Douglas L. Sondak

Simulation of Vortex Shedding in a Turbine Stage

Vortex shedding in a turbomachine blade row is affected by the passing of blades in the adjacent downstream blade row, but these effects have not been examined in the literature. A series of flow simulations has been performed to study vortex shedding in a turbine stage, and to quantify the blade interaction effects on the unsteady pressure response. The numerical issues of spatial order of accuracy and the use of Newton subiterations were investigated first. Second-order spatial accuracy was shown to be inadequate to model the shedding frequency response and time-averaged base pressure accurately. For the small time step employed for temporal accuracy, Newton iterations were shown to be unnecessary. The effects of the adjacent blade row were examined by comparing the shedding frequency response for the stage simulations to the response for isolated cascades. The vane shedding was shown to occur exactly on a series of harmonics of the blade passing frequency for the stage case, compared to a single predominant frequency for the isolated cascade. Losses were also examined in the wake region. It was shown that close to the trailing edge, losses were mainly due to wake mixing. Farther downstream of the trailing edge, losses were predominantly due to the trailing edge shock wave.

Simulation of Coupled Unsteady Flow and Heat Conduction in Turbine Stage

A three-dimensional, unsteady, compressible, e nite difference Navier ‐Stokes solver has been coupled with a three-dimensional, unsteady, e nite difference conduction heat-transfer solver to study conjugate heat-transfer problems in turbomachinery. The heat-transfer solver was validated by computing unsteady heat transfer in a cylinderandcomparingtheresultswithananalyticalsolution.Thecodewasthenappliedtoahigh-pressureturbine stage, typical of those found in modern high-bypassturbofan engines, with a nonuniform inlettemperatureproe le. The unsteady temperature e eld of a rotor blade, both at the surface and within the blade, has been examined in detail. The surface-temperature results have also been compared with those from a e ow simulation in which the blade surfaces were assumed to be adiabatic, demonstrating the need for the coupled approach.

A SURVEY OF HOT STREAK EXPERIMENTS AND SIMULATIONS

Experimental and computational data have shown that the flow exiting gas-turbine combustors can contain large circumferential and radial temperature non-uniformities. The temperature non-uniformities, or hot streaks, can have a significant impact on the performance and durability of first-stage turbine airfoils. This paper contains a survey of the hot streak experiments and simulations that have been performed during the last two decades, and the impact they have had on the design of high-pressure turbine stages.

VORTEX SHEDDING IN A TURBINE CASCADE

Flow simulations have been performed to examine vortex shedding in a turbine cascasde. Four progressively finer grids were used, and it was found that a very fine grid is required to accurately capture the unsteady flow in the trailing edge region. The predominant shedding frequency in the simulations was in good agreement with the experimental result. The simulations showed that the well-known uniform time-averaged pressure in the base pressure region is composed of the time average of two individual uniform pressure regions oscillating out of phase.

FULL-ANNULUS SIMULATIONS OF AIRFOIL CLOCKING IN A 1-1/2 STAGE AXIAL COMPRESSOR

Axial compressors have inherently unsteady flow fields because of relative motion between rotor and stator airfoils. This relative motion leads to viscous and inviscid (potential) interactions between blade rows. As the number of stages increases in a turbomachine, the buildup of convected wakes can lead to progressively more complex wake/wake and wake/airfoil interactions. Variations in the relative circumferential positions of stators or rotors can change these interactions, leading to different unsteady forcing functions on airfoils and different compressor efficiencies. In addition, as the Mach number increases the interaction between blade rows can be intensified due to potential effects.It has been shown, both experimentally and computationally, that airfoil clocking can be used to improve the efficiency and reduce the unsteadiness in multiple-stage axial turbomachines with equal blade counts in alternate blade rows. While previous investigations have provided an improved understanding of the physics associated with airfoil clocking, more research is needed to determine if airfoil clocking is viable for use in modern gas-turbine compressors. This paper presents the results of a combined experimental/computational research effort to study the physics of airfoil clocking in a high-speed axial compressor. Computational simulations have been performed for eight different clocking positions of the stator airfoils in a 1-1/2 stage high-speed compressor. To accurately model the experimental compressor, full-annulus simulations were conducted using 34 IGV, 35 rotor and 34 stator airfoils. It is common practice to modify blade counts to reduce the computational work required to perform turbomachinery simulations, and this approximation has been made in all computational clocking studies performed to date. A simulation was also performed in the present study with 1 inlet guide vane, 1 rotor airfoil, and 1 stator airfoil to model blade rows with 34 airfoils each in order to examine the effects of this approximation. Time-averaged and unsteady data (including performance and boundary layer quantities) were examined. The predicted results indicate that simulating the full annulus gives better qualitative agreement with the experimental data, as well as more accurately modeling the interaction between adjacent blade rows.Copyright

GENERAL EQUATION SET SOLVER FOR COMPRESSIBLE AND INCOMPRESSIBLE TURBOMACHINERY FLOWS

Turbomachines for pro pulsion applications operate with many different working fluids and flow conditions. The flow may be incompressible, such as in the liquid hydrogen pump in a rocket engine, or supersonic, such as in the turbine that may drive the hydrogen pump. Separate codes have traditionally been used to solve incompressible and compressible flows. The General Equation Set (GES) method can be used to solve both incompressible and compressible flows, and it is not restricted to perfect gases, as are many compressible -flow turbomachinery solvers. It also contains preconditioning for acceleration of steady -state simulations or rapid elimination of initial transients in unsteady simulations. An unsteady GES turbomachinery flow solver has been developed and applied to flow s through turbomachinery components operating in both liquids and gases. It has been shown to be a versatile technique for solving a wide range of turbomachinery flow problems. NOMENCLATURE

Effects of Tip Clearance on Hot Streak Migration in a High-Subsonic Single-Stage Turbine

Experimental data have shown that combustor temperature nonuniformities can lead to the excessive heating of first-stage rotor blades in turbines. This heating of the rotor blades can lead to thermal fatigue and degrade turbine performance. The results of recent studies have shown that variations in the circumferential location, or clocking, of the first-stage vane airfoils can be used to minimize the adverse effects of the hot streaks due to the hot fluid mixing with the cooler fluid contained in the vane wake. In addition, the effects of the hot streak/airfoil count ratio on the heating patterns of turbine airfoils have been quantified. In the present investigation, three-dimensional unsteady Navier-Stokes simulations have been performed for a single-stage high-pressure turbine geometry operating in high subsonic flow to study the effects of tip clearance on hot streak migration. Baseline simulations were initially performed without hot streaks to compare with the experimental data. Two simulations were then performed with a superimposed combustor hot streak; in the first the tip clearance was set at the experimental value, while in the second the rotor was allowed to scrape along the outer case (i.e., the limit as the tip clearance goes to zero). The predicted results for the baseline simulations show good agreement with the available experimental data. The simulations with the hot streak indicate that the tip clearance increases the radial spreading of the hot fluid, and increases the integrated rotor surface temperature compared to the case without tip clearance.

Experimental and Numerical Investigation of Airfoil Clocking and Inner-Blade-Row Gap Effects on Axial Compressor Performance

An experimental and computational research program has been performed to study the effects of stator indexing (clocking) in a high-speed axial compressor stage with an inlet guide vane. The work is directed at determining the basic flow mechanisms causing the performance variations associated with airfoil clocking. A comparison of the results obtained for two axial gaps between the rotor and the stator has been carried out to determine the role of potential interaction between blade rows as the airfoils are clocked.

Effect of Rotor Wake Structure on Fan Interaction Tone Noise

The prediction of noise produced by fan-wake interaction with the fan exit guide vane (FEGV) is studied. The acoustic response of the FEGV is computed using LINFLUX, a three-dimensional, frequency-domain, linearized Euler solver for turbomachinery. The research focuses on tonal noise predictions for the advanced ducted propulsor (ADP) and the source diagnostic test (SDT) scaled turbofan rigs. The sensitivity of the 2BPF noise prediction to the inow specication is quantied. Inow models are generated based on both experimental and computational fan-wake data. The computational data are provided by four dierent Reynolds-Averaged Navier Stokes (RANS) computational uid dynamic (CFD) solutions. When compared to experimental results, the computations provide comparable mean ow solutions but overpredict the wake decit. The CFD solutions dier more for higher wheel speed cases, especially in the tip region. It is shown that inputs generated from the various experimental and CFD data also dier, but that these dierences weakly impact the prediction of the sound power level (PWL) at the lower wheel speed. At the higher wheel speeds, where the dierences in the inputs near the tip are greatest, the PWL reects this dierence. For a given wheel speed, the predictions using input values based on the dierent CFD data and the experimental wake data agree to within 5 dB. However, they vary from the experimentally measured acoustic PWLs by up to 14 dB.

Full and Partial Admission Performance of the Simplex Turbine

The turbines used in rocket-engine applications are often partial-admission turbines, meaning that the flow enters the rotor over only a portion of the annulus. These turbines have been traditionally analyzed, however, assuming full-admission characteristics. This assumption enables the simulation of only a portion of the 360-deg annulus with periodic boundary conditions applied in the circumferential direction. Whereas this traditional approach to simulating the flow in partial-admission turbines significantly reduces the computational requirements, the accuracy of the solutions has not been evaluated or compared to partial-admission data. In the current investigation, both full-admission and partial-admission three-dimensional unsteady Navier‐Stokes simulations were performed fo ra partial-admission turbine designed and tested at NASA Marshall Space Flight Center. The results indicate that the partial-admission nature of the turbine should be included in simulations to properly predict the performance and flow unsteadiness of the turbine.