Solvers

A generalized interpolative interface is developed to provide a coupling between multielement unstructured grids potentially in relative motion, including those that require highly stretched anisotropic meshes. The method centers around an extruded interpolative interface that does not require matched unstructured grids on the corresponding surfaces. The method is also constructed such that is generally applicable to any data that must be transported between the adjacent solution domains; as such, the same mechanism can be used regardless of the underlying governing equations of the field. This feature is utilized in the current work by using dierent sets of governing equations for the presented solutions, as well as using the same interpolative procedure in the process of solving the turbulence models. Mesh extrusions are constructed from the interpolative interface, which allow for closure and the solution of control volumes that lie on the interface. For each extruded point, a corresponding virtual point is created in order to control the exact location at which the client data interpolations are performed. This allows interpolation from locations that are reflective of the physics of the problem. Special procedures, such as utilization of surface projections and parallel boundary layer displacement algorithms, are required for support of highly stretched anisotopic grids commonly used in the resolution of large gradients most common in fluid flow solvers. All algorithms used to extrude the interpolative surface, place virtual points, and interpolate for the client data must be parallelized for compatibility with modern parallel field solvers. To this end, fully general parallel mechanisms are implemented in order to transport data from its native storage to a possibly remote location. Interpolation also requires a parallel unstructured multielement search algorithm, which is a concerted eort by itself, and is the subject of an upcoming paper. This interpolative interface scheme is validated on a ramp immersed in supersonic flow, where the shock passes through the interface; comparisons with the theoretical solution for an oblique shock are excellent. Also, an unsteady pitching airfoil in which the airfoil is pitched with the aid of a surrounding interpolative interface is examined and compared against experimental data and a baseline case where the entire grid is pitched; comparisons here are excellent as well. The NASA SDT2-R4 turbofan stage is also presented as a demonstration of capability of the interpolative interface in real-world problems. Agreement with experimental data for demonstrated case at 100% speed is good.

A numerical investigation of s-duct flows with boundary-layer ingestion

The essence of this paper is to report the computational research in furthering NASAs experimental study on active flow control systems for a flush-mounted inlet with significant boundary layer ingestion. In conjunction with a NASA sponsored research grant, the aim is to further accumulate knowledge and insight on the effectiveness of flow control devices in reducing circumferential distortion. With the use of Computational Fluid Dynamics, this study seeks to validate wind tunnel results recorded by the NASA Langley Research Center. For this study, NASAs air jets configuration was chosen as the flow control device. The accuracy of a numerical solution is determined by the ability to match experimental values for mass flow rate through the S-duct, pressure measurements along the center line, and distortion coefficient at the fan-face location. The “Tenasi” code was able to calculate agreeable numerical solutions for the baseline case of no jet flow and for the case of jet flow with a low mass flow ratio. The high mass flow ratio cases have proven more difficult to validate. This paper reveals the steps taken in achieving early success in a computational validation of experimental data and discusses current actions being made to obtain complete valid numerical solutions.

An Artificial Compressibility Algorithm for Convective Heat Transfer

An implicit algorithm is presented for the unsteady three-dimensional incompressible Navier-Stokes equations including the thermal energy equation, formulated in a strongly coupled manner. Thermal buoyancy is addressed using the Boussinesq approximation. The method of artificial compressibility is introduced to cast the resulting equations into a time marching form. A finite volume discretization is applied for general unstructured grids with nonsimplical elements. Numerical flux formulations are presented for both Roe’s approximate Riemann solver and an HLLC approach. A backward Euler scheme is utilized for temporal discretization. The resulting implicit nonlinear equation is solved at each time step using an approximate Newton iteration algorithm. All boundary conditions are applied implicitly. The resulting algorithm has been implemented in Tenasi, an in-house developed flow solver, and validated for canonical test cases from the three regimes (forced, natural, and mixed) of laminar heat transfer and one regime (mixed) of turbulent heat transfer. A final application of the current algorithm is the prediction of forced turbulent heat transfer in a two-pass ribbed turbulator. All computed solutions are in close agreement with analytical solutions, other benchmark simulations, or experimental data.

Effect of Casing and Tip Modifications on the Performance of an Axial Flow Stage

This paper is a presentation of a computational study focused on modifications to the rotor blade tip and casing end wall for the purpose of enhancing the performance and increasing the stall margin of a model turbofan stage. Grooves in the casing outboard of the rotor blade tips, as well as a very small ring protruding inward from the casing just upstream of the leading edge and downstream of the trailing edge of the rotor blade constitute the modifications that will be examined. The results will be compared to experiment as well as simulations of the baseline configuration in order to demonstrate the degree of effectiveness of the respective modifications. It is felt that this work serves as a good starting point for further investigations into possible improvements of performance and stability for both fan and compressor stages.

Effect of Casing Groove Locations on the Performance of an Axial Flow Stage

This paper is a presentation of a computational study focused on modifications to the circumferential groove casing treatments for the purpose of enhancing the performance and increasing the stall margin of a model turbofan stage. An attempt to understand the combined behavior of a five-groove configuration leads to qualitative and quantitative analyses of the individual groove configurations. The following conclusions can be drawn from this analysis: the first is that the blockage should be shifted towards the downstream and pressure side of the rotor blade for the groove to be effective. The second is that the reduction of the incidence angle, the improved radial transport of axial momentum, and the reduction of tip blade loading are not found to be strongly related to the stall margin improvement for this fan stage. It is also found that the stall pattern is changed by the presence of grooves in this configuration.

Computational Simulation of the Fan and Low-pressure Compressor Stages of the Energy Efficient Engine

The capability of computational resources (processor speed, memory capacity, mass storage, etc.) has allowed simulations of complex fluid dynamic processes to become relatively commonplace. Even so, large-scale turbomachinery problems remain somewhat of a challenge for a number of reasons. This paper will show that reasonable solutions can be obtained for a problem that is representative of the highly unsteady aerodynamic environment that results from the interaction of multiple rotating and stationary rows of blades and vanes. A single passage of the fan stage and two passages of the low-pressure “boost” compressor stage for this machine are used in the current simulations so as to keep the computational size to a reasonable level. Adjustments in blade and vane counts are necessary, but the resulting performance values are still reasonable. Since this is a summary of the initial results, performance of the components will be the focus.

A Generalized Axisymmetric Boundary Condition Method for Parallel Unstructured Field Solvers

A generalized interpolative interface is developed to provide an axisymmetric boundary condition for multielement unstructured field solvers, including those that require highly stretched anisotropic meshes. The method centers around an extruded interpolative interface that does not require matched unstructured grids on the corresponding axisymmetric surfaces. Arc-based mesh extrusions are constructed from each periodic interface, which allow for closure and the solution of control volumes that lie on the interface. For each extruded point, a corresponding virtual point is created in order to control the exact location at which the client data interpolations are performed; for axisymmetric surfaces, the virtual points are rotated and placed inside of the opposite side of the domain. Special procedures, such as utilization of surface projections and parallel boundary layer displacement algorithms, are required for support of highly stretched anisotropic grids commonly used in the resolution of boundary layers in fluid flow solvers. All algorithms used to extrude the interpolative surface, place virtual points, and interpolate for the client data must be parallelized for compatibility with modern parallel field solvers. No restrictions are to be placed on the subdomain decomposition. To this end, fully general parallel mechanisms are implemented in order to transport data from its native storage to a possibly remote location. This overall axisymmetric boundary condition scheme is implemented in the Tenasi code for testing. Interpolation requires a parallel unstructured multielement search algorithm, which is a concerted effort by itself, and is the subject of an upcoming paper. This axisymmetric interface scheme is validated on an empty rotor passage, as well as on a Rotor 37 standard test case at full design speed. For these simulations, the two equation Menter SST [1] turbulence model is utilized. Profiles of the relative Mach number aft of the blade and pressure ratio data match very well with experimental results, demonstrating the validity of the proposed approach.Copyright

Testing Protruding Studs As a Form of Casing Treatment on a Transonic Turbofan: A Computational Study

This paper presents the findings of an ongoing CFD study of using protruding studs as a form of casing treatment on a transonic turbofan stage. Simulations have been performed on the subject turbomachine with and without the casing treatment in order to validate computations with available experimental results and to compute any difference in performance. The results of the simulations with the casing treatment suggest that protruding studs have the potential to extend the stall margin of the turbofan while resulting in a slight reduction in pressure rise and efficiency. From the use of an initial configuration of studs, the computed increase in stall margin based on mass flow rate was 5.46%, and the greatest decrease in pressure ratio and adiabatic efficiency were 0.25% and 1.59%, respectively. Flowfield visualizations of simulations at computed near-stall conditions without casing treatment show regions of low momentum flow near the casing in the rotor blade passage, and low momentum regions near the hub in the stator section. Visualization from simulations with casing treatment at computed near-stall conditions show a large blockage imposed by the studs in the rotor blade passage, and a low momentum region near the casing in the stator section. Computed performance maps obtained from using other configurations of studs suggest that further increase in stall margin is possible at other levels of protrusion of the studs.© 2017 ASME

Numerical Simulation of Reacting and Non-Reacting Nozzle Flows

The purpose of this paper is to present the results of several numerical studies of internal nozzle flows. The simulated flow fields are of non-reacting, ideal-gas flows, as well as flows with chemical reactions of the combustion products passing through the nozzle. The combustion process itself is not being simulated. The simulations are conducted using a number of flow solvers contained within the in-house suite of flow solvers collectively referred to as Tenasi. The older variants of these flow solvers are serial, structured-grid codes which allow for multiple sub-domain decomposition with arbitrary block-to-block connectivity. The newer code is a parallel, unstructured-grid solver that has been in a state of continual, and ongoing, development over the past seven years; sub-domain decomposition for this solver is completely arbitrary. Although there is noticeable difference in the code structure of the older and newer versions, the basic numerical algorithms are the same. Results from both the old and new solvers are compared to each other and to experiment for the non-reacting case. For a hypothetical reacting flow case, results from both the old and new solvers are compared to each other and to the results from the rocket engine analysis code, TDK, which is widely used throughout the industry. Comparison of these results to experiment (non-reacting) and TDK results (reacting) serves the purpose of validating the Tenasi solvers for nozzle flow applications.