Stephen Guillot

Experimental Investigation of the Effects of a Moving Shock Wave on Compressor Stator Flow

Linear cascade testing was performed to simulate the flow conditions experienced by stator blades in an axial compressor with supersonic relative Mach numbers at the inlet to the downstream embedded rotors. Experiments were conducted in a transonic blow-down wind tunnel with a nominal inlet Mach number of 0.65. A single moving normal shock introduced at the exit of the stator cascade simulated the bow shock from a downstream rotor. The shock was generated using a shock tube external to the wind tunnel. Pressure measurements indicated that the stator matched its design intent loading, turning, and loss under steady flow conditions. Effects of the passing shock on the stator flowfield were investigated using shadowgraph photography and digital particle image velocimetry (DPIV). Measurements were taken with three different shock strengths. In each case, the passing shock induced a vortex around the trailing edge of the stator. The size and strength of these vortices were directly related to the shock strength. A suction side separation on the trailing edge of the stator was observed and found to correlate with the vortex blockage.

Effect of Loading Level and Distribution on LPT Losses

This paper reports the results of an experimental and analytical study dealing with the effect of loading level and distribution on low-pressure turbine (LPT) blade performance. Only a single blade row is considered here, and the study is conducted in a stationary linear cascade that simulates the aero characteristics of a modern LPT design. The loading level and distribution are systematically varied by changing the number of blades (solidity), the stagger angle, and the unguided turning angle. The exit Mach number for this high-speed test is set at 0.64. The Zweifel number ranges from ∼ 1 (nominal lift) to ∼ 1.27 (high lift). The Reynolds number (based on chord and exit velocity) is varied from ∼70,000 to ∼350,000, a range that is broad enough to cover typical cruise and take-off conditions. While some data is taken near the end-walls, the primary focus of this study is on measurements at the mid-span. In addition to the profile loss, measurements include static pressure distribution on the blade surface (loading) and flow visualization. Data demonstrates increased suction side separation and consequent high losses as the loading level increases, the loading is moved aft, or the Reynolds number decreases. Three-dimensional CFD simulations, in conjunction with a turbulence transition model, corroborate these findings.Copyright

A Three-Dimensional CFD Design Study of a Circulation Control Inlet Guide Vane for a Transonic Compressor

A three-dimensional computational investigation was conducted to evaluate the performance of a circulation control inlet guide vane (IGV) designed for a transonic compressor test rig. The general configuration of the IGV is an uncambered airfoil designed to exploit the Coanda effect for flow turning through the use of a jet that exhausts along a curved trailing edge. Computational fluid dynamics (CFD) was used in conjunction with an optimization approach to design an IGV that balances performance potential with manufacturing cost and complexity. Optimization parameters include geometry specifications such as jet height and trailing edge radius, as well as operational parameters such as jet supply pressure and inlet Mach number. The potential for positive flow turning was shown for all geometries examined. Operational goals such as increased turning at lower off-design inlet Mach numbers were demonstrated, and an optimum configuration was established.Copyright

2D Parametric Study Using CFD of a Circulation Control Inlet Guide Vane

Circulation control inlet guide vanes (IGVs) may provide significant benefits over current IGVs that employ mechanical means for flow turning. This paper presents the results of a two-dimensional computational study on a circulation control IGV that takes advantage of the Coanda effect for flow vectoring. The IGV in this study is an uncambered airfoil that alters circulation around itself by means of a Coanda jet that exhausts along the IGV’s trailing edge surface. The IGV is designed for an axial inlet flow at a Mach number of 0.54 and an exit flow angle of 11 degrees. These conditions were selected to match the operating conditions of the 90% span section of the IGV of the TESCOM compressor rig at the Compressor Aero Research Laboratory (CARL) located at Wright-Patterson AFB, the hardware that is being used as the baseline in this study. The goal of the optimization was to determine the optimal jet height, trailing edge radius, and supply pressure that would meet the design criteria while minimizing the mass flow rate and pressure losses. The optimal geometry that was able to meet the design requirements had a jet height of h/Cn = 0.0057 and a trailing edge Radius R/Cn = 0.16. This geometry needed a jet to inflow total pressure ratio of 1.8 to meet the exit turning angle requirement. At this supply pressure ratio the mass flow rate required by the flow control system was 0.71 percent of the total mass flow rate through the engine. The optimal circulation control IGV had slightly lower pressure losses when compared with a reference cambered IGV.Copyright

PIV Investigation of the Flow Field in a Transonic Compressor Cascade with a Moving Shock Wave

The unsteady flow field produced during the interaction of a shock wave with stator blades is investigated in a linear cascade using particle image velocimetry (PIV). The study investigates the interaction that occurs when shock waves traveling with the rotor blades in axial transonic compressors interact with upstream stator blades. This interaction produces unsteady phenomena such as vortices and separation that induce blockage and losses. Flow visualization and PIV data, synchronized with shock-wave-passage locations provide details of the flow field in various areas of the cascade passage. The experiments are conducted in a transonic blow-down wind tunnel with a nominal inlet Mach number of 0.65. A single moving normal shock is generated using a shock tube external to the wind tunnel, and this shock is introduced at the exit of the stator cascade to simulate the bow shock from a downstream rotor. PIV instantaneous measurements are made for three different shock strengths at various regions of interest and are synchronized with various instants of the shock passage. In each case, the passing shock induces a vortex of varying size and strength around the trailing edge of the stator. The flow pattern includes the disruption and recovery of the transonic free stream, shock waves, vortex flow, vortex blockage, suction-side separation, spiraling arms, secondary vortices, and endwall clearance flows.

An Experimental Investigation of the Performance Impact of Swirl on a Turbine Exhaust Diffuser/Collector for a Series of Diffuser Strut Geometries

A comprehensive experimental investigation was initiated to evaluate the aerodynamic performance of a gas turbine exhaust diffuser/collector for various strut geometries over a range of inlet angle. The test was conducted on a 1/12th scale rig developed for rapid and accurate evaluation of multiple test configurations. The facility was designed to run continuously at an inlet Mach number of 0.40 and an inlet hydraulic diameter-based Reynolds number of 3.4×105. Multi-hole pneumatic pressure probes and surface oil flow visualization were deployed to ascertain the effects of inlet flow angle and strut geometry. Initial baseline diffuser-only tests with struts omitted showed a weakly increasing trend in pressure recovery with increasing swirl, peaking at 14° before rapidly dropping. Tests on profiled struts showed a similar trend with reduced recovery across the range of swirl and increased recovery drop beyond the peak.Subsequent tests for a full diffuser/collector configuration with profiled struts revealed a rising trend at lower swirl when compared to diffuser-only results, albeit with a reduction in recovery. When tested without struts, the addition of the collector to the diffuser not only reduced the pressure recovery at all angles but also resulted in a shift of the overall characteristic to a peak recovery at a lower value of swirl. The increased operation range associated with the implementation of struts in the full configuration is attributed to the de-swirling effects of the profiled struts. In this case the decreased swirl reduces the flow asymmetry responsible for the reduction in pressure recovery attributed to the formation of a localized reverse-flow vortex near the bottom of the collector. This research indicates that strut setting angle and, to a lesser extent, strut shape can be optimized to provide peak engine performance over a wide range of operation.© 2015 ASME

2D CFD Studies Using Different Turbulence Models of a Circulation Control Inlet Guide Vane

The IGV in this study is an uncambered airfoil, designed to use the Coanda effect to achieve flow vectoring. For this internal flow problem, two isotropic turbulence models are compared to an anisotropic model. Good trend comparison was seen for turning angle and pressure loss performance characteristics given a nominal fixed trailing edge geometry. However, as the trailing edge geometry is modified in an attempt to increase turning, discrepancies become evident, possibly due to the effects of streamline curvature. As the trailing edge radius decreases, significant variations in the jet separation location are predicted, which translates directly into flow turning predictions. Further, one turbulence model was examined using a second CFD code to ensure software independence of the solutions.Copyright

Numerical Study of Embedded Supersonic Compressor Stages

A numerical case study of a multistage, highly-loaded, relative supersonic compressor is presented. The purpose of the investigation was to highlight the changing shock structure while throttling the compressor and to give insight into possible compressor instabilities. The computational fluid dynamic (CFD) study was conducted with the NASA code ADPAC, utilizing the mixing-plane assumption for the boundary condition between adjacent, relatively-rotating blade rows. A steady, five-blade-row, numerical simulation using the Baldwin-Lomax turbulence model was performed, creating several constant speed lines. The results show that the shock structure in the downstream rotor isolates the upstream rotor from the exit conditions until the shock detaches from the leading edge. The shock structure in the upstream rotor then moves, changing the conditions for the downstream rotor. This continues as the compressor is throttled until the shock in the upstream rotor detaches from the leading edge. CFD indicates that this event causes a rapid drop in the mass flow rate, creating a mismatch between stage-one and stage-two that results in compressor instability.Copyright