Literatura académica sobre el tema "COMPRESSOR BLADES DETERIORATION"

Deterioration of axial compressors is in general a major concern in aircraft engine maintenance. Among other effects, roughness in high-pressure compressor reduces the pressure rise and thus efficiency, thereby increasing the specific fuel consumption of an engine. Therefore, it is important to improve the understanding of roughness on compressor blading and their impact on compressor performance. To investigate the surface roughness of rotor blades of a compressors, different stages of an axial high-pressure compressor and a first-stage blisk (BLade–Integrated–dISK) of a regional aircraft engine is measured by a three-dimensional laser scanning microscope. Fundamental types of roughness structures can be identified: impacts in different sizes, depositions as isotropically distributed single elements with steep flanks and anisotropic roughness structures direct approximately normal to the flow direction. To characterise and quantify the roughness structures in more detail, roughness parameters were determined from the measured surfaces. The quantification showed that the roughness height varies through the compressor depending on the stage, position and the blade side. Overall complex roughness structures of different shape, height and size are detected regardless of the type of the blades.

Gas turbine axial compressor blades may encounter damage during service for various reasons such as damage by debris from casing or foreign objects impacting the blades, typically near the rotor’s tip. This may lead to deterioration of performance and reduction in the surge margin. The damage breaks the cyclic symmetry of the rotor assembly; thus, computational fluid dynamics simulations have to be performed using full annulus compressor assembly. Moreover, downstream boundary conditions are unknown during rotating stall or surge, and simulations become difficult. This paper presents unsteady computational fluid dynamics analyses of compressor performance with tip curl damage. Computations were performed near the stall boundary. The primary objectives are to understand the effect of the damage on the flow behaviour and compressor stability. Computations for the undamaged rotor assembly were also performed as a reference case. A transonic axial compressor rotor was used for the time-accurate numerical unsteady flow simulations, with a variable area nozzle downstream simulating an experimental throttle. Computations were performed at 60% of the rotor design speed. Two different degrees of damage for one blade and multiple damaged blades were investigated. Rotating stall characteristics differ including the number of stall cells, propagation speed and rotating stall cell characteristics. Contrary to expectations, damaged blades with typical degrees of damage do not show noticeable effects on the global compressor performance near stall.

Erosion of compressor blades due to operation in particulate environments is a serious problem for the manufacturers and users of industrial and aeronautical gas turbines, because of drastic degradations in performance, mostly through blunting of blade leading edges, reduction of chord and increase of tip clearance and surface roughness. This paper presents a numerical study to assess the effects of erosion by sand ingestion on blade geometry deterioration and the subsequent performance degradation. These computations were carried out for an axial turbomachine in steps; first, calculations of particle trajectories and erosion resulting from cumulative impacts by sand particles (MIL-E 5007E, 0–1000 μm) were carried out, then, the required data were used in the estimation of performance degradation based on a mean-line method that included Lieblein and Koch-Smith loss correlations, in addition to an erosion fault model derived from blade geometry deterioration. This global procedure was successfully validated upon an axial fan stage, and can be generalized easily to other axial compressor designs.

It is well known that variations in stagger angle between rotor blades affect compressor performance. In this paper, the stagger angle of blade No. 8 is increased or decreased by six degrees for non-uniformity, and the influence of rotor non-uniformity caused by the change in only one blade stagger angle on the performance and stability of the compressor is investigated. The experimental results show that whether the local rotor stagger angle increases or decreases, the compressor stability will deteriorate. If the stagger angle of blade No. 8 is reduced by six degrees, the flow coefficient at the stall point increases by 8.5%. If the stagger angle of blade No. 8 is increased by six degrees, the flow coefficient at the stall point increases by 1.5%. The reason for the deterioration of compressor stability caused by the local non-uniform rotor stagger angle is explored. When the stagger angle of rotor blade No. 8 deviates from the designed state, the load of blade No. 8 and the surrounding blades will change. The load on rotor blade No. 8 increases when the stagger angle decreases. In the near-stall condition, blade No. 8 becomes the “dangerous blade” that triggers the stall. As the stagger angle of rotor blade No. 8 increases, the load on blade No. 8 decreases. However, the load on blade No. 9 increases due to flow redistribution and blade No. 9 becomes a “dangerous blade” that triggers stall. The “dangerous blade” caused by the non-uniformity of stagger angle is the direct reason for the advance of the compressor rotating stall.

ABSTRACTThis paper investigates the role degradation of protective diffusion aluminide coating on Inconel 713LC used for CT blades of short-haul aircraft fleet played in having the blades prematurely retired from service at 6378 hours, as opposed to their pre-set service time of 10000 hours. The blade samples were subjected to various examinations; X-ray diffraction (XRD), X-ray fluorescence (XRF), scanning electron microscopy (SEM) and energy dispersive spectroscopy (EDS) analyse at the; tips, airfoil, as well as the base, transverse and longitudinal, sectioned and unsectioned. As affirmed by both the transverse and longitudinal sections examinations, it was established that thermal attack leading to deterioration of the coating was greater at the tip and airfoils of the blades (the hotter zones) and lesser towards the bases (colder zones). As a result, severe degradation of the core material at the tips and airfoils compared to the bases and more prevalent at the leading edges than trailing edges at the tips. The results further suggest that both active outward Ni diffusion and inward Al diffusion can coexist during exploitation of the blades in service. The study illustrates the role played by the aluminide coating in early failure of CT blades with the aim of bettering the surface coatings and enhancing coating technologies, managing CT blade material monitoring as well as to give insights on advancing CT blades maintenance practices.

In gas turbines and jet engines, stagger angle and tip gap variations between adjacent blades lead to the deterioration of performance. To evaluate the effect of manufacturing tolerance on performance, a CFD-based uncertainty quantification analysis is performed in this work. However, evaluating dozens of thousands of rotor assembly through CFD simulations would be computationally prohibitive. A surrogate model is thus developed to predict compressor performance given an ordered set of manufactured blades. The model is used to predict the influence of tip gap and stagger angle variations on maximum isentropic efficiency. The results confirm that the best arrangement is obtained by minimizing the stagger angle variation between adjacent blades, and by maximizing the tip gap variation. Another finding is that the best arrangement yields the lowest variability, the range of maximum efficiency being 4 times sharper (resp. 2 times) than worst arrangement for stagger angle variations (resp. tip gap variations). Not measuring manufacturing tolerance, or not specifying any strategy for the blade arrangement, lead to variability as large as the worst arrangement.

In modern jet propulsion systems the core engine has an essential influence on the total engine performance. Especially the high pressure compressor plays an important role in this scheme. Substantial factors here are losses due to tip clearance effects and aerodynamic airfoil quality. During flight operation the airfoils are subject to wear and tear on the leading edge. These effects cause a shortening of the chord length and the leading edge profiles become deformed. This results in a deterioration of the engine efficiency performance level and a reduced stall margin.The paper deals with the re-contouring of the leading edges of compressor airfoils by application of a new developed method for the profile definition. The common procedure of smoothing out the leading edges manually on a wheel grinding machine can not provide a defined contour nor a reproducible result of the overhaul process. In order to achieve optimized flow conditions in the compressor blade rows, suitable leading edge contours have to be defined for the worn airfoils. In an iterative process the flow behavior of these redesigned profiles is checked by numerical flow simulations and the shape of the profiles is improved. The following machining of the new defined leading edge contours is achieved on a grinding station handled by an appropriately programmed robot.

The performance deterioration of a high-speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported. A 0.025 mm (0.001 in.) thick rough coating with a surface finish of 2.54–3.18 rms μm (100–125 rms μin.) is applied to the pressure and suction surface of the rotor blades. Coating both surfaces increases the leading edge thickness by 10 percent at the hub and 20 percent at the tip. Application of this coating results in a loss in efficiency of 6 points and a 9 percent reduction in the pressure ratio across the rotor at an operating condition near the design point. To separate the effects of thickness and roughness, a smooth coating of equal thickness is also applied to the blade. The smooth coating surface finish is 0.254–0.508 rms μm (10–20 rms μin.), compared to the bare metal blade surface finish of 0.508 rms pm (20 rms μin.). The smooth coating results in approximately half of the performance deterioration found from the rough coating. Both coatings are then applied to different portions of the blade surface to determine which portions of the airfoil are most sensitive to thickness/roughness variations. Aerodynamic performance measurements are presented for a number of coating configurations at 60, 80, and 100 percent of design speed. The results indicate that thickness/roughness over the first 2 percent of blade chord accounts for virtually all of the observed performance degradation for the smooth coating, compared to about 70 percent of the observed performance degradation for the rough coating. The performance deterioration is investigated in more detail at design speed using laser anemometer measurements as well as predictions generated by a quasi-three-dimensional Navier–Stokes flow solver, which includes a surface roughness model. Measurements and analysis are performed on the baseline blade and the full-coverage smooth and rough coatings. The results indicate that adding roughness at the blade leading edge causes a thickening of the blade boundary layers. The interaction between the rotor passage shock and the thickened suction surface boundary layer then results in an increase in blockage, which reduces the diffusion level in the rear half of the blade passage, thus reducing the aerodynamic performance of the rotor.

Due to the complexity of the internal flow field of compressors, the aerodynamic design and optimization of a highly loaded axial compressor with high performance still have three problems, which are rich engineering design experience, high dimensions, and time-consuming calculations. To overcome these three problems, this paper takes an engineering-designed 2.5-stage highly loaded axial flow compressor as an example to introduce the design process and the adopted design philosophies. Then, this paper verifies the numerical method of computational fluid dynamics. A new Bezier surface modeling method for the entire suction surface and pressure surface of blades is developed, and the multi-island genetic algorithm is directly used for further optimization. Only 32 optimization variables are used to optimize the rotors and stators of the compressor, which greatly overcome the problem of high dimensions, time-consuming calculations, and smooth blade surfaces. After optimization, compared with the original compressor, the peak efficiency is still improved by 0.12%, and the stall margin is increased by 2.69%. The increase in peak efficiency is mainly due to the rotors. Compared with the original compressor, for the second-stage rotor, the adiabatic efficiency is improved by about 0.4%, which is mainly due to the decreases of total pressure losses in the range of above 30% of the span height and 10%–30% of the chord length. Besides, for the original compressor, due to deterioration of the flow field near the tip region of the second-stage stator, the large low-speed region eventually evolves from corner separation into corner stall with three-dimensional space spiral backflow. For the optimized compressor, the main reason for the increased stall margin is that the flow field of the second-stage stator with a span height above 50% is improved, and the separation area and three-dimensional space spiral backflow are reduced.

Purpose. Improving the quality of manufacturing of blanks for compressor rotor blades by hot extrusion. Research methods and equipment. The research was carried out using a crank press with a force of 1000 kN, in split dies in accordance with a serial technological process. The dies were heated up to 150 ... 200 °С, to improve the work when extruding the blanks of the rotor blades made from the titanium alloy ВT8. The thickness of the copper coating was measured with an ИTMП-3 magnetic induction device with an error of ± 2 μm. X-ray spectral microanalysis was performed on an ISM-6360ALA scanning microscope. The billets were heated in an MП-2В furnace. Results. It has been established that the quality of blade blanks made of ВT8 titanium alloy obtained by hot extrusion is influenced by the state of the copper coating, which is preliminarily applied to the surface of the original blank. When the initial blanks are heated, copper is oxidized and in the temperature range of 250…700 °С the oxidation rate proceeds according to a linear pattern, and after 700…750 °С – according to a parabolic pattern. Oxidation of the copper coating occurs unevenly not only within one workpiece, but also within the batch, which leads to a decrease in durability and deterioration of the surface quality of the blade workpieces obtained by hot extrusion. Research carried out by X-ray spectral microanalysis of the copper coating revealed the presence of aluminum oxides of varying degrees of dispersion. The source of this material in the copper coating is caricature of corundum used in blowing into the surface of the billet, which is the reason for the appearance of scoring on the blade blank. It was also found that longitudinal marks on the blade are a consequence of the appearance of a matrix of tubercles (sagging) on the working surface of the die, caused by the adhesion of the deformable material of the blade to the base metal of the tool. Scientific novelty. The regularity of the influence of the heating temperature of the initial blank of the blade on the oxidation rate of the copper coating has been established. The mechanism of the influence of the oxidation of the copper coating and the adhesion of contacting materials during hot extrusion on the surface condition of the resulting blanks is disclosed. Practical value. The results obtained make it possible to improve the quality of the manufactured blanks of the compressor rotor blades by hot extrusion.

The efficiency of gas turbines used in power plants is largely dependent on their aerodynamic performance. The components like stators and rotors of the turbines and compressor are subject to abrasive and erosive wear. There are various contaminants which get deposited over the blades and produce roughnesses. The roughness magnitude both in turbines & compressors varies along the height and chord of blade and also over different stages of turbine. In actual turbines and compressors, roughness is not only found over entire surfaces of the blades but over a small portion of the surface of the blades also. The roughness is found, in the form of bands also on leading edge, middle chord and trailing edge of pressure and suction surfaces of blades of turbines and compressors. The flow through turbines and compressors is inherently three dimensional due to the vane/blade passage geometry and other variety of reasons. The blade profile continuously changes in the span wise direction. The flow structure, along the end walls, is strongly three dimensional. This effect is tremendous in case of lower span to chord ratio. These characteristics of the flow lead to increase profile losses which in turn adversely affect the efficiency of turbomachines. The secondary flows cause to generate a non-uniform flow at exit of the blade row thereby efficiency of the blade row downstream gets further reduced. This research work in fact is an attempt to capture complex three- dimensional secondary flow vortices near end wall region along with getting results for total losses. The total losses are segregated to obtain profile & secondary loss i.e. ends losses numerically. The Computational fluid dynamics (CFD) uses numerical methods and algorithms to solve and analyze problems that involve fluid flows. The present research work is carried out using the CFD, commercial softwares, Gambit 2.4.6® and FLUENT 6.2.16®. These softwares are used for designing stage and working on it without any actual manufacturing and installation of such cascade in real working situation. The cascades of turbine and compressor are simulated to carry out study of effect of blade deterioration on various losses using three different blade profiles for turbines and one profile for compressor. Total three number blade profiles titled 6030, 5530 and 3525 as selected by Samsher [2002], from impulse and reaction turbine are chosen. Of the three profiles selected, blade profile 3525 was nearly impulse type and blade profiles titled 5530 and 6030 iv were of reaction type with different degree of reactions. The study is conducted for a number of cascades. There are total 13 numbers of cascades to be simulated separately for each of combinations of roughness magnitude, location of roughness and the given single blade profile for application of roughness on entire surface. In addition study of localised roughnesses of varying magnitudes is also conducted. There are 6 numbers of locations over the suction and pressure surfaces of given cascade for application of localised roughness. The total pressure at inlet and total pressure and static pressure at exit measurement planes for numerous number of cascades are measured with the help of 'Fluent' software. These parameters are required to calculate local loss coefficients relative to nondimensional distance in the pitch wise direction along the measurement plane. The mass averaged loss coefficients is representative loss coefficient for selected pitch wise positions and calculated using the pitch wise local loss coefficients. The total loss, secondary loss and profile loss, for a cascade, are calculated on the basis of mass averaged loss coefficients for all selected span wise positions. The mass averaged loss coefficients near the end walls at both ends are higher than their values at mid span of the blade for each of cascade for all blade profiles i.e. 6030, 5530 and 3525. The local increase in mass averaged loss coefficients is observed due to the secondary flow cores near the hub and casing for all cascades. The results with regard to the total, profile and secondary losses for BSR and PSR cascade based on mass averaged loss coefficients show that the magnitudes of total loss for these cascades for all roughness values are higher than that of the smooth cascade for all blade profiles i.e. 6030, 5530 and 3525. The total loss increases as roughness increases on the blade surfaces of each cascade for all blade profiles. The losses increase in the same order when roughness was increased from lower roughness value to high roughness value for all type of cascades. It is observed that change in respective losses with the increase in magnitude of roughness at high roughness values (such as 750 μm) are negligibly small. The increasing of roughness on suction surface is found to be more detrimental than the same for pressure surface in terms of generation of total loss. It is found that the total loss increases as the roughness is increased on suction surfaces. The effect is combined when roughness is applied on both the surface of the blades. The v contribution of the profile loss in the total loss increases as roughness increases on the blade surfaces. The secondary loss increases with the increase of roughness on pressure surfaces of blades of cascades. The same is decreased with the increase of roughness on suction surfaces of blades of cascades. The combined effect of increase of roughness on both the surfaces of blades is seen. The effect of roughness on secondary loss is more pronounced in blade profile 6030. The results also show that shape or geometry of the blades of the cascades significantly affect the losses. The localised roughness and the effect of the same on losses vary from one blade profile to other blade profile. The application of localised roughness on leading edge on pressure surfaces of the cascade employing blade profile 3525 leads to more generation of total and profile losses than localised roughness on middle chord and trailing edge of same surfaces for the same cascade. On the other hand, the localised roughness on trailing edge of the pressure surfaces for blade profile 6030 and 5530 leads to more generation of total and profile losses. The application of localised roughness on blade profile 6030 and 5530 effect the phenomena of losses generation in same way. The localised roughness on various surfaces of blade profile 3525 effect the phenomenon of losses generation differently comparing with that of blade profile 6030 and 5530. The trend of increase in total loss with the increase of surface roughness on various surfaces for compressor cascade is similar to that of turbine cascades. It is noticeable that the profile loss contributes very significantly in the total loss for all type of compressor cascades and that the roughness magnitudes do not affect secondary loss very appreciably for all type of cascades.

During the operation of a jet engine, deterioration will constantly reduce its performance. This results in an increase in specific fuel consumption (SFC) and exhaust gas temperature (EGT); the main characteristics to describe the efficiency of a jet engine. Thereby, the high pressure compressor (HPC) is particularly affected by deterioration. Multiple effects take place and decrease the efficiency of the HPC. Erosion is one of the main effects and leads to thinner or thicker leading- and trailing edges, thinner airfoils, a reduction of chord length and an increase in tip clearance. In addition, erosion and fouling may also lead to increased surface roughness on airfoils and endwalls. An additional parameter which is also dependent on the on-wing time are changes in the stagger angle of the different blade heights. The objective is to estimate the quantitative effect of the different wear mechanisms on the stage parameters, like throttle line and efficiency. Therefore, a geometry setup process is implemented to create HPC blade models with independent values of erosion. With these blades, CFD calculations based on realistic boundary conditions were carried out with the CFD solver ANSYS CFX. It could be proven that the deterioration of leading edge thickness has the major influence on stage performance, followed by the max. profile thickness and the stagger angle. The operational blade deterioration of leading edge thickness leads to an efficiency range of about 0.173 %. Moreover, the deterioration of stagger angle leads to an offset of the throttle lines towards higher or smaller loadings, depending on the direction of change.

Turbines and compressors operating in polluted atmosphere with solid particles are subjected to performance deterioration. This paper presents an investigation carried out on two-stage gas turbine with blunt leading edge blades and on a single-stage axial flow compressor to study the effects of particulates and erosion on performance deterioration.

Gas turbine performance degradation over time is mainly due to the deterioration of compressor and turbine blades, which, in turn, causes a modification of the compressor and turbine performance maps. Since detailed information about the actual modification of the compressor and turbine performance maps is usually unavailable, component performance can be modeled and investigated (i) by scaling the overall performance map, or (ii) by using stage-by-stage models of the compressor and turbine and by scaling each single stage performance map to account for each stage deterioration, or (iii) by performing 3D numerical simulations, which allow to both highlight the fluid-dynamic phenomena occurring in the faulty component and grasp the effect on the overall performance of the component. In this paper, the authors address the most common and experienced source of loss for a gas turbine, i.e. compressor fouling. With respect to the traditional approach, which mainly aims at the identification of the overall effects of fouling, authors investigate a micro-scale representation of compressor fouling (e.g. blade surface deterioration and flow deviation). This allows (i) a more detailed investigation of the fouling effects (e.g. mechanism, location along blade height, etc.), (ii) a more extensive analysis of the causes of performance deterioration and (iii) the assessment of the effect of fouling on stage performance coefficients and on stage performance maps. The effects of a non-uniform surface roughness on both rotor and stator blades of an axial compressor stage are investigated by using a commercial CFD code. The NASA Stage 37 test case was used as the baseline geometry. The numerical model already validated against experimental data available in literature was used for the simulations. Different non-uniform combinations of surface roughness levels on rotor and stator blades were imposed. This makes it possible to highlight how the localization of fouling on compressor blades affects compressor performance, both at an overall and at a fluid-dynamic level.

This paper describes a methodology for obtaining correct blade geometry of high aspect ratio axial compressor blades during running condition taking into account of blade untwist and bending. It discusses the detailed approach for generating cold blade geometry for axial compressor rotor blades from the design blade geometry using fluid structure interaction technique. Cold blade geometry represents the rotor blade shape at rest, which under running condition deflects and takes a new operating blade shape under centrifugal and aerodynamic loads. Aerodynamic performance of compressor primarily depends on this operating rotor blade shape. At design point it is expected to have the operating blade shape same as the intended design blade geometry and a slight mismatch will result in severe performance deterioration. Starting from design blade profile, an appropriate cold blade profile is generated by applying proper lean and pre-twist calculated using this methodology. Further improvements were carried out to arrive at the cold blade profile to match the stagger of design profile at design operating conditions with lower deflection and stress for first stage rotor blade. In rear stages, thermal effects will contribute more towards blade deflection values. But due to short blade span, deflection and untwist values will be of lower values. Hence difference between cold blade and design blade profile would be small. This methodology can especially be used for front stage compressor rotor blades for which aspect ratio is higher and deflections are large.

Gas turbine axial compressor blades may encounter damage during service for various reasons. Debris from casing or foreign objects may impact blades causing damage near the rotor’s tip. This may result in deterioration of performance and reduction in the surge margin. Ability to assess the effect of damaged blades on the compressor performance and stability is important at both the design stage and in service. The damage to compressor blades breaks the cyclic symmetry of the compressor assembly. Thus computations have to be performed using the whole annulus. Moreover, if rotating stall or surge occurs, the downstream boundary conditions are not known and simulations become difficult. This paper presents an unsteady CFD analysis of compressor performance with tip curl damage. Tip curl damage typically occurs when rotor blades hit a loose casing liner. The computations were performed up to the stall boundary, predicting rotating stall patterns. The aim is to assess the effect of blade damage on stall margin and provide better understanding of the flow behaviour during rotating stall. Computations for the undamaged rotor are also performed for comparison. A transonic axial compressor rotor is used for the time-accurate numerical unsteady flow simulations, with a variable choked nozzle downstream simulating an experimental throttle. One damaged blade was introduced in the rotor assembly and computations were performed at 60% of the design rotational speed. It was found that there is no significant effect on the compressor stall margin due to one damaged blade despite the differences in rotating stall patterns between the undamaged and damaged assemblies.

This is the second part of a two-part paper. It presents four case studies. The first case is the redesign of a transonic rotor (NASA rotor 67) at a post peak efficiency operating point. The second case is a redesign of a transonic compressor stage originally designed by DLR. The redesign is carried out at the stage peak efficiency point. The third and fourth cases look at the redesign of blade rows within a three-stage transonic test compressor that was originally designed by Siemens Industrial Turbomachinery Ltd known as the ATC compressor. Specifically the third case is a redesign of the IGV-rotor-stator configuration. It is carried out at two operating points: one is at the stage peak efficiency point; the other is at a lower stagnation pressure ratio choked flow point. Initially the redesign at the stage peak efficiency point produces considerable efficiency gain, but leads to noticeably reduced choked mass flow rate. The redesign at a near choked mass flow rate point, on the other hand, leads to considerable performance deterioration at operating points with lower mass flow rate, though the choked mass flow rate is even increased. Subsequently, a parallel multi-point approach has been implemented. Results show that a two-point design optimization avoids unacceptable performance deterioration at off design conditions. In the fourth case a redesign is applied across all 7 blade rows of the ATC compressor at the compressor design point. All these case studies are aimed to increase isentropic efficiency whilst meeting the specified constraints.

Abstract The wear and damage of High-Pressure Compressor (HPC) blades due to erosion or Foreign Object Damage (FOD) have a significant influence on HPC aerodynamic performance, vibration resistance against High-Cycle Fatigue (HCF) and thus component lifetime. The changes in airfoil geometry reduce the overall engine efficiency. Furthermore extended off-wing engine maintenances due to blade failures are increasing the cost of ownership. The safe operation of every engine within a reduced number of shop visits requires a reliable prediction of future deterioration. This enables the optimization of services and off-wing time. One contribution to this is a better understanding of the component’s dynamics and based on this providing an improved wear modeling to reliably predict the remaining lifetime and the decreased efficiency. This contribution determines the material removal of HPC blades due to sand erosion. Originally, this stage was built as a blisk (Blade Integrated Disk). After sand erosion test completion, the blisk was cut into segments containing one airfoil only. First, the material removal is determined for ten blades of one exemplary rotor. A blue light fringe projector is employed to identify the geometrical differences between the eroded blades and the nominal design. Second, realistic finite element models are generated to enable comparable modal analyses of eroded blades. This procedure suffers from unavoidable and mostly random imperfections due to the manufacturing process, which significantly affects the blade surface before the erosion test can be conducted. Therefore, an already published approach is implemented in the third step to predict the blade surface after erosion based on nominal blade design. The investigation is completed by comparing measured and predicted surfaces. Finally, the aforementioned tool is employed to predict the locations and intensities of the material losses and the accompanying change in modal properties of this compressor blade concerning operational time.

Fouling of compressor blades is an important mechanism leading to performance deterioration in gas turbines over time. Experimental and simulation data are available for the impact of specified amounts of fouling on performance, as well as the amount of foulants entering the engine for defined air filtration systems and ambient conditions. This study provides experimental data on the amount of foulants in the air that actually stick to a blade surface for different conditions of the blade surface. Quantitative results both indicate the amount of dust as well as the distribution of dust on the airfoil, for a dry airfoil, as well as airfoils that were wet from ingested water, as well as different types of oil. The retention patterns are correlated with the boundary layer shear stress. The tests show the higher dust retention from wet surfaces compared to dry surfaces. They also provide information about the behavior of the particles after they impact on the blade surface, showing that for a certain amount of wet film thickness, the shear forces actually wash the dust downstream, and off the airfoil. Further, the effect of particle agglomeration of particles to form larger clusters was observed, which would explain the disproportional impact of very small particles on boundary layer losses.

Inlet air filters are widely used to remove solids and liquid droplets from the ambient air before it enters the compressor of a gas turbine. Clean inlet air provides many advantages: Less corrosion of the compressor and of gas-path hot parts, such as the turbine, decreased compressor fouling, less erosion of the compressor bladeThese in turn prevent deterioration of output and heat rate, and reduce maintenance costs. Compressor fouling is caused by the ingestion of substances that deposit and adhere to blade surfaces, resulting in reduced aerodynamic efficiency and decreased available output. Air contamination could be significantly reduced by the use of more efficient air filtration systems, especially through the reduction of the quantity of smaller particles ingested. The consequent lower loss of output power and decreased cleaning efforts provide lower costs of operation and increased shaft power. This work was composed of three major efforts: 1) A literature search was performed to establish the state of the art for particle removal from gases, particularly by electrostatic precipitation, and to identify the leading vendors of the equipment-considering both experience and technical expertise. 2) Two chosen companies were visited to determine their technical capabilities as they apply to gas turbine inlet air filtration. 3) A representative gas turbine was specified by PRCI as being the equivalent of a GE Model 3002J turbine, with airflow of 91,200 acfm. A specification based upon that airflow was prepared and submitted to the two vendors. Each vendor prepared a proposal for a filter system compliant with the specification. The proposed air filtration equipment is sufficiently different from existing products that it was judged not beneficial to visit manufacturing facilities. Both vendors are reputable suppliers of air filtration equipment. This study is intended to provide definitive information relative to the use of new technology for air inlet filtration on gas turbines in gas pipeline pumping applications.