The Effect of Inlet Guide Vanes Wake Impingement on the Flow Structure and Turbulence Around a Rotor Blade

2005 ◽  
Vol 128 (1) ◽  
pp. 82-95 ◽  
Author(s):  
Francesco Soranna ◽  
Yi-Chih Chow ◽  
Oguz Uzol ◽  
Joseph Katz
Keyword(s):  
Flow Structure ◽  
Rotor Blade ◽  
Normal Component ◽  
Leading Edge ◽  
Suction Side ◽  
Guide Vanes ◽  
Inlet Guide Vanes ◽  
Entire Flow ◽  
Turbulence Production ◽  
Wall Blockage

The flow structure and turbulence around the leading and trailing edges of a rotor blade operating downstream of a row of inlet guide vanes (IGV) are investigated experimentally. Particle image velocimetry (PIV) measurements are performed in a refractive index matched facility that provides unobstructed view of the entire flow field. Data obtained at several rotor blade phases focus on modification to the flow structure and turbulence in the IGV wake as it propagates along the blade. The phase-averaged velocity distributions demonstrate that wake impingement significantly modifies the wall-parallel velocity component and its gradients along the blade. Due to spatially non-uniform velocity distribution, especially on the suction side, the wake deforms while propagating along the blade, expanding near the leading edge and shrinking near the trailing edge. While being exposed to the nonuniform strain field within the rotor passage, the turbulence within the IGV wake becomes spatially nonuniform and highly anisotropic. Several mechanisms, which are consistent with rapid distortion theory (RDT) and distribution of turbulence production rate, contribute to the observed trends. For example, streamwise (in rotor frame reference) diffusion in the aft part of the rotor passage enhances the streamwise fluctuations. Compression also enhances the turbulence production very near the leading edge. However, along the suction side, rapid changes to the direction of compression and extension cause negative production. The so-called wall blockage effect reduces the wall-normal component.

The Effect of IGV Wake Impingement on the Flow Structure and Turbulence Around a Rotor Blade

2005 ◽  
Author(s):  
Francesco Soranna ◽  
Yi-Chih Chow ◽  
Oguz Uzol ◽  
Joseph Katz
Keyword(s):  
Flow Structure ◽  
Rotor Blade ◽  
Normal Component ◽  
Leading Edge ◽  
Suction Side ◽  
Image Velocimetry ◽  
Trailing Edges ◽  
Entire Flow ◽  
Turbulence Production ◽  
Wall Blockage

The flow structure and turbulence around the leading and trailing edges of a rotor blade operating downstream of a row of Inlet Guide Vanes (IGV) are investigated experimentally. Particle Image Velocimetry (PIV) measurements are performed in a refractive index matched facility that provides unobstructed view of the entire flow field. Data obtained at several rotor blade phases focus on modification to the flow structure and turbulence in the IGV wake as it propagates along the blade. The phase-averaged velocity distributions demonstrate that wake impingement significantly modifies the wall-parallel velocity component and its gradients along the blade. Due to spatially non-uniform velocity distribution, especially on the suction side, the wake deforms while propagating along the blade, expanding near the leading edge and shrinking near the trailing edge. While being exposed to the non-uniform strain field within the rotor passage, the turbulence within the IGV wake becomes spatially non-uniform and highly anisotropic. Several mechanisms, which are consistent with rapid distortion theory (RDT) and distribution of turbulence production rate, contribute to the observed trends. For example, streamwise (in rotor frame reference) diffusion in the aft part of the rotor passage enhances the streamwise fluctuations. Compression also enhances the turbulence production very near the leading edge. However, along the suction side, rapid changes to the direction of compression and extension cause negative production. The so-called wall blockage effect reduces the wall-normal component.

Experimental Characterization of the Evolution of Global Flow Structure in the Passage of an Axial Compressor

2021 ◽  
Author(s):  
Ayush Saraswat ◽  
Subhra Shankha Koley ◽  
Joseph Katz
Keyword(s):  
Flow Structure ◽  
Rotor Blade ◽  
Incidence Angle ◽  
Substantial Reduction ◽  
Axial Compressor ◽  
Suction Side ◽  
Rotor Wake ◽  
Tip Leakage ◽  
Blade Tip

Abstract Ongoing experiments conducted in a one-and-half stages axial compressor installed in the JHU refractive index-matched facility investigate the evolution of flow structure across blade rows. After previously focusing only on the rotor tip region, the present stereo-PIV (SPIV) measurements are performed in a series of axial planes covering an entire passage across the machine, including upstream of the IGV, IGV-rotor gap, rotor-stator gap, and downstream of the stator. The measurements are performed at flow rates corresponding to pre-stall condition and best efficiency point (BEP). Data are acquired for various rotor-blade orientations relative to the IGV and stator blades. The results show that at BEP, the wakes of IGV and rotor are much more distinct and the wake signatures of one row persists downstream of the next, e.g., the flow downstream of the stator is strongly affected by the rotor orientation. In contrast, under pre-stall conditions, the rotor orientation has minimal effect on the flow structure downstream of the stator. However, the wakes of the stator blades, where the axial momentum is low, are now wider. For both conditions, the flow downstream of the rotor is characterized by two regions of axial momentum deficit in addition to the rotor wake. A deficit on the pressure side of the rotor wake tip is associated with the tip leakage vortex (TLV) of the previous rotor blade, and is much broader at pre-stall condition. A deficit on the suction side of the rotor wake near the hub appears to be associated with the hub vortex generated by the neighboring blade, and is broader at BEP. At pre-stall, while the axial momentum upstream of the rotor decreases over the entire tip region, it is particularly evident near the rotor blade tip, where the instantaneous axial velocity becomes intermittently negative. Downstream of the rotor, there is a substantial reduction in mean axial momentum in the upper half of the passage, concurrently with an increase in the circumferential velocity. Consequently, the incidence angle upstream of the stator increases in certain regions by up to 30 degrees. These observations suggest that while the onset of the stall originates from the rotor tip flow, one must examine its impact on the flow structure in the stator passage as well.

Ice Crystal Icing Investigation on a Honeywell Uncertified Research Engine in an Altitude Simulation Icing Facility

2020 ◽  
Author(s):  
Ashlie B. Flegel
Keyword(s):  
Particle Size ◽  
Guide Vane ◽  
Leading Edge ◽  
Inlet Guide Vane ◽  
Inlet Temperature ◽  
Ice Crystal ◽  
Guide Vanes ◽  
Inlet Guide Vanes ◽  
The Core ◽  
Break Up

Abstract A Honeywell Uncertified Research Engine was exposed to various ice crystal conditions in the NASA Glenn Propulsion Systems Laboratory. Simulations using NASA’s 1D Icing Risk Analysis tool were used to determine potential inlet conditions that could lead to ice crystal accretion along the inlet of the core flowpath and into the high pressure compressor. These conditions were simulated in the facility to develop baseline conditions. Parameters were then varied to move or change accretion characteristics. Data were acquired at altitudes varying from 5 kft to 45 kft, at nominal ice particle Median Volumetric Diameters from 20 μm to 100 μm, and total water contents of 1 g/m3 to 12 g/m3. Engine and flight parameters such as fan speed, Mach number, and inlet temperature were also varied. The engine was instrumented with total temperature and pressure probes. Static pressure taps were installed at the leading edge of the fan stator, front frame hub, the shroud of the inlet guide vane, and first two rotors. Metal temperatures were acquired for the inlet guide vane and vane stators 1–2. In-situ measurements of the particle size distribution were acquired three meters upstream of the engine forward fan flange and one meter downstream of the fan in the bypass in order to study particle break-up behavior. Cameras were installed in the engine to capture ice accretions at the leading edge of the fan stator, splitter lip, and inlet guide vane. Additional measurements acquired but not discussed in this paper include: high speed pressure transducers installed at the trailing edge of the first stage rotor and light extinction probes used to acquire particle concentrations at the fan exit stator plane and at the inlet to the core and bypass. The goal of this study was to understand the key parameters of accretion, acquire particle break-up data aft of the fan, and generate a unique icing dataset for model and tool development. The work described in this paper focuses on the effect of particle break-up. It was found that there was significant particle break-up downstream of the fan in the bypass, especially with larger initial particle sizes. The metal temperatures on the inlet guide vanes and stators show a temperature increase with increasing particle size. Accretion behavior observed was very similar at the fan stator and splitter lip across all test cases. However at the inlet guide vanes, the accretion decreased with increasing particle size.

Flow Structure and Turbulence in the Tip Region of a Turbomachine Rotor Blade

2007 ◽  
Author(s):  
Francesco Soranna ◽  
Yi-Chih Chow ◽  
Oguz Uzol ◽  
Joseph Katz
Keyword(s):  
Kinetic Energy ◽  
Turbulent Kinetic Energy ◽  
Flow Structure ◽  
Turbulent Diffusion ◽  
Rotor Blade ◽  
Trailing Edge ◽  
Left Hand ◽  
Tip Leakage ◽  
The Subject ◽  
Turbulence Production

The flow structure and turbulence in the tip region of a rotor blade operating downstream of a row of Inlet Guide Vanes (IGVs) are investigated experimentally in a refractive index matched facility that provides unobstructed view of the entire flow field. Stereo-PIV measurements are performed in closely spaced radial planes near the blade tip in a region extending from (slightly upstream of) the blade trailing edge to about 40% of the chord downstream of it. The data enable calculations of all the components of the phase-averaged velocity and vorticity vectors, as well as the strain rate, Reynolds stress, and turbulent diffusion tensors. Each rotor blade is confined between two tip-leakage vortices, a right hand vortex (RHV), generated by the subject blade and propagating along its right hand side, and a left hand vortex (LHV), generated by the previous blade in the same row and propagating along the left hand side of the subject blade. In addition, a trailing edge vortex (TEV) lays underneath the LHV and is subject to intense shearing/deformation by the LHV. RHV-induced radial gradients of radial phase-averaged velocity cause negative turbulence production, P, along the RHV-axis, and formation of a region of low P in the gap between the RHV and the blade suction surface. Trends of turbulent kinetic energy k and P within the RHV do not agree due to the effects of advection by the phase-averaged flow. To the left of the blade, shearing of the TEV by the LHV enhances turbulence production in the region between the two vortices and the rotor wake. Trends of turbulent kinetic energy and its production rate are in good agreement and peaks of k and P occur at the same location. As the TEV migrates away from the LHV, shearing effects become weaker and the dominant contributors to production are terms containing vortex-induced radial gradients of axial and radial velocities. Turbulent diffusion is a minor contributor to the evolution of turbulent kinetic energy in the tip region. It is also shown that the tip-leakage flow/vortex deteriorates the rotor blade performance, causing a ∼66% increase in shaft power input (per unit mass flow-rate) in the tip region in comparison with midspan.

Investigation of the Influence of Hub Gap Size to the Performance of a Subsonic Compressor Stator

2012 ◽  
Author(s):  
Ke Shi ◽  
Haixin Chen ◽  
Song Fu ◽  
Ruben van Rennings ◽  
Frank Thiele
Keyword(s):  
Flow Behavior ◽  
Leading Edge ◽  
Surface Flow ◽  
Channel Height ◽  
Leakage Flow ◽  
Gap Size ◽  
Flow Angle ◽  
Back Flow ◽  
Guide Vanes ◽  
Inlet Guide Vanes

This paper presents a RANS study of the hub clearance effects on the performance of a subsonic compressor stator. The inlet boundary conditions are from the calculation of inlet guide vanes. The k-ω SST turbulence model is adapted to resolve the Reynolds stresses. The present numerical results are compared with the experiment carried out at Technical University Berlin. The circumferentially averaged total pressure has a strong decrease in the lower span region from hub to nearly 50% channel height, while the tangential flow angle reduces from approximately 40% channel height to the hub, linearly. The above phenomenon indicates that the leakage flow in the gap between stator blade and the hub does not turn sufficiently. This leads to a smaller incidence angle of the flow to the stator, thus, the lower span of the stator works in smaller attack angle, 0 to 13 degrees lower than the higher span. Surface flow patterns on the hub and both side of the blade surfaces are compared with the oil flow visualization in the experiment. The compressor stator is shown to operate under large separation and strong back flow conditions. The hub leakage flow is studied together with the endwall flow phenomenon for full gap configuration. Two separation lines are observed on the hub. One is lying in front of the blade leading edge plane indicating the separation due to the leading edge leakage flow which spills out of the passage before the flow enters the passage. The other is caused by the interaction between the strong hub leakage flow and the incoming flow. This separation line undergoes an abrupt turning just after the flow leaving the stator passage. The effect of the hub gap size on the leakage flow and the whole flow passage in the stator, including the strength and location of the vortex structure, the location and size of the separation bubble, as well as the back flow behavior, is analyzed. With the help of a novel vortex identification method, the flow field of this subsonic compressor stator and the inlet guide vanes can be visualized illustrating the behavior at the operation point when rotating instability occurs. The parameter η4 can help identifying the stretching and relaxation of the vortex. This approach reveals significant flow details [1]. Combined with DPH (Dynamic Pressure Head) contour and streamlines, the detailed vortices structures and topology in a subsonic compressor can also be further elucidated. The study illustrates different vortices structures in the compressor, as well as their behavior in different gap size configurations.

scholarly journals Development of a Centrifugal Compressor With a Variable Geometry Split-Ring Pipe Diffuser

1998 ◽  
Author(s):  
J. W. Salvage
Keyword(s):  
Leading Edge ◽  
Lessons Learned ◽  
Operating Point ◽  
Maximum Efficiency ◽  
Variable Geometry ◽  
Impeller Blade ◽  
Guide Vanes ◽  
Part Load ◽  
Inlet Guide Vanes ◽  
Surge Line

Higher noise levels resulted when a compressor was scaled to larger capacity. The machine’s sound pressure level was relieved by increasing the distance between the impeller blade tip and diffuser leading edge. However, the part-load surge line deteriorated severely as a consequence. A variable geometry pipe diffuser solved this problem, permitting operation at stringent off-design conditions. The addition of a variable diffuser permits compressor selection very near its most efficient full-load operating point, without regard for limitations normally imposed by part-load requirements. The principal lessons learned during aerodynamic design refinement include (a) how performance and surge depend upon positioning the variable inlet guide vanes and variable diffuser, and (b) how to define simultaneous variation of inlet guide vanes and diffuser for specific operational objectives. Generally, each operating point requires a unique setting of the variable components to achieve maximum efficiency. However, linked movement is shown to yield both a satisfactory surge line and improved performance for most applications when compared to a compressor without the variable geometry pipe diffuser.

Flow Nonuniformities and Turbulent “Hot Spots” Due to Wake-Blade and Wake-Wake Interactions in a Multi-Stage Turbomachine

2002 ◽  
Vol 124 (4) ◽  
pp. 553-563 ◽  
Author(s):  
Yi-Chih Chow ◽  
Oguz Uzol ◽  
Joseph Katz
Keyword(s):  
Flow Structure ◽  
Hot Spots ◽  
Rotor Blade ◽  
Suction Side ◽  
Direct Result ◽  
Shear Stresses ◽  
Pressure Side ◽  
Rotor Wake ◽  
Multi Stage ◽  
Wake Interactions

This experimental study provides striking examples of the complex flow and turbulence structure resulting from blade-wake and wake-wake interactions in a multi-stage turbomachine. Particle image velocimetry (PIV) measurements are performed within the entire 2nd stage of a two-stage turbomachine. The experiments are performed in a facility that allows unobstructed view of the entire flow field, facilitated using transparent rotor and stator and a fluid that has the same optical index of refraction as the blades. This paper contains data on the phase-averaged flow structure including velocity, vorticity and strain-rate, as well as the turbulent kinetic energy and shear stress, at mid span, for several orientation of the rotor relative to the stator. Two different test setups with different blade geometries are used in order to highlight and elucidate complex phenomena involved, as well as to demonstrate that some of the interactions are characteristic to turbomachines and can be found in a variety of geometries. The first part of the paper deals with the interaction of a 2nd-stage rotor with the wakes of both the rotor and the stator of the 1st stage. Even before interacting with the blade, localized regions with concentrated mean vorticity and elevated turbulence levels form at the intersection of the rotor and stator wakes of the 1st stage. These phenomena persist even after being ingested by the rotor blade of the 2nd stage. As the wake segment of the 1st-stage rotor blade arrives to the 2nd stage, the rotor blades become submerged in its elevated turbulence levels, and separate the region with negative vorticity that travels along the pressure side of the blade, from the region with positive vorticity that remains on the suction side. The 1st-stage stator wake is chopped-off by the blades. Due to difference in mean lateral velocity, the stator wake segment on the pressure side is advected faster than the segment on the suction side (in the absolute frame of reference), creating discontinuities in the stator wake trajectory. The nonuniformities in phase-averaged velocity distributions generated by the wakes of the 1st stage persist while passing through the 2nd-stage rotor. The combined effects of the 1st-stage blade rows cause 10–12 deg variations of flow angle along the pressure side of the blade. Thus, in spite of the large gap between the 1st and 2nd rotors (compared to typical rotor-stator spacings in axial compressors), 6.5 rotor axial chords, the wake-blade interactions are substantial. The second part focuses on the flow structure at the intersection of the wakes generated by a rotor and a stator located upstream of it. In both test setups the rotor wake is sheared by the nonuniformities in the axial velocity distributions, which are a direct result of the “discontinuities” in the trajectories of the stator wake. This shearing creates a kink in the trajectory of the rotor wake, a quadruple structure in the distribution of strain, regions with concentrated vorticity, high turbulence levels and high shear stresses, the latter with a complex structure that resembles the mean strain. Although the “hot spots” diffuse as they are advected downstream, they still have elevated turbulence levels compared to the local levels around them. In fact, every region of wake intersection has an elevated turbulence level.

Heat Transfer and Film-Cooling Measurements on a Stator Vane With Fan-Shaped Cooling Holes

2005 ◽  
Author(s):  
W. Colban ◽  
A. Gratton ◽  
K. A. Thole ◽  
M. Haendler
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Pressure Side ◽  
Guide Vanes ◽  
Single Row ◽  
Inlet Guide Vanes ◽  
Adiabatic Effectiveness ◽  
Lift Off

In a typical gas turbine engine, the gas exiting the combustor is significantly hotter than the melting temperature of the turbine components. The highest temperatures in an engine are typically seen by the turbine inlet guide vanes. One method used to cool the inlet guide vanes is film-cooling, which involves bleeding comparatively low-temperature, high-pressure air from the compressor and injecting it through an array of discrete holes on the vane surface. To predict the vane surface temperatures in the engine, it is necessary to measure the heat transfer coefficient and adiabatic film-cooling effectiveness on the vane surface. This study presents heat transfer coefficients and adiabatic effectiveness levels measured in a scaled-up, two-passage cascade with a contoured endwall. Heat transfer measurements indicated that the behavior of the boundary layer transition along the suction side of the vane showed sensitivity to the location of film-cooling injection, which was simulated through the use of a trip wire placed on the vane surface. Single row adiabatic effectiveness measurements without any upstream blowing showed jet lift-off was prevalent along the suction side of the airfoil. Single row adiabatic effectiveness measurements on the pressure side, also without upstream showerhead blowing, indicated jet lifted-off and then reattached to the surface in the concave region of the vane. In the presence of upstream showerhead blowing, the jet lift-off for the first pressure side row was reduced, increasing adiabatic effectiveness levels.

Heat Transfer and Film-Cooling Measurements on a Stator Vane With Fan-Shaped Cooling Holes

2005 ◽  
Vol 128 (1) ◽  
pp. 53-61 ◽  
Author(s):  
W. Colban ◽  
A. Gratton ◽  
K. A. Thole ◽  
M. Haendler
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Pressure Side ◽  
Guide Vanes ◽  
Single Row ◽  
Inlet Guide Vanes ◽  
Adiabatic Effectiveness ◽  
Lift Off

In a typical gas turbine engine, the gas exiting the combustor is significantly hotter than the melting temperature of the turbine components. The highest temperatures in an engine are typically seen by the turbine inlet guide vanes. One method used to cool the inlet guide vanes is film cooling, which involves bleeding comparatively low-temperature, high-pressure air from the compressor and injecting it through an array of discrete holes on the vane surface. To predict the vane surface temperatures in the engine, it is necessary to measure the heat transfer coefficient and adiabatic film-cooling effectiveness on the vane surface. This study presents heat transfer coefficients and adiabatic effectiveness levels measured in a scaled-up, two-passage cascade with a contoured endwall. Heat transfer measurements indicated that the behavior of the boundary layer transition along the suction side of the vane showed sensitivity to the location of film-cooling injection, which was simulated through the use of a trip wire placed on the vane surface. Single-row adiabatic effectiveness measurements without any upstream blowing showed jet lift-off was prevalent along the suction side of the airfoil. Single-row adiabatic effectiveness measurements on the pressure side, also without upstream showerhead blowing, indicated jet lifted-off and then reattached to the surface in the concave region of the vane. In the presence of upstream showerhead blowing, the jet lift-off for the first pressure side row was reduced, increasing adiabatic effectiveness levels.

Experimental Study of Tip-Gap Turbulent Flow Structure

2006 ◽  
Author(s):  
Genglin Tang ◽  
Roger L. Simpson ◽  
Qing Tian
Keyword(s):  
Turbulent Flow ◽  
Turbulent Flows ◽  
Flow Structure ◽  
Three Dimensional ◽  
Leading Edge ◽  
Suction Side ◽  
Normal Velocity ◽  
Compressor Cascade ◽  
Custom Made ◽  
Turbulent Flow Structure

Experimental results are presented from a study of the tip-gap turbulent flow structure in a low-speed linear compressor cascade wind tunnel at Virginia Tech by utilizing surface oil flow visualization, endwall pressure measurements, and instantaneous velocity measurements with a custom-made 3-orthogonal-velocity-component fiber-optic laser-Doppler velocimetry (LDV) system. Tip gap flows are pressure-driven and highly skewed three-dimensional turbulent flows. The crossflow velocity normal to the blade chord is nearly uniform in the mid tip gap and changes substantially from the pressure to suction side due to the local tip pressure loading while the TKE does not vary much across the mid tip gap. The tip gap flow correlations of streamwise and wall normal velocity fluctuations decrease significantly from the leading edge to the trailing edge of the blade due to flow skewing.

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