scholarly journals Influence of Upstream and Downstream Compressor Stators on Rotor Exit Flow Field

2014 ◽  
Vol 2014 ◽  
pp. 1-10 ◽  
Author(s):  
Nicole L. Key
Keyword(s):  
Potential Field ◽  
Computational Models ◽  
Leading Edge ◽  
Suction Side ◽  
Rotor Wake ◽  
Flow Angle ◽  
Absolute Flow ◽  
Measurement Plane ◽  
Vane Clocking

Measurements acquired at the rotor exit plane illuminate the interaction of the rotor with the upstream vane row and the downstream vane row. The relative phase of the upstream and downstream vane rows is adjusted using vane clocking so that the effect of the upstream propagating potential field from the downstream stator can be distinguished from the effects associated with the wakes shed from the upstream stator. Unsteady absolute flow angle information shows that the downstream potential field causes the absolute flow angle to increase in the vicinity of the downstream stator leading edge. The presence of Stator 1 wake is also detected at this measurement plane using unsteady total pressure data. The rotor wakes are measured at different circumferential locations across the vane passage, and the influence of Stator 1 wake on the suction side of the rotor wake is evident. Also, the influence of the downstream stator is detected on the pressure side of the rotor wake for a particular clocking configuration. Understanding the role of the surrounding vane rows on rotor wake development will lead to improved comparison between experimental data and results from computational models.

An Investigation of the Flow Physics of Vane Clocking Using Unsteady Flow Measurements

2008 ◽  
Author(s):  
Nicole L. Key ◽  
Patrick B. Lawless ◽  
Sanford Fleeter
Keyword(s):  
Boundary Layer ◽  
Unsteady Flow ◽  
Leading Edge ◽  
Suction Side ◽  
Boundary Layer Separation ◽  
Flow Measurements ◽  
Rotor Wake ◽  
Efficiency Condition ◽  
Vane Clocking ◽  
Stage Efficiency

Vane clocking, the circumferential indexing of adjacent vane rows with similar vane counts, has been shown to affect stage efficiency in compressors and turbines. Steady flow measurements acquired in the embedded stage of the Purdue 3-Stage Compressor showed a change in stage efficiency with vane clocking, as discussed in a companion paper. The optimum efficiency condition at design loading occurred when the upstream vane wake impinged on the downstream vane, as had been reported by other vane clocking studies. However, at high loading, the impingement of the upstream vane wake triggered a vane suction side boundary layer separation and resulted in the worst efficiency condition. The objective of this research is to experimentally investigate the maximum and minimum efficiency clocking configurations with unsteady flow measurements to illuminate the flow physics associated with the measured changes in Stage 2 performance. Vane exit unsteady total pressure, velocity, and flow angle measurements were acquired at 50 pitchwise locations spanning one vane passage. Fourier decomposition is used to identify the impact of the upstream rotor wake on the shedding characteristics of the Stator 2 boundary layer and how the placement of the upstream vane wake affects this phenomenon. For the clocking configuration that located the Stator 1 wake at the leading edge of the Stator 2 vane at design loading, it dampened the boundary layer response to the fluctuating incidence associated with rotor wake chopping, leading to a reduction in the size of the structures shed in the Stator 2 vane wake. At the high loading condition, the placement of the Stator 1 wake at the leading edge of Stator 2 triggered a suction side boundary layer separation, resulting in an absence of the upstream rotor blade pass frequency in the spectrum measured in the Stator 2 wake.

An Experimental Study of Stall Suppression and Associated Changes to the Flow Structures in the Tip Region of an Axial Low Speed Fan Rotor by Axial Casing Grooves

2017 ◽  
Author(s):  
Huang Chen ◽  
Yuanchao Li ◽  
Subhra Shankha Koley ◽  
Nick Doeller ◽  
Joseph Katz
Keyword(s):  
Flow Rate ◽  
Large Scale ◽  
Pressure Rise ◽  
Leading Edge ◽  
Suction Side ◽  
Rotating Stall ◽  
Flow Structures ◽  
Flow Angle ◽  
Low Speed ◽  
Flow Characterization

The effects of axial casing grooves on the performance and flow structures in the tip region of an axial low speed fan rotor have been studied experimentally in the JHU refractive index-matched liquid facility. The four-per-passage semicircular grooves are skewed by 45° in the positive circumferential direction, and have a diameter of 65% of the rotor blade axial chord length. A third of the groove overlaps with the blade front, and the rest extends upstream. These grooves have a dramatic effect on the machine performance, reducing the stall flow rate by 40% compared to the same machine with a smooth endwall. However, they reduce the pressure rise at high flow rates. The flow characterization consists of qualitative visualizations of vortical structures using cavitation, as well as stereo-PIV (SPIV) measurements in several meridional and (z,θ) planes covering the tip region and interior of the casing grooves. The experiments are performed at a flow rate corresponding to pre-stall conditions for the untreated machine. They show that the flow into the downstream sides of the grooves and the outflow from their upstream sides vary periodically. The inflow peaks when the downstream end is aligned with the pressure side (PS) of the blade, and decreases, but does not vanish, when this end is located near the suction side (SS). These periodic variations have three primary effects: First, substantial fractions of the leakage flow and the tip leakage vortex (TLV) are entrained periodically into the groove. Consequently, in contrast to the untreated flow, The TLV remnants remain confined to the vicinity of the entrance to the groove, and the TLV strength diminishes starting from the mid-chord. Second, the grooves prevent the formation of large scale backflow vortices (BFVs), which are associated with the TLV, propagate from one blade passage to the next, and play a key role in the onset of rotating stall in the untreated fan. Third, the flow exiting from the grooves causes periodic variations of about 10° in the relative flow angle around the blade leading edge, presumably affecting the blade loading. The distributions of turbulent kinetic energy provide statistical evidence that in contrast to the untreated casing, very little turbulence originating from a previous TLV, including the BFVs, propagates from the PS to the SS of the blade. Hence, the TLV-related turbulence remain confined to the entrance to groove. Elevated, but lower turbulence is also generated as the outflow from the groove jets into the passage.

Numerical Investigation of Diffuser Flow Field and Rotating Stall in a Centrifugal Compressor With Vaned Diffuser

2017 ◽  
Author(s):  
Yang Zhao ◽  
Jiayi Zhao ◽  
Zhiheng Wang ◽  
Guang Xi
Keyword(s):  
Centrifugal Compressor ◽  
Leading Edge ◽  
Suction Side ◽  
Rotating Stall ◽  
Tip Leakage Flow ◽  
Vaned Diffuser ◽  
Flow Angle ◽  
Diffuser Flow ◽  
Flow Phenomena ◽  
Reversal Flow

The diffuser rotating stall in a centrifugal compressor with vaned diffuser is one of important unsteady flow phenomena, which limits the operating range of the compressor. In this paper, the unsteady CFD analysis on a low-speed centrifugal compressor has been performed to investigate the flow characteristic in the diffuser and the propagation of the diffuser rotating stall. The flow behaviors at the outlet of the impeller at design and off-design conditions are firstly investigated. It is found that a reversal flow, induced by the tip leakage flow, exists near the shroud at the impeller outlet and becomes serious with the mass flow rate reduced. Due to the span-wise variation of the flow angle at the diffuser inlet and the inversed pressure gradient in the passage, the leading-edge vortex (LEV) generates on the diffuser leading edge. The LEV then induces the secondary flow in the diffuser passage and then causes the hub-corner separation. Furthermore, the propagation of the diffuser rotating stall is presented in details. The suction-side separation near the hub induces the blockage in the passage. And the shedding vortex from the suction side moves toward the leading edge of the adjacent blade. When the vortex reaches to the leading edge of the adjacent blade, the incidence increase and a new separation occurs on the suction side. With the development of the new separation, the passage becomes blocked gradually and the upstream stalled passage recovers to a normal condition. The rotating stall propagates along the direction of the impeller rotation at about 4.5% of the impeller rotational speed.

scholarly journals Optimization of the LS89 Axial Turbine Profile Using a CAD and Adjoint Based Approach †

2018 ◽  
Vol 3 (3) ◽  
pp. 20 ◽  
Author(s):  
Ismael Sanchez Torreguitart ◽  
Tom Verstraete ◽  
Lasse Mueller
Keyword(s):  
Entropy Generation ◽  
Computer Aided Design ◽  
Stokes Equations ◽  
Leading Edge ◽  
Suction Side ◽  
Optimized Design ◽  
Geometric Continuity ◽  
Cfd Validation ◽  
Flow Angle ◽  
Axial Turbine

The LS89 high pressure axial turbine vane was originally designed and optimized for a downstream isentropic Mach number of 0.9. This profile has been widely used for computational fluid dynamics (CFD) validation in the open literature but very few attempts have been made to improve the already optimized design. This paper presents a sound methodology to design and optimize the LS89 using computer-aided design (CAD) at design conditions. The novelty of the study resides in the parametrization of design space, which is done at the CAD level, and the detailed analysis of the aerodynamic performance of the optimized design. Higher level constraints are imposed on the shape, such as the trailing edge thickness, the axial chord length, and G2 geometric continuity between the suction side and pressure side at the leading edge. The gradients used for the optimization are obtained by applying algorithmic differentiation to the CAD kernel and grid generator and the discrete adjoint method to the CFD solver. A reduction of almost 12% entropy generation is achieved, which is equivalent to a 16% total pressure loss reduction. The entropy generation is reduced while keeping the exit flow angle as a flow constraint, which is enforced via the penalty formulation. The resulting unconstrained optimization problem is solved by the L-BFGS-B algorithm. The flow is governed by the Reynolds-averaged Navier-Stokes equations and the one-equation transport Spalart-Allmaras turbulence model. The optimal profile is compared and benchmarked against the baseline case.

Flow Non-Uniformities and Turbulent “Hot Spots” Due to Wake-Blade and Wake-Wake Interactions in a Multistage Turbomachine

2002 ◽  
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 ◽  
Flow Angle ◽  
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 positive vorticity that travels along the pressure side of the blade, from the region with negative vorticity that remains on the suction side. The 1st stage stator wake is chopped-off by the blades. Due to difference in mean tangential 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 non-uniformities 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° 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 non-uniformities in the horizontal 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.

Secondary Flow and Loss Distribution in a Radial Compressor With Untwisted Backswept Vanes

1990 ◽  
Author(s):  
G. Sipos
Keyword(s):  
Three Dimensional ◽  
Leading Edge ◽  
Suction Side ◽  
Meridional Flow ◽  
Specific Work ◽  
Suction Surface ◽  
Inlet Flow ◽  
Loss Distribution ◽  
Flow Angle ◽  
Radial Compressor

The unshrouded impeller and the vaneless diffuser of a single-stage radial compressor have been investigated at three flow rates. Three-dimensional velocities and pressures were measured at a tip speed of 84 m/s by an L2F-velocimeter, a slanted single hot-wire probe and piezoresistive pressure transducers. The measurements show that upstream the blading the averaged meridional inlet flow angle is about 54 degree and a periodical variation of the meridional flow angle of about 25 degree occurs near the casing wall. Further, an inlet vortex of clockwise direction appears and an initial whirl is induced. The specific work of the initial whirl corresponds to approximately 12% of the enthalpy losses between inlet pipe and diffuser outlet. In the beginning of the passage, the inlet vortex is suppressed and a solid body vortex of counterclockwise direction can be observed. At the outlet, a heavy flow deceleration at the blade suction side with subsequent separation can be seen. Increasing the flow rate decreases the wake and causes a more uniform loss distribution in this area. The measured secondary vortex flow and rotary stagnation pressure gradients are compared with test results from impellers with inducer. The incidence of the investigated impeller is greater than that of the impellers with inducer, but the wake-jet outlet flows are very similar. Inlet losses could be reduced by improving incidence angles by matching the blade angles to the inlet flow angles. Smaller blade angles at the shroud would reduce or eliminate separation at the leading edge, and the resulting reduction in low momentum fluid along the suction surface would help to avoid separation on that surface near the outlet.

Influencing the Secondary Losses in Compressor Cascades by a Leading Edge Bulb Modification at the Endwall

2002 ◽  
Author(s):  
R. Mu¨ller ◽  
H. Sauer ◽  
K. Vogeler ◽  
M. Hoeger
Keyword(s):  
Leading Edge ◽  
Suction Side ◽  
Side Branch ◽  
Pressure Measurements ◽  
Numerical Work ◽  
Low Speed ◽  
Passage Vortex ◽  
Turbine Cascades ◽  
Horse Shoe Vortex

Recent investigations have shown a significant reduction of secondary losses in turbine cascades using a modification of the blade at the endwall, a so called bulb. This paper deals with the same objective but is focussed on experimental and numerical work in compressor cascades. The cascades are modified near the endwall with a similar bulb as the earlier turbine cascades. The investigations have been carried out on a modified profile hub section of the Dresden Low Speed Research Compressor (LSRC) rotor blade, a compressor profile with a nominal turning of 18 degree. A datum configuration and a bulb configuration were tested in the Dresden Low Speed Cascade Wind Tunnel. An intensified suction side branch of the horse shoe vortex by a bulb was expected counterrotating to the passage vortex with an influence on its propagation. The interaction of the passage vortex and the boundary layer on the blade suction side is influenced. The superposition of both is decreased and the losses developing by this effect are significantly lower. The cases show a reduction in losses of 0.5–1.5% as a function of the blade turning. This equals a reduction of the isolated secondary losses by 15–25% with respect to the reference profile. It supports the physical understanding of the role of the horse shoe vortex in the loss production due to the passage vortex in compressor cascades. Detailed results of total pressure measurements are presented for both cascades.

Bio-Inspired High Pressure Turbine Optimization Strategy for Hybrid-Electric Engines Operating at Off-Design

2020 ◽  
Author(s):  
Paht Juangphanich ◽  
Guillermo Paniagua ◽  
Vikram Shyam
Keyword(s):  
High Pressure ◽  
Separation Bubble ◽  
Leading Edge ◽  
Suction Side ◽  
Superior Performance ◽  
Optimization Strategy ◽  
High Pressure Turbine ◽  
Flow Angle ◽  
Diverse Range ◽  
Pressure Losses

Abstract Incident tolerant turbine design is a major challenge for any turbomachinery designer. High Pressure Turbines experience large aerodynamic losses when operating at reduced massflow and lower RPM. Turbine performance is adversely impacted at positive incidence angles due to shifting of the stagnation point towards the pressure side. This can cause a separation bubble in the aft suction side region. In marine life, a diverse range of animals have developed wavy surfaces along their fins and bodies to prevent stall or flow separation at engine-relevant Reynolds numbers, but for incompressible fluids. This paper describes a novel parameterization strategy for optimizing wavy-shaped airfoils to offer superior performance at off-design operation, in the present case, at positive incidence. The methodology can be applied to all types of aircraft engines: one, two, or three spool engine configurations. The parameterized geometries are compatible with existing gas turbine manufacturing processes including casting and additive manufacturing [1,2]. The objective of the optimization was to discover the appropriate waveform combinations at the airfoil leading edge, trailing edge, and suction side characterized by their amplitude, phase, and frequency, such that the airfoils offer the lowest possible pressure losses at 15 degrees positive incidence. The optimization was performed on a high pressure turbine passage, optimized for best efficiency at nominal conditions, while maintaining the same exit flow angle and massflow. The Reynolds number is 850,000. Based on 286 designs produced, the results of the optimization show a clear benefit at positive incidence, at the expense of a slightly lower performance at nominal conditions. A final comparison of the optimized rotor with stage is included in the analysis.

Secondary Flow and Loss Distribution in a Radial Compressor With Untwisted Backswept Vanes

1991 ◽  
Vol 113 (4) ◽  
pp. 686-695 ◽  
Author(s):  
G. Sipos
Keyword(s):  
Three Dimensional ◽  
Leading Edge ◽  
Suction Side ◽  
Periodic Variation ◽  
Specific Work ◽  
Suction Surface ◽  
Inlet Flow ◽  
Loss Distribution ◽  
Flow Angle ◽  
Radial Compressor

The unshrouded impeller and the vaneless diffuser of a single-stage radial compressor have been investigated at three flow rates. Three-dimensional velocities and pressures were measured at a tip speed of 84 m/s by an L2F-velocimeter, a slanted single hotwire probe, and piezoresistive pressure transducers. The measurements show that upstream of the blading the averaged meridional inlet flow angle is about 54 deg and a periodic variation of the meridional flow angle of about 25 deg occurs near the casing wall. Further, an inlet vortex in the clockwise direction appears and an initial whirl is induced. The specific work of the initial whirl corresponds to approximately 12 percent of the enthalpy losses between inlet pipe and diffuser outlet. In the beginning of the passage, the inlet vortex is suppressed and a solid body vortex in the counterclockwise direction can be observed. At the outlet, a heavy flow deceleration at the blade suction side with subsequent separation can be seen. Increasing the flow rate decreases the wake and causes a more uniform loss distribution in this area. The measured secondary vortex flow and rotary stagnation pressure gradients are compared with test results from impellers with inducer. The incidence of the investigated impeller is greater than that of the impellers with inducer, but the wake-jet outlet flows are very similar. Inlet losses could be reduced by improving incidence angles by matching the blade angles to the inlet flow angles. Smaller blade angles at the shroud would reduce or eliminate separation at the leading edge, and the resulting reduction in low-momentum fluid along the suction surface would help to avoid separation on that surface near the outlet.

Lung Computational Models and the Role of the Small Airways in Asthma

PEDIATRICS ◽  
2020 ◽  
Vol 146 (Supplement 4) ◽  
pp. S359.2-S360
Author(s):  
Jennilee Eppley ◽  
Todd Mahr
Keyword(s):  
Computational Models ◽  
Small Airways
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