Seung Jin Song

Clearance Effects on the Onset of Instability in a Centrifugal Compressor

This report intends to shed an insight into the effect of large relative tip clearances on the onset of instability in a highly loaded centrifugal compressor. Time-resolved pressure measurements have been performed along the casing of a scaled-up model of a small compressor for two clearances at a wide range of operating conditions. Based on these time-resolved measurements, the pressure distribution along the meridional length and the blade loading distribution are calculated for each operating condition. In addition, the phase locked pressure fluctuation and its deviation are computed. The results show the behavior of each subcomponent of the compressor at different flow conditions and explain the role of the relative tip clearance on the onset of instability. For high mass-flow rates, the steady pressure distribution along the casing reveals that the inducer acts as an accelerating nozzle. Pressure is only built up in the radial part due to the centrifugal forces and in the subsequent diffuser due to area change. For off-design conditions, incidence effects are seen in the blade loading distribution at the leading edge while the inducer is unloaded. A region of high pressure deviation originates at the leading edge of the main blade and convects downstream. This feature is interpreted as the trajectory of the leakage vortex. The trajectory of these vortices is strongly affected by the mass-flow coefficient. If the mass-flow rate is sufficiently small, the trajectory of the leakage vortex becomes perpendicular to the axis of rotation, the leakage vortex interacts with the adjacent blade, and inlet tip recirculation is triggered. If the flow rate is further reduced, the leakage vortex vanishes and rotating stall is initiated in the diffuser. For larger clearances, stronger vortices are formed, stall is triggered at higher flow rates, and the overall compressor performance deteriorates.

Rotordynamic Forces Due to Turbine Tip Leakage: Part I—Blade Scale Effects

An experimental and theoretical investigation has been conducted on rotordynamic forces due to nonaxisymmetric turbine tip leakage effects. This paper presents an actuator disk model that describes the flow response to a finite clearance at the rotor tip. The model simplifies the flow field by assuming that the radially uniform flow splits into two streams as it goes through the rotor. The stream associated with the tip clearance, or the underturned stream, induces radially uniform unloading of the rest of the flow, called the bladed stream. Thus, a shear layer forms between the two streams. The fraction of each stream and the strength of shear layer between the two are found as functions of the turbine loading and flow parameters without resorting to empirical correlations. The results show that this model`s efficiency predictions compare favorably with the experimental data and predictions from various correlations. A companion paper builds on this analysis to yield a model of the three-dimensional disturbances around an offset turbine and to predict the subsequent cross forces.

Rotordynamic Forces Due to Turbine Tip Leakage: Part II—Radius Scale Effects and Experimental Verification

This paper presents a radius scale actuator disk model, which describes the flow response to a whirling/spinning rotor in an unshrouded turbine. At each azimuth, the upstream-downstream flow variables are matched by the results from a steady blade scale analysis presented in a companion paper, with allowance for mass storage in the stator-rotor region. The new model can accurately predict the magnitude of both direct and cross excitation forces as well as their breakdown into work extraction and pressure effects. The trends versus the mean flow coefficient and interblade distance are predicted. While underpredicted, a trend versus mean rotor tip clearance height is also indicated. Thus, the new model captures the dominant physical effects caused by a whirling/spinning rotor in an unshrouded turbine.

Performance Degradation Due to Blade Surface Roughness in a Single-Stage Axial Turbine

Turbine blades experience significant surface degradation with service. Previous studies indicate that an order of magnitude or greater increase in roughness height is typical, and these elevated levels of surface roughness significantly influence turbine efficiency and heat transfer. This paper presents measurement and a mean line analysis of turbine efficiency reduction due to blade surface roughness. Performance tests have been conducted in a low speed, single-stage, axial flow turbine with roughened blades. Sheets of sandpaper with equivalent sandgrain roughnesses of 106 and 400 μm have been used to roughen the blades. The roughness heights correspond to foreign deposits on real turbine blades measured by Bons et al. [1]. In the transitionally rough regime (106 μm), normalized efficiency decreases by approximately 4 percent with either roughened stator or roughened rotor and 8 percent with roughness on both the stator and rotor blades. In the fully rough regime (400 μm), normalized efficiency decreases by 2 percent with roughness on the pressure side and by 6 percent with roughness on the suction side. Also, the normalized efficiency decreases by 11 percent with roughness only on stator vanes; 8 percent with roughness only on rotor blades; and 19 percent with roughness on both the stator and rotor blades.Copyright

Effects of Reynolds Number and Surface Roughness Magnitude and Location on Compressor Cascade Performance

An experimental investigation has been conducted to characterize the influence of Reynolds number and surface roughness magnitude and location on compressor cascade performance. Flow field surveys have been conducted in a low-speed, linear compressor cascade. Pressure, velocity, and loss have been measured via a five-hole probe, pitot probe, and pressure taps on the blades. Four different roughness magnitudes, Ra values of 0.38 lm (polished), 1.70 lm (baseline), 2.03 lm (rough 1), and 2.89 lm (rough 2), have been tested. Furthermore, various roughness locations have been examined. In addition to the as manufactured (baseline) and entirely rough blade cases, blades with roughness covering the leading edge, pressure side, and 5%, 20%, 35%, 50%, and 100% of suction side from the leading edge have been studied. All of the tests have been carried out for Reynolds numbers ranging from 300,000 to 640,000. For Reynolds numbers under 500,000, the tested roughnesses do not significantly degrade compressor blade loading or loss. However, loss and blade loading become sensitive to roughness at Reynolds numbers above 550,000. Cascade performance is more sensitive to roughness on the suction side than pressure side. Furthermore, roughness on the aft 2=3 of suction side surface has a greater influence on loss. For a given roughness location, there exists a Reynolds number at which loss begins to significantly increase. Finally, increasing the roughness area on the suction surface from the leading edge reduces the Reynolds number at which the loss begins to increase. [DOI: 10.1115/1.4003821]

Performance degradation due to blade surface roughness in a single-stage axial turbine

Turbine blades experience significant surface degradation with service. Previous studies indicate that an order-of-magnitude or greater increase in roughness height is typical, and these elevated levels of surface roughness significantly influence turbine efficiency and heat transfer. This paper presents measurement and a mean-line analysis of turbine efficiency reduction due to blade surface roughness. Performance tests have been conducted in a low-speed, single-stage, axial flow turbine with roughened blades. Sheets of sandpaper with equivalent sandgrain roughnesses of 106 and 400 μm have been used to roughen the blades. The roughness heights correspond to foreign deposits on real turbine blades measured by Bons et al. [1]. In the transitionally rough regime (106 μm), normalized efficiency decreases by approximately 4% with either roughened stator or roughened rotor and by 8% with roughness on both the stator and rotor blades. In the fully rough regime (400 μm), normalized efficiency decreases by 2% with roughness on the pressure side and by 6% with roughness on the suction side. Also, the normalized efficiency decreases by 11% with roughness only on stator vanes, 8% with roughness only on rotor blades, and 19% with roughness on both the stator and rotor blades.

Lateral forces from single gland rotor labyrinth seals in turbines

Even though interest in labyrinth seal flows has increased recently, an analytical model capable of predicting turbine flow response to labyrinth seals is still lacking. Therefore, this paper presents a new model to predict flow response in an axial turbine stage with a shrouded rotor. A concentric model is first developed, and this model is used to develop an eccentric model. Basic conservation laws are used in each model, and a nonaxisymmetric sealing gap is prescribed for the eccentric model. Thus, the two models can predict the evolution of a uniform upstream flow into a nonuniform downstream flow. In turbines with concentric shrouded rotors, the seal flow is retarded in the axial direction and tangentially underturned. In turbines with eccentric shrouded rotors, flow azimuthally migrates away from and pressure reaches its peak near the maximum sealing gap region. Finally, the rotordynamic implications of such flow nonuniformities are discussed and compared against eccentric unshrouded turbine predictions.

Design, Fabrication, and Testing of a MEMS Microturbine

This paper describes the design, fabrication, and testing of a microturbine developed at Seoul National University Here, the term “microturbine” refers to a radial turbine with a diameter on the order of a centimeter Such devices can be used to transmit power for various systems The turbine is designed using a commercial CFD code, and it has a design flow coefficient of 0 238 and work coefficient of 0 542 It has 31 stator blades and 24 totor blades A hydrodynamic journal bearing and hydrostatic thrust bearings counteract radial and axial forces on the rotor The test turbine consists of a stack of five wafers and is fabricated by MEMS technology, using photolithography, DRIE, and bonding processes The first, second, fourth, and fifth layers contain plumbing, and hydrostatic axial thrust bearings for the turbine The third wafer contains the turbine’s stator, rotor, and hydrodynamic journal bearings Furthermore, a turbine test facility containing a flow control system and instrumentation has been designed and constructed In performance tests, a maximum rotation speed of 11,400 rpm and flow rate of 16,000 seem have been achieved

Pressure Gradient Effects on Smooth and Rough Surface Turbulent Boundary Layers—Part I: Favorable Pressure Gradient

The effects of the pressure gradient and surface roughness on turbulent boundary layers have been experimentally investigated. In Part I, smooth- and rough-surface turbulent boundary layers with and without favorable pressure gradients (FPG) are presented. All of the tests have been conducted at the same Reynolds number (based on the length of the flat plate) of 900,000. Streamwise time-mean and fluctuating velocities have been measured using a single-sensor hot-wire probe. For smooth surfaces, the FPG decreases the mean velocity defect and increases the wall shear stress; however, the friction coefficient hardly changes due to the increased freestream velocity. The FPG effect on the streamwise normal Reynolds stress has been examined. The FPG increases the streamwise normal Reynolds stress for y less than 0.6 times the boundary layer thickness. With the zero pressure gradient (ZPG), the roughness increases the mean velocity defect throughout the boundary layer and increases the normal Reynolds stress for y greater than twice the average roughness height. The roughness effect on the mean velocity defect is stronger under the FPG than under the ZPG for y less than 30 times the average roughness height. For y less than 25 times the average roughness height, the roughness effect of increasing normal Reynolds stress is also stronger under the FPG than under the ZPG. Consequently, for a rough surface, the FPG increases the integrated streamwise turbulent kinetic energy, friction coefficient, roughness Reynolds number, and roughness shift. Furthermore, the FPG increases the roughness effects on the integral boundary layer parameters—the boundary layer thickness, momentum thickness, ratio of the displacement thickness to the boundary layer thickness, and shape factor. Thus, the FPG augments the roughness effects on turbulent boundary layers.

Influence of Surface Roughness on the Performance of a Compressor Blade in a Linear Cascade: Experiment and Modeling

Compressor blades experience significant surface degradation with service. Elevated levels of surface roughness reduce compressor efficiency and mass flow rate. This paper presents measurement and a new model of compressor blade performance degradation due to blade surface roughness. Performance tests have been conducted in a low-speed, linear cascade with roughened compressor blades. Equivalent sandgrain roughnesses of 12, 180, 300, 425, and 850 microns have been tested. These roughness values are representative of compressor blade roughnesses found in actual gas turbines in service. Flow angle, flow rate, and loss have been measured. For the tested roughnesses of 180, 300, 425, and 850 microns, the axial velocity ratio decreases by 0.1, 2.1, 2.5, and 5.4%, respectively. For the same cases, the deviation increases by 24, 38, 51, and 70%, respectively. Finally, the loss increases by 12, 44, 132, and 217%, respectively. Thus, among the three parameters, the loss responds most sensitively to changes in compressor blade roughness. Furthermore, a new mean-line model based on the assumption of 50% reaction stages has been developed to estimate the effects of roughness on the performance of a multi-stage compressor. The data from the cascade data are used as input to predict the performance of a single compressor stage. Subsequently, a stage-stacking method is used to enable prediction for a multi-stage compressor. According to the model, the pressure ratio, and mass flow rate are significantly influenced by the blade surface roughness.© 2009 ASME