Дисертації з теми "Compressors stage"

Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2015.
Cataloged from PDF version of thesis.
Includes bibliographical references (pages 74-75).
Loss generation features for the first stage impeller in a multistage centrifugal compressor are examined using three-dimensional RANS computations. The calculations were carried out for a baseline configuration and for seven other impeller configurations, with the constraints of constant mass flow and constant work per unit mass flow. The computations showed an 8 percent reduction in loss, compared to the baseline, for a configuration that incorporated 60% of the total casing blade angle change in the front 20% chord. Twodimensional interactive boundary layer computations were carried out to demonstrate links between the loss variation and the changes in boundary layer behavior in the front 20% of the blade passage.
by Kristina Barile.
S.M.

Thesis (M.S. in Mechanical Engineering)--Naval Postgraduate School, June 2009.
Thesis Advisor(s): Gannon, Anthony J. "June 2009." Description based on title screen as viewed on July 13, 2009. DTIC Identifiers: TCR (Transonic Compressor Rig). Author(s) subject terms: Compressor, Transonic, Stall, Surge. Includes bibliographical references (p. 73-74). Also available in print.

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2012.
Cataloged from PDF version of thesis.
Includes bibliographical references (p. 61-63).
This thesis presents concepts for improving the performance of return channels in multi-stage centrifugal compressors. Geometries have been developed to reduce both separation and viscous losses. A number of different features with potential to reduce separation have also been investigated. The final proposed geometry uses a vaneless diffuser which narrows on the shroud side at the beginning of the 180' bend, an axially extended 1800 bend with increasing radius of curvature, and return channel vane leading edge radial position at an increased radius compared to the baseline. Three-dimensional calculations showed a 9% loss reduction compared to previous work [1], with a cumulative loss reduction of 19% compared to a baseline geometry. The geometry developed was based on specified inlet conditions. To examine the potential for increased performance if this constraint was removed, a return channel geometry was also defined that incorporated the same features but allowed modified inlet conditions, specifically radial inlet flow. The design of the impeller required for this new inlet flow was not considered. An overall loss reduction of 23% compared to baseline was found from the calculations. Modification of the impeller geometry is thus proposed as future work.
by Anne-Raphaëlle Aubry.
S.M.

A one-dimensional, stage-by-stage axial compression system mathematical model has been constructed which can describe system behavior during post-stall events such as surge and rotating stall. The model uses a numerical technique to solve the nonlinear conservation equations of mass, momentum, and energy. Inputs for blade forces and shaft work are provided by a set of quasi-steady stage characteristics modified by a first order lagging equation to simulate dynamic stage characteristics. The model was validated with experimental results for a three-stage, low-speed compressor and a nine-stage, high-pressure compressor. Using these models, a parametric study was conducted to determine the effect of inlet resistance, combustor performance, heat transfer, and stage characteristic changes due to hardware modification on post—stall system behavior.
Ph. D.
incomplete_metadata

A one-dimensional, stage-by-stage axial compression system mathematical model has been constructed which can describe system behavior during post-stall events such as surge and rotating stall. The model uses a numerical technique to solve the nonlinear conservation equations of mass, momentum, and energy. Inputs for blade forces and shaft work are provided by a set of quasi-steady stage characteristics modified by a first order lagging equation to simulate dynamic stage characteristics. The model was validated with experimental results for a three-stage, low-speed compressor and a nine-stage, high-pressure compressor. Using these models, a parametric study was conducted to determine the effect of inlet resistance, combustor performance, heat transfer, and stage characteristic changes due to hardware modification on post—stall system behavior.
Ph. D.
incomplete_metadata

The numerical simulation of a three-dimensional viscous flow in cascade of the axial compressor of low pressure of the gas-turbine engine is presented. The results of a flow in the first stage o f the compressor in nonstationary three-dimensional statement are obtained in the solver F. Velocity and pressure fields are received as a result.

The objectives of this research are the prediction of the thermodynamic behavior of a multi-stage intercooled reciprocating compressor and its progressive loss of performance due to leakage. A theoretical model is developed to simulate the thermodynamics of the compressor system and the lubricating condition and wear of the piston ring pack for a multi-stage intercooled reciprocating compressor. A first law of thermodynamics approach is used to determine the thermodynamic properties of the gas inside the cylinders, the intercoolers and the inlet and discharge manifold. The compressor valves are modeled as single degree-of-freedom, spring-mass=damper systems. The flows through the valves are calculated based on the steady flow equations for equivalent orifices. The lubricating condition of the piston ring pack are determined on the basis of hydrodynamic lubrication theory. The wear of the piston rings is assumed to occur when the hydrodynamic oil film between the piston ring and cylinder bore breaks down. Based on the theoretical model, a computer program is developed. This program is tested on an IngersoH-Rand Model 242, two stage aircooled reciprocating air compressor. The comparison of the experimental values of the pressure variations in the first cylinder with the value predicted by the computer program shows a reasonable match. The computer program predicts the pressure, temperature and mass flow rates for each cylinder and the intercooler. Also predicted is the wear rate of each piston ring. The progressive loss in the compressor mass discharge, and hence the loss in its performance, is determined by calculating the leakage losses several times, updating the leakage area each time based on the wear rate of the piston rings. The result shows a drop of about 15 percent in the discharge rate of the Model 242 compressor after 8000 hours of running time.
Master of Science

The correlation of cascade experimental data is one method for obtaining compressor stage characteristics. These correlations specify pressure loss and flow turning caused by the blades. Current open literature correlations used in streamline curvature codes are inadequate for general application to high-speed transonic axial-flow compressors. The objective of this research was to investigate and evaluate the available correlations and ultimately discover sets of correlations which best fit the empirical data to be used in streamline curvature codes. Correlations were evaluated against experimental data from NASA Rotor 1-B and NASA Stage 35. It was found that no universal set of correlations was valid for minimum-loss point predictions. The Bloch shock loss model showed promising results in the stall regime for supersonic relative inlet Mach numbers. The Hearsey and Creveling off-minimum-loss deviation angle prediction performed consistently better than all other correlations tested.
Master of Science

The need was identified to develop an analytical performance prediction code for subsonic multistage axial compressors that can be included in network analysis software. It was found that performance calculations based on an elementary one-dimensional meanline prediction method could achieve remarkable accuracy, provided that sound models are used for the losses, deviation and the onset of rotating stall. Consequently, this study focuses on gaining more expertise on the modelling of losses in such compressors through investigating the mechanisms responsible, the methods of predicting them, their implementation and possible usage. Internal losses are seen as mechanisms that increase the entropy of the working fluid through the compressor and it was found that, at a fundamental level, all internal losses are a direct result of viscous shearing that occurs wherever there are velocity gradients. Usually the methodology employed to predict the magnitudes of these mechanisms uses theoretically separable loss components, ignoring the mechanisms with negligible velocity gradients. For this study these components were presented as: Blade profile losses, endwall losses including tip leakage and secondary losses, part span shroud losses, other losses, losses due to high subsonic Mach numbers and incidence loss. A preliminary performance prediction code, with the capability of interchanging of the different loss models, is presented. Verification was done by comparing the results with those predicted by a commercial software package and the loss models were evaluated according to their ease of implementation and deviation from the predictions of the commercial package. Conclusions were made about the sensitivity of performance prediction to using the different loss models. Furthermore, the combination of loss models that include the most parameters and gave the best comparison to the commercial software predictions was selected in the code to perform parametric studies of the loss parameters on stage efficiency. This was done to illustrate the ability of the code for performing such studies to be used as an aid in understanding compressor design and performance or for basic optimization problems. It can therefore be recommended that the preliminary code can be implemented in an engineering tool or network analysis software. This may however require further verification, with a broader spectrum of test cases, for increased confidence as well as further study regarding aspects like multi-stage annulus blockage and deviation
Thesis (M.Ing. (Mechanical Engineering))--North-West University, Potchefstroom Campus, 2004.

Thesis (PhD)--Stellenbosch University, 2012.
ENGLISH ABSTRACT: The aims of this thesis are to identify all possible modes of operaton for a multi-stage axial flow compressor; then to characterise the performance, attempt to numerically model operation, and determine the main flow field features for each mode. Four quadrant axial flow compressor operation occurs when the direction of flow through the compressor or the sign of the pressure difference across the compressor reverses, or any combination of these. Depending on the direction of rotation of the compressor, six modes of operation are possible in the four quadrants of the performance map. The rotor rotates in the design direction for three modes, and in the opposite direction for the other three. The stationary-rotor pressure characteristic is S-shaped and passes through the second and fourth quadrants. A three-stage axial flow compressor operating in the incompressible flow regime was used for the experimental investigation. Flow through the compressor was reversed or augmented by means of an auxiliary axial flow fan. Compressor performance was measured by means of static pressure tappings, a turbine anemometer calibrated to measure forward and reversed volumetric flow and a load cell for torque measurement. The inter-blade row flow fields were measured with pneumatic probes and 50 μm cylindrical hot film probes. Three dimensional single blade-passage Navier-Stokes simulations were performed using the Numeca FineTurbo package. Steady state simulations used a mixing plane approach. A nonlinear harmonic approximation was used for time-unsteady simulations. Unstalled first quadrant operation was unremarkable, and good agreement was obtained between experimental and numerical data. A single stall cell was detected experimentally during stalled operation, which was not modelled numerically. In the fourth quadrant for positive rotation, (windmilling), the compressor acts as an inefficient turbine. Flow separates from the pressure surface of the blade, rendering the steady-state mixing plane approach unsuitable. The performance characteristic curves for second quadrant for positive rotation, are discontinuous with those of first quadrant operation. The temperature rise in the working fluid is significantly higher than at design point. Periodic flow structures occurring across two blade passages were detected at all flow coefficients investigated, invalidating numerical modelling assumptions. Better agreement was obtained between experimental and numerical data from a case found in literature. If the compressor operates as a compressor in reverse (third quadrant operation), significant separation occurs on the pressure surface of all blades, and flow conditions resemble severe first quadrant stall. Separation becomes less severe at larger flow rates, allowing numerical simulation, though this is sensitive to the initial flow field. In the the part of the second quadrant, where the compressor rotates in reverse, it operates as a turbine. The blade angles and the direction of curvature match the flow angles and turning well, leading to high turbine efficiencies. Numerical simulations yielded good agreement with measured results, but were again sensitive to the initial flow field. Fourth quadrant operation with negative rotation occurs when flow is forced through the compressor in the design direction. Large separation bubbles are attached to the pressure surfaces of rotor and stator blades, so virtually all throughflow occurs near the hub and casing
AFRIKAANSE OPSOMMING: Die doelwitte van hierdie tesis is om al die moontlike werkmodusse vir ’n bestaande multi-stadium aksiaalvloei kompressor uit te ken; om dan die gedrag te gekarakteriseer, ’n poging aan te wend om die werking numeries te modelleer, en die belangrikste vloeiveldkenmerke vir elke modus te bepaal. Vier-kwadrant aksiaalvloei kompressor werking vind plaas as die rigting van die vloei deur die kompressor, of die teken van die drukverskil oor die kompressor omkeer, of enige kombinasie daarvan. Afhangende van die rigting van rotasie van die kompressor is ses operasionele modusse moontlik in die vier kwadrante van die kompressorkaart. Die rotor draai in die ontwerprigting vir drie van die modes, en in die teenoorgestelde rigting vir die ander drie. Die stilstaande-rotor drukkarakteristiek is S-vormig gaan deur die tweede en vierde kwadrante. ’n Drie-stadium onsamedrukbare vloei aksiaalvloei kompressor is vir die eksperimentele ondersoek gebruik. Vloei deur die kompressor is omgekeer of aangehelp deur middel van ’n aksiaalvloei hulpwaaier. Kompressor werking is gemeet deur middel van statiese druk meetpunte in die omhulsel, ’n turbine anemometer wat gekalibreer is om vorentoe en omgekeerde volumetriese vloei te meet, en ’n lassel vir wringmoment meting. Interlemryvloeivelde is opgemeet met pneumatiese sensors en 50-μm silindriese warm film sensors. Drie-dimensionele Navier-Stokes simulasies is uitgevoer vir ’n enkele lem van elke lemry, met behulp van die Numeca FineTurbo sagtewarepakket. ’n Mengvlakbenadering is gebruik vir bestendige toestand simulasies, terwyl ’n nie-linere harmoniese benadering gebruik is vir die tyd-afhanklike simulasies. Ongestaakte eerste kwadrant werking was alledaags, en goeie ooreenkoms is gevind tussen die eksperimentele en numeriese data. ’n Enkele staak-sel is eksperimenteel ontdek tydens gestaakte werking. Gestaakte werking is nie numeries gemodelleer nie. In die vierde kwadrant vir positiewe rotasie, (”windmeulwerking”), werk die kompressor as ’n ondoeltreffende turbine. Vloei-wegbrekinging op die lem drukoppervlaktes maak die bestendige toestand mengvlakbenadering ongeskik. In die kenlyne vir tweede kwadrant positiewe rotasie, is daar ’n diskontinu¨ıteit in die prestasie karakteristiekkrommes vir die eerste en tweede kwadrant werking. Die temperatuurstyging in die werk- vloeistof is beduidend ho¨er as as by die ontwerppunt. Periodiese vloeistrukture wat oor twee lemme plaasvind is gevind by alle vloei ko¨effisi¨ente wat ondersoek is, en dit maak die numeriese modellering aannames ongeldig. Beter ooreenkoms tussen die eksperimentele en numeriese data iss verkry met ’n geval wat uit die literatuur gevind is. Indien die kompressor werk as ’n kompressor in omgekeerde (derde kwadrant weking), vind beduidende wegbreking op die drukoppervlak van al die lemme plaas, wat lyk soos ernstige gestaakte eerste kwadrant werking. Die vloeiskeiding raak minder ernstig by ’n groter vloeitempo, wat numeriese nabootsing toelaat, maar die nabootsings is sensitief vir die aanvanklike vloeiveld. In die tweede kwadrant, by omgekeerde rotasie, werk die kompressor as ’n turbine. Die lemhoeke en die rigting van lemkromming stem ooreen met die vloeihoeke en verwringing, wat lei tot ho¨er turbine doeltreffendheid. Numeriese nabootsings stem goed ooreen met gemete resultate, maar is weereens sensitief vir die keuse van die aanvanklike vloeiveld. Vierde kwadrant werking met negatiewe rotasie vind plaas wanneer die lug gedwing word om deur die kompressor in die ontwerprigting te vloei. Groot skeidingborrels sit vas aan die drukoppervlaktes van alle lemme, sodat meeste deurvloei naby die naaf en die omhulsel plaas vind.

[ES] Esta tesis doctoral presenta un estudio de compresores scroll con inyección de vapor (SCVI) para bombas de calor que operan en climas fríos o para aplicaciones de calentamiento de agua a alta temperatura. Para ello, se comparó experimentalmente un SCVI con un compresor de dos etapas de pistones (TSRC) trabajando con R-407C en condiciones extremas. La comparación se realizó en términos de eficiencias del compresor, capacidad, COP y rendimientos estacionales tanto para el modo calefacción como para el modo refrigeración. Los resultados proporcionan una idea general sobre el rango de aplicación de los compresores estudiados y sobre las diferencias en los rendimientos de los compresores. Sin embargo, se identificaron varias limitaciones en la caracterización de los compresores y en el análisis del ciclo. Esto motivó a profundizar en el estudio del ciclo de compresión de dos etapas y sus componentes. El siguiente paso fue realizar un análisis teórico de los ciclos de compresión de dos etapas para aplicaciones de calefacción, en donde se identificó a la presión intermedia y a la relación de inyección como los parámetros del sistema más influyentes sobre el COP. La presión intermedia se optimizó para dos configuraciones de inyección (tanque de separación y economizador) utilizando varios refrigerantes. Basándose en los resultados de la optimización, se propuso una correlación que permite obtener la presión intermedia óptima del ciclo, considerando la influencia del subenfriamiento a la salida del condensador. Además, se analizó la influencia del diseño de los componentes del sistema sobre el COP del ciclo. Posteriormente, el estudio se profundizó a nivel de componentes. El factor más crítico en el sistema es el rendimiento del compresor. Por lo tanto, el siguiente paso fue evaluar la influencia de varios sistemas de compresión con inyección de vapor sobre el COP. Se tomaron en cuenta tres tecnologías de compresores, un SCVI, un TSRC y un compresor scroll de dos etapas (TSSC). Estas tecnologías de compresores fueron caracterizadas y modeladas para estudiar su rendimiento. Para ello, se propuso una nueva metodología para caracterizar compresores scroll con inyección de vapor. Esta metodología permite evaluar el rendimiento del compresor independientemente del mecanismo de inyección que se utiliza en el ciclo. Se identificó una correlación lineal entre la relación de inyección de refrigerante y la relación de compresión intermedia. Esta correlación se utiliza para determinar el flujo másico de inyección en función de la presión intermedia. Posteriormente, se propuso un modelo semi-empírico de compresores scroll y una metodología para extender dicho modelo para compresores scroll con inyección de vapor. Los modelos fueron ajustados y validados usando datos experimentales de cuatro compresores scroll trabajando con R-290 y un SCVI trabajando con R-407C. Finalmente, se comparó un SCVI con dos compresores de dos etapas, un TSSC y un TSRC, trabajando en condiciones extremas. Se optimizó la relación de volúmenes de los compresores de dos etapas. Los resultados muestran que, en las condiciones nominales de funcionamiento (Te=-15 °C, Tc=50 °C), la relación de volúmenes óptima del TSSC es 0.58, y del TSRC es 0.57. El TSSC consigue un COP 6% mayor que el SCVI y un COP 11.7% mayor que el TSRC. Bajo un amplio rango de condiciones de operación, el SCVI presenta una mejor eficiencia y COP para relaciones de presión inferiores a 5. Para relaciones de presión más altas, el TSSC presenta mejor rendimiento y consigue una temperatura de descarga más baja. Se concluye que el SCVI es una solución fácil de implementar, desde el punto de vista del mecanizado, y que permite extender el mapa de trabajo de los compresores de una etapa. Sin embargo, los resultados muestran que la compresión en dos etapas consigue mejorar en mayor medida el COP del ciclo y la capacidad, con una mayor redu
[CA] Aquesta tesi doctoral presenta un estudi de compressors scroll amb injecció de vapor (SCVI) per a bombes de calor que operen en climes freds o per a aplicacions d'escalfament d'aigua a alta temperatura. Per a això, es va comparar experimentalment un SCVI amb un compressor de dues etapes de pistons (TSRC) treballant amb R-407C en condicions extremes. La comparació es va realitzar en termes d'eficiències del compressor, capacitat, COP i rendiments estacionals tant per al mode calefacció com per al mode refrigeració. Els resultats proporcionen una idea general sobre el rang d'aplicació dels compressors estudiats i sobre les diferències en els rendiments dels compressors. No obstant això, es van identificar diverses limitacions en la caracterització dels compressors i en l'anàlisi del cicle. Això va motivar a aprofundir en l'estudi del cicle de compressió de dues etapes i els seus components. El següent pas va ser realitzar una anàlisi teòrica dels cicles de compressió de dues etapes per a aplicacions de calefacció, on es va identificar la pressió intermèdia i la relació d'injecció com els paràmetres del sistema més influents sobre el COP. La pressió intermèdia es va optimitzar per a dues configuracions d'injecció (tanc de separació i economitzador) utilitzant diversos refrigerants. Basant-se en els resultats de l'optimització, es va proposar una correlació que permet obtindre la pressió intermèdia òptima del cicle, considerant la influència del subrefredament a l'eixida del condensador. A més, es va analitzar la influència del disseny dels components del sistema sobre el COP del cicle. Posteriorment, l'estudi es va aprofundir a nivell de components. El factor més crític en el sistema és el rendiment del compressor. Per tant, el següent pas va ser avaluar la influència de diversos sistemes de compressió amb injecció de vapor sobre el COP. Es van prendre en compte tres tecnologies de compressors, un SCVI, un TSRC i un compressor scroll de dues etapes (TSSC). Aquestes tecnologies de compressors van ser caracteritzades i modelades per a estudiar el seu rendiment. Per a això, es va proposar una nova metodologia per a caracteritzar compressors scroll amb injecció de vapor. Aquesta metodologia permet avaluar el rendiment del compressor independentment del mecanisme d'injecció que s'utilitza en el cicle. Es va identificar una correlació lineal entre la relació d'injecció de refrigerant i la relació de compressió intermèdia. Aquesta correlació s'utilitza per a determinar el flux màssic d'injecció en funció de la pressió intermèdia. Posteriorment, es va proposar un model semi-empíric de compressors scroll i una metodologia per a estendre aquest model per a compressors scroll amb injecció de vapor. Els models van ser ajustats i validats utilitzant dades experimentals de quatre compressors scroll treballant amb R-290 i un SCVI treballant amb R-407C. Finalment, es va comparar un SCVI amb dos compressors de dues etapes, un TSSC i un TSRC, treballant en condicions extremes. Es va optimitzar la relació de volums dels compressors de dues etapes. Els resultats mostren que, en les condicions nominals de funcionament (Te=-15 °C, Tc=50 °C), la relació de volums òptima del TSSC és 0.58, i del TSRC és 0.57. El TSSC aconsegueix un COP 6% major que el SCVI i un COP 11.7% major que el TSRC. Sota un ampli rang de condicions d'operació, el SCVI presenta una millor eficiència i COP per a relacions de pressió inferiors a 5. Per a relacions de pressió més altes, el TSSC presenta millor rendiment i aconsegueix una temperatura de descàrrega més baixa. Es conclou que el SCVI és una solució fàcil d'implementar, des del punt de vista del mecanitzat, i que permet estendre el mapa de treball dels compressors d'una etapa. No obstant això, els resultats mostren que la compressió en dues etapes aconsegueix millorar en major mesura el COP del cicle i la capacitat, amb una major reducció de la
[EN] This Ph.D. thesis presents a study of scroll compressors with vapor-injection (SCVI) for heat pumps operating in cold climates or in high-temperature water heating applications. To do so, firstly, an SCVI was experimentally compared with a two-stage reciprocating compressor (TSRC) working with R-407C under extreme conditions. The comparison was made in terms of compressor efficiencies, capacity, COP, and seasonal COP, both for heating and cooling modes. The results give a general idea about the application range of the studied compressors and the differences in the compressors' performance. Nevertheless, several restrictions in the compressors' characterization and the cycle analysis were identified. This motivated us to deepen in the study of the two-stage compression cycle and its components. The next step was performing a theoretical analysis of two-stage compression cycles for heating applications, where the intermediate pressure and the injection ratio were identified as the most influential system parameters on the COP. The intermediate pressure was optimized for two vapor-injection configurations (flash tank and economizer) using several refrigerants. Based on the optimization results, a correlation was proposed that allows obtaining the optimal intermediate pressure of the cycle, considering the influence of the subcooling at the condenser outlet. In addition, a theoretical analysis of the influence of the design of the system components on the COP of the cycle was performed. Once the thermodynamic analysis of the two-stage cycle was carried out, the study was deepened at the component level. The most critical factor in the system is the compressor performance. Hence, the next step was evaluating the influence of several compression systems with vapor-injection on the COP. Three compressor technologies were taken into account, an SCVI, a TSRC and a two-stage scroll compressor (TSSC). These compressor technologies were characterized and modeled in order to study their performance. To do so, a new methodology to characterize SCVI was proposed. This methodology allows evaluating the compressor performance independently of the injection mechanism used in the cycle. A linear correlation was identified between the refrigerant injection ratio and the intermediate compression ratio. This correlation is used to determine the injection mass flow as a function of the intermediate pressure. Then, a semi-empirical model of scroll compressors and a methodology to extend the model for scroll compressors with vapor-injection was proposed. The models were adjusted and validated using experimental data from four scroll compressors working with R-290 and an SCVI compressor working with R-407C. Finally, an SCVI was compared with two two-stage compressors, a TSSC, and a TSRC, working in extreme conditions. The displacement ratio of the two-stage compressors was optimized. Results show that, at the nominal operating conditions (Te=-15 °C, Tc=50 °C), the optimal displacement ratio of the TSSC is 0.58, and of the TSRC is 0.57. The TSSC achieves 6% larger COP than the SCVI and 11.7% larger COP than the TSRC. Under a wide range of operating conditions, the SCVI presents a better efficiency and COP for pressure ratios below 5. For higher-pressure ratios, the TSSC presents better performance and achieves lower discharge temperature. It is concluded that the SCVI is an easy solution to implement from the point of view of machining, which allows extending the working map of the single-stage compressors. However, the results show that the two-stage compression technology gets further improve the COP of the cycle and the capacity, with a greater reduction of the discharge temperature operating under extreme conditions.
I thank the financial support provided by the Secretaría de Educación Superior, Ciencia, Tecnología e Innovación (SENESCYT) of Ecuador, through the international scholarship program for postgraduate studies “Convocatoria Abierta 2013 Segunda Fase, Grant No 2015-AR37665”.
Tello Oquendo, FM. (2019). Study of scroll compressors with vapor-injection for heat pumps operating in cold climates or in high-temperature water heating applications [Tesis doctoral no publicada]. Universitat Politècnica de València. https://doi.org/10.4995/Thesis/10251/120473
TESIS

The reliable operation of Integrally Geared Centrifugal Compressors (IGCCs), used in the coal-fired power generation industry of South Africa, is essential for economic, environmental and safety considerations. However, due to the unavailability of individual stage performance curves, the ability of a compressor owner to identify underperforming stages to maintain these compressors proactively remains limited. This study addresses the stage performance prediction of an IGCC when only the compressor’s overall performance characteristic, in conjunction with the impeller diameters and tip speeds, are known. The study is limited to IGCCs used in the coal-fired power generation industry of South Africa. Based on the limited inputs, two performance modelling methods were considered for this application, namely stage stacking and 1-dimensional modelling. However, stage stacking requires known operating points on each stage performance curve from which the rest of the curve can be extrapolated while 1-dimensional models require detailed stage design information to model stage performance. This study developed a revised stage stacking procedure which in contrast to the traditional stage stacking procedure, does not require a known operating point on each stage’s performance curve, for it assesses the relative stage performance at the compressor’s surge flow rate. The relative maximum pressure ratio of each stage is acquired through the application of similarity principles while a simplified 1-dimensional impeller analysis model is used to assess relative impeller head coefficients. The modelling process was developed based on performance and design data for IGCCs obtained from a compressor manufacturer. Performance data of four IGCCs, consisting of 13 stages, were obtained, including the design data for ten impellers. Hence, the IGCCs satisfy the requirements of geometric and aerodynamic similarity, unveiling a linear relationship between the stage impeller tip speed and maximum pressure ratio. A simplified 1-dimensional performance model was used to assess relative impeller head coefficients. A verification procedure ensured the integrity of the findings of the 1-dimensional model was maintained by comparing the model results to findings obtained using commercial compressor performance modelling software. A sensitivity analysis was conducted on the 1-dimensional performance model to ascertain which input parameters could be scaled as a function of the impeller tip diameter. For the four IGCCs for which data were obtained, the stage-discharge pressure and isentropic efficiency curves were calculated using the developed model. The maximum variation between the measured and calculated pressure and isentropic efficiency curves equaled 8.20% and 10.84%, respectively. The prediction accuracy of the developed modelling procedure is similar to map-based models found in literature and is considered adequate for identifying an underperforming stage. Thus, the developed model could serve as a valuable conditioning monitoring tool for site-based compressor owners.
Dissertation (MEng)--University of Pretoria, 2019.
Mechanical and Aeronautical Engineering
MEng
Unrestricted

Previous research on blade boundary layers in turbomachinery have been recognised to crucially influence the stability and performance of the gas turbine components. The interactions between rotating and stationary blade rows inevitably make the flow environment within a multistage axial compressor unsteady. Research conducted at midspan in Low Pressure turbines has shown that patches of transitional flow can withstand higher levels of deceleration, helping the boundary layer stay attached. An experimental investigation into unsteadiness in a embedded stage was conducted in the third-stage of the Cranfield four-stage Low Speed Research Compressor at two operating points: peak efficiency and near stall. This build of the Cranfield Rig was equipped with three-dimensional blading. A three-hole pressure probe was traversed at the exit of Rotor 3 in the rotating frame of reference and at the exit of Stator 3 in the stationary frame of reference. In addition measurements were made at the exit of both Rotor and Stator 3 using a slanted hotwire rotated about its axis. This measurement technique gave time-resolved three-dimensional velocities. Coupled to the exit traverses a series of boundary layer traverses were performed along Stator 3 suction surface covering the midchord region at midspan and close to the casing endwall. To aid in the understanding and interpretation of the experimental campaign, three-dimensional computations of Stator 3 were made at the two operating points using the commercial Computational Fluid Dynamics code ANSYS-CFX . A two-dimensional unsteady calculation of Rotor 3, Stator 3 and Rotator 4 at midspan and peak efficiency was also performed. The time-resolved measurements downstream of Rotor 3 showed that the rotor wake was characterised by high levels of random unsteadiness and increased incidence onto the stator row. The increase in incidence across the wake was two to three times that experienced with change in flow coefficient. Therefore the increased incidence and turbulence in a rotor wake will have a significant influence on the unsteady development of a downstream boundary layer. Measurements of the boundary layer at design condition at midspan show evidence of laminar and transitional flow up to 50% of the suction surface length. The boundary layer flow periodically undergoes transition due to the convection of the wake-induced strip that was generated close to the leading edge. Towards the casing the picture is altered slightly due to the stator-casing separation region. Boundary layer transition is completed farther forward and the transition length reduced. At off-design the picture is completely altered. Transition is completed upstream of 25% suction surface length and the flow shows only a modulating variation with blade passing. The stator-casing separation region grows in spanwise extent and the boundary layer flow on the stator surface is completely separated aft of 50% of suction surface length.

Thesis (M.S. in Aeronautical Engineering)--Naval Postgraduate School, Sept. 2001.
Thesis advisor, Shreeve, Raymond P. "September 2001." Includes bibliographical references (p. 117-119). Also available in print.

Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1989.
Includes bibliographical references (v.2, leaves 212-219).
by Petros Anestis Kotidis.
Ph.D.

Flow features inside centrifugal compressor stages are very complicated to simulate with numerical tools due to the highly complex geometry and varying gas conditions all across the machine. For this reason, a big effort is currently being made to increase the fidelity of the numerical models during the design and validation phases. Computational Fluid Dynamics (CFD) plays an increasing role in the assessment of the performance prediction of centrifugal compressor stages. Historically, CFD was considered reliable for performance prediction on a qualitatively level, whereas tests were necessary to predict compressors performance on a quantitatively basis. In fact "standard" CFD with only the flow-path and blades included into the computational domain is known to be weak in capturing efficiency level and operating range accurately due to the under-estimation of losses and the lack of secondary flows modeling. This research project aims to fill the gap in accuracy between "standard" CFD and tests data by including a high fidelity reproduction of the gas domain and the use of advanced numerical models and tools introduced in the author's OEM in-house CFD code. In other words, this thesis describes a methodology by which virtual tests can be conducted on single stages and multistage centrifugal compressors in a similar fashion to a typical rig test that guarantee end users to operate machines with a confidence level not achievable before. Furthermore, the new "high fidelity" approach allowed understanding flow phenomena not fully captured before, increasing aerodynamicists capability and confidence in designing high efficiency and high reliable centrifugal compressor stages.

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2003.
Includes bibliographical references (p. 75-76).
This electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections.
The impact of geometric variability on compressor performance is investigated using a compressor repeating-stage model based on well-known correlations for profile losses, endwall blockage, deviation, and the onset of stall. Previous computations with a quasi-two dimensional cascade analysis code are used to link geometric variability to performance deviations. Performance variability is then introduced probabilistically through random perturbations to tip clearances, profile losses and turning. For the variation input, at design incidence, the mean efficiency is found to decrease by 1%, mostly due to the mean shift in profile losses, and the mean pressure rise is reduced by 2.5%, mostly because of the mean shift in turning. A parametric study for compressor stages of different designs shows a lower degradation of mean performance and a lower performance variability for stages which have higher work coefficient, lower degree of reaction, and higher blade aspect ratio. It was found that the influence of blade profile effects was well represented, but the impact of tip clearance variation was not well captured when compared to three-dimensional computations. It is concluded that to address the effects of tip clearance variability, emphasis should be placed on development of models which both can include the alteration of end-wall displacement thickness within the compressor stage and are appropriate for probabilistic description.
by Antoine Vincent.
S.M.

A computer model of the J-85 gas turbine engine has been used in a investigation for potential benefits in performance simulation, arising from the adoption of compressor stage stacking models. The eight-stage, axial flow compressor of the engine was simulated by means of two separate stage stacking methods: the Howell program and the projects code, the latter being developed on the basis of the pre-existing J-85 complete engine simulation model. Results obtained from the use of the Howell program generally support its suitability for the prediction of single stage and overall compressor performance. This includes the capability of the code to render the effect from the Variation of specific design parameters and from the incorporation of variable geometry. Certain weaknesses originating from the empirical nature and the logic of the program are identified. Recommendations for the assessment of the validity of produced results, a well as modifications for the improvement of the code, are proposed. The projects stage stacking code was tested with sets of stage characteristics derived from different techniques. The produced results depend on the topology and shape of the utilised characteristics. A special feature embedded in the program's logic denotes the tendency of individual stages to work within their stall areas or below the hypothetical choke points of the corresponding characteristics. Although in certain instances the predicted overall compressor performance appears satisfactory, analysis of results in individual stages indicates the need for improvement of the code, in order to obtain a closer approach to the physical mechanism and the associated limitations of stage matching. This is also true for the prediction of variable geometry effects on the overall compressor performance. The complete engine simulation model incorporating the projects stage stacking code provides generally satisfactory results, a compared to those produced by a equivalent model of the Cranfield's Turbomatch Scheme. Specific problems encountered are attributed to certain weaknesses of the embedded stage stacking code, which is susceptible of improvement as already reported.

There is a demand for increasing efficiency of automotive engines, and one way to achieve this is through downsizing and turbocharging. In the design compromises are made, for example the maximum power of the engine determines the size of the compressor, but since the compressor mass flow range is limited, this affects the torque for low engine speeds. A two stage system, with two different sized turbochargers, reduces this compromise, but the system complexity increases. To handle the complexity, models have come to play a central role where they aid engineers in the design. Models are used in simulation, for design optimization and also in the control synthesis. In all applications it is vital that the models have good descriptive capabilities for the entire operating range studied. A novel control oriented compressor model is developed, with good performance in the operating regions relevant for compressors in a two stage system. In addition to the nominal operating regime, also surge, choke and operation at pressure ratios less than unity, are modeled. The model structure can be automatically parametrized using a compressor map, and is based on static functions for low computational cost. A sensitivity analysis, isolating the important characteristics that influence surge transients in an engine is performed, and the gains of a novel surge controller are quantified. A compressor map is usually measured in a gas stand, that has different surrounding systems, compared to the application where the compressor is used. A method to automatically determine a turbo map, when the turbo is installed on an engine in an engine test stand is developed. The map can then be used to parametrize the developed compressor model, and effectively create a model parametrized for its intended application. An experimental analysis of the applicability of the commonly used correction factors, used for estimating compressor performance when the inlet conditions deviate from nominal, is presented. Correction factors are vital, to e.g. estimate turbocharger performance for driving at high altitude or to analyze second stage compressor performance, where the variations in inlet conditions are large. The experimental campaign uses measurements from an engine test cell and from a gas stand, and shows a small, but clearly measurable trend, with decreasing compressor pressure ratio for decreasing compressor inlet pressure, for points with equal corrected shaft speed and corrected mass flow. A method is developed, enabling measurements to be analyzed with modified corrections. An adjusted shaft speed correction quantity is proposed, incorporating also the inlet pressure in the shaft speed correction. A high altitude example is used to quantify the influence of the modified correction.

The effects of impeller-diffuser interaction are investigated through numerically simulating the modified wedge vane profiles. Steady and time-accurate, 3D- viscous RANSsolver is used to perform flow field computations. The original design is modified to obtain better aerodynamic performance. Five morechanges are made to the leading edge profile of the new design, in order to assess different degrees of unsteadiness. These changes show that their contribution on stageefficiency is rather minor, while they have a huge reduction on blade loadings. Moreover, it is shown that the shorter radial distance of vaneless space does not necessarilymean an increased loading thanks to the eliminating in-phase fluctuations on pressureand suction sides. It is found that the impeller reacts to the upstream static pressure disturbance, whichis caused by the applied geometry change and its resultant flow field in the wedge diffuser, but not to the radial location of a certain profile. In addition, the results indicatethat the wedge diffuser aerodynamic performance is driven by time-averaged flow fieldbehaviour.

The purpose of this study is to analyze the behavior of a reversible two-cylinder refrigerant compressor manufactured by Bristol Compressor Incorporated. This compressor contains a specialized linkage that causes the compressor to transition from a two-cylinder compressor to a single-cylinder compressor when the direction of rotation of the crankshaft is reversed. The linkage accomplishes this by reducing the throw of one cylinder to zero. Of interest are the conditions to which this linkage is subjected when the direction of rotation is again reversed, causing the compressor to return to its two-cylinder functioning. When this reversal takes place, a collision occurs within the linkage. These repeated collisions are thought to be the cause of fatigue failure of the linkage in many of these compressors. To verify that this collision is the problem, an understanding of the stress state during the collision is needed. This thesis begins the work necessary to determine the dynamic stress state present within the system. A FORTRAN program was developed that modeled the kinematic behavior of the system under operating conditions. The program predicts the accelerations, velocities, positions, and internal forces present within the system during startup conditions. Also, a method has been developed to model rotary sliding contact between two cylindrical surfaces. This method is developed and investigated in hopes that it will facilitate the modeling of the behavior of the compressor linkage in a dynamic finite element analysis.
Master of Science

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2006.
Includes bibliographical references (p. 137).
Numerical experiments have been implemented to characterize the unsteady loading on the rotating impeller blades in a modem centrifugal compressor. These consist of unsteady Reynolds-averaged Navier Stokes simulations of three-dimensional and quasi-two dimensional approximate models. The interaction between the rotating impeller and the stationary downstream diffuser has been identified as strong source of unsteady loading on the impeller blades. First of a kind unsteady calculations haven been carried out to elucidate an upstream manifestation of a downstream stimulus experienced in a particular centrifugal compressor stage. Here the upstream manifestation is the considerable unsteady loading in the splitter blade leading edge while the downstream stimulus is the unsteady impeller-diffuser interaction Three key parameters that control the level and extent of the unsteady loading are the impeller-diffuser gap, stage loading, and the impeller passage relative Mach number. Impeller-diffuser gap has been shown to control the peak level of unsteady loading on the blade. Stage loading has been shown to control the upstream attenuation of the loading.
(cont.) A hypothesis has been put forward that increased diffusion associated with increased stage loading increases the impeller sensitivity to the downstream disturbance. The relative Mach number has been shown to set the chordwise distribution of the unsteady load on the blade. Unsteady blade loading has been computed through a quasi two-dimensional model in which an unsteady pressure boundary condition is imposed at the impeller exit to approximate the presence of the downstream diffuser. Results of this approximate model have been shown to yield unsteady loading characteristics that are in accord with the full three-dimensional unsteady model. An implied utility of this result is that a quasi-2D approximation could be used during the design phase to approximate the unsteady loading in a timeframe that is compatible with the design environment. The effect of unsteady flow on mass flow capacity of a fluid device is eliminated as a source for over-predictions in mass flow when a steady-state approximation is used.
by Kenneth A. Gould.
S.M.

Thesis (Ph.D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1999.
Includes bibliographical references (p. 145-150).
The pressure ratio of axial compressor stages can be significantly increased by controlling the development of blade and endwall boundary layers in regions of adverse pressure gradient by means of boundary layer suction. This concept is validated and demonstrated through the design and analysis of two unique aspirated compressor stages: a low-speed stage with a design pressure ratio of 1.6 at a tip speed of 750 ft/s, and a high-speed stage with a design pressure ratio of 3.5 at a tip speed of 1500 ft/s. The aspirated compressor stages were designed using a new procedure which is a synthesis of low speed and high speed blade design techniques combined with a flexible inverse design method which enabled precise independent control over the shape of the blade suction and pressure surfaces. Integration of the boundary layer suction calculation into the overall design process is an essential ingredient of the new procedure. The blade design system consists of two axisymmetric through-flow codes coupled with a quasi three-dimensional viscous cascade plane code with inverse design capability. Validation of the completed designs were carried out with three-dimensional Euler and Navier-Stokes calculations. A single spanwise slot on the blade suction surface is used to bleed the boundary layer. The suction mass flow requirement for the low-speed and high-speed stages are 1 % and 4% of the inlet mass flow, respectively. Additional suction between 1-2% is also required on the compressor end walls near shock impingement locations. The rotor is modeled with a tip shroud to eliminate tip clearance effects and to discharge the suction flow radially from the flowpath. Three-dimensional viscous evaluation of the designs showed good agreement with the quasi three-dimensional design intent, except in the endwall regions. The suction requirements predicted by the quasi three-dimensional calculation were confirmed by the three-dimensional viscous calculations. The three-dimensional viscous analysis predicted a peak pressure ratio of 1.59 at an isentropic efficiency of 89% for the low-speed stage, and a peak pressure ratio of 3.68 at an isentropic efficiency of 94% for the high-speed rotor.
by Ali M. Merchant.
Ph.D.

Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2017.
Cataloged from PDF version of thesis.
Includes bibliographical references (pages 269-270).
This research characterizes the performance of a centrifugal compressor stage with a special focus on the pipe diffuser. Two diffuser configurations are studied, one of which is a truncated version of the other. Experimental data acquired on a research compressor stage is interrogated along with a set of well-designed Reynolds-Averaged Navier Stokes computations, complemented by reduced order flow modeling. The fundamental performance-limiting flow mechanisms in the diffuser are identified and used to physically relate important geometry features and operating conditions to the observed compressor pressure rise, efficiency, and operability characteristics. Despite large differences in their geometry, the two diffuser configurations exhibit similar pressure recovery characteristics due to differences in exit nonuniformity and flow angle which result in similar effective area ratios. Variations in the diffuser pressure recovery coefficient with operating point are found to be most influenced by the diffuser inlet flow angle, and secondly by the inlet Mach number. The diffuser inlet flow angle has the primary effect of setting the diffuser inlet one-dimensional area ratio, increasing diffusion at high flow angles. In addition, the diffuser incidence angle influences the formation of counter-rotating vortex pairs that persist throughout the diffuser passage. Using a two-dimensional integral boundary layer model that is modified to accommodate three-dimensional effects as source terms, these secondary flows are shown to detrimentally impact the diffuser pressure rise capability by accumulating high loss flow along the diffuser wall near the plane of symmetry between the vortices. This contributes to the extent and location of a large diffuser passage separation, especially for the baseline diffuser. The impact of the vortices on the boundary layer growth rate is shown to scale inversely with diffuser aspect ratio. The major performance difference between the two diffuser configurations is that the truncated diffuser configuration experiences enhanced stall margin over the baseline diffuser at the design speed. These differences are traced to reduced secondary flows influence and thus reduced separation extent for the higher aspect ratio truncated diffuser. It is hypothesized that the onset of stall for the baseline diffuser configuration is initiated by the transition of the vortex location and corresponding passage separation between diffuser pressure and suction sides with increasing cusp incidence. Conversely, because the extent of the passage separation in the truncated diffuser is diminished due to the higher aspect ratio, the switch in separation side does not immediately initiate instability. The fact that secondary flows have a large influence on diffuser pressure rise capability and compressor stability is counter to conventional preliminary diffuser design approaches which neglect such 3D effects. The findings of this research may therefore be considered during preliminary design optimization to produce better-performing diffuser designs.
by David W. Erickson.
S.M.

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2006.
Includes bibliographical references (p. 96-97).
A major challenge in the design of multi-stage compressors is the matching of stages to enable stable operation over a large range of mass flows and operating conditions. Particularly in turbofan low-pressure compressors, where a variable geometry cannot be implemented, design strategies for maximum efficiency at high speed can compromise the surge margin at low speed. In this thesis, a design optimization framework has been implemented to an industry-strength compressor-matching problem. The optimization framework combines a mean-line flow solver and a dynamic stability analysis of a six-stage low-pressure compressor of a modern turbofan engine to optimize the blade row geometry for enhanced stability at flight idle conditions. To assess the potential improvements in compressor stability at low speed, a number of optimization strategies are employed using different objective functions and stability metrics. To estimate the performance and stability of the six-stage compressor, a mean-line flow solver is developed and coupled with a previously developed dynamic compressor-stability analysis. A fan-root flow model and an endwall loss correlation are developed using performance data provided by industry.
(cont.) The analysis reveals that the models enable an adequate estimation of the datum compressor performance. This methodology is then used in an optimization effort searching for the optimum compressor design. A compressor blade parametrization based on Bezier splines is developed to explore a range of possible blade geometries. A CFD-based blade-row performance database is established using the blade-to-blade solver MISES. This facilitates an effective means to predict the blade performance for various geometries defined by the optimizer. To find the best solution for the compressor-matching problem, a number of optimization strategies are applied to the datum compressor. The best result is obtained using an optimization strategy based on industry surge margin. An improvement of 14.8% in flight idle surge margin is achieved while maintaining the design pressure ratio and efficiency at climb speed within 1% and 0.3 points of the design values respectively. A compressor design optimization based on a dynamic-stability metric is also employed. Due to time constraints, this strategy could not be fully explored and the preliminary results suggest that further work is required.
(cont.) The best results is a 14.8% improvement in the flight idle surge margin, but the re-matching of the compressor and the associated increase in the rotor loading of the second stage entail high-risk design modifications. This suggests that, given these design limitations, the best matching is achieved by the datum configuration. In summary, the thesis demonstrates that the developed compressor design optimization methodology is applicable to industry-strength design problems, and the framework is shown to have the potential to investigate compressor designs for optimum matching.
by Jérôme Bert.
S.M.

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Mechanical Engineering, 2012.
Cataloged from PDF version of thesis.
Includes bibliographical references (p. 111-114).
Compressor efficiency variation with rotor tip gap is assessed using numerical simulations on an embedded stage representative of that in a large industrial gas turbine with Reynolds number being approximately 2 x 106 to 7 x 106. The results reveal three distinct behaviors of efficiency variation with tip gap. For relatively small tip gap (less than 0.8% span), the change in efficiency with tip gap is non-monotonic with an optimum tip gap for maximum efficiency. The optimum tip gap is set by two competing flow processes: decreasing tip leakage mixing loss and increasing viscous shear loss at the casing with decreasing tip gap. An optimum tip gap scaling is established and shown to satisfactorily quantify the optimal gap value. For medium tip gap (0.8% - 3.4% span), the efficiency decreases approximately on a linear basis with increasing tip clearance. However, for tip gap beyond a threshold value (3.4% span for this rotor), the efficiency becomes less sensitive to tip gap as the blade tip becomes more aft-loaded thus reducing tip flow mixing loss in the rotor passage. The threshold value is set by the competing effects between increasing tip leakage flow and decreasing tip flow induced mixing loss with increasing tip gap. Thus, to desensitize compressor performance variation with blade gap, rotor should be tip aft-loaded and hub fore-loaded while stator should be tip fore-loaded and hub aft-loaded as much as feasible. This reduces the opportunity for clearance flow mixing loss and maximizes the benefits of reversible work from unsteady effects in attenuating the clearance flow through the downstream blade-row. The net effect can be an overall compressor performance enhancement in terms of efficiency, pressure rise capability, robustness to end gap variation and potentially useful operable range broadening. Preliminary assessment of a stage redesign with a 4% chord more tip aft-loaded blade design for 1.7 % span tip clearance yields 0.2 point stage efficiency benefit.
by Sitanun Sakulkaew.
S.M.

Thesis: S.M., Massachusetts Institute of Technology, Department of Mechanical Engineering, 2015.
Cataloged from PDF version of thesis.
Includes bibliographical references (pages 117-119).
Changes in loss generation associated with altering the rotor tip loading of an embedded compressor stage is assessed. Steady and unsteady three-dimensional computations, complemented by control volume analyses, for varying rotor tip loading distributions provided results for determining if aft-loading rotor tip would yield a stage performance benefit in terms of a reduction in loss generation. Aft-loading rotor blade tip yields a relatively less-mixed-out tip leakage flow at the rotor exit and a reduction in overall tip leakage mass flow hence a lower loss generation; however, the attendant changes in tip flow angle distribution are such that there is an overall increase in the flow angle mismatch between tip flow and main flow leading to higher loss generation. The latter outweighs the former so that rotor passage loss from aft-loading rotor tip is marginally higher unless a constraint is imposed on tip flow angle distribution so that associated induced loss is negligible; a potential strategy for achieving this is proposed. In the course of assessing the benefit from unsteady tip leakage flow recovery in the downstream stator, it was determined that tip clearance flow is inherently unsteady with a time-scale distinctly different from the blade passing time. The disparity between the two timescales: (i) defines the periodicity of the unsteady rotor-stator flow, which is an integral multiple of blade passing time; and (ii) causes tip leakage vortex to enter the downstream stator at specific pitchwise locations for different blade passing cycles, which is a tip leakage flow phasing effect. Because of an inadequate grid resolution defining the unsteady interaction of tip flow with downstream stator, the benefit from unsteady tip flow recovery is the lower bound of its actual benefit. A revised design hypothesis is thus as follows: "rotor should be tip-aft-loaded and hub-fore-loaded while stator should be hub-aft-loaded and tip-fore-loaded with tip/hub leakage flow angle distribution such that it results in no additional loss". For the compressor stage being assessed here, an estimated 0.15% enhancement in stage efficiency is possible from aft-loading rotor tip only.
by Aniwat Tiralap.
S.M.

This thesis investigates over tip leakage where the tip clearance is large. In the high pressure stages of axial compressors the tip clearance can be typically 6% of span and the total blockage due to tip clearance can consume in excess of forty percent of the annulus height. Experimental and computational investigations of large tip clearance in a linear cascade have been used to investigate this phenomenon. Two cascade builds have been used the first (Build A) consisted of a controlled diffusion aerofoil of low stagger and thirty degrees flow turning. The second cascade (Build B) consisted of an engine representative design with high stagger and around ten degrees of flow turning. The diffusion factor of both cascades was around 0.3. The major findings are that: Large tip clearances have a smaller detrimental influence on single row performance than the previous research would have suggested, for Build B the loss at 10% tip clearance was the same as the 0% tip clearance loss, though the overall flow turning was much reduced. An increase in blade loading towards the tip was observed with both builds. Both these phenomenon were attributed to the small amount of movement of the over tip leakage vortex. An engine representative level of inlet skew was implemented using upstream injection so to assess its influence. This was found to have a remarkably small influence on the performance of a single row with the tip clearance and geometry of the blading having a much greater influence. Finally a circumferential grooved casing treatment was applied in the linear cascade but this was found not to be an appropriate tool for such an investigation.

Tens of thousands of centrifugal compressors are installed worldwide in chemical and petrochemical plants. The performance of these compressors degrade during the operation as a consequence of effects such as fouling, erosion, corrosion, and abrasion. A performance monitoring method that could detect and assess the magnitude of the degradation would be greatly beneficial to schedule production and maintenance leading to an economic profit for the operation. After searching the literature, it was concluded that such a method is yet to be developed for industrial centrifugal compressors. This thesis shows the development of an adaptive monitoring framework that simultaneously considers the degradation of the state of health and identifies the malfunctioning of sensors. Indeed, because of degradation, the state of the components changes over time and this change can be observed in the measurements, however, the measurements are also affected by random or persistent errors. The approach adopted in this thesis aims at reconciling two distinct features that are normally separated: (1) to account for the degraded state by recursively matching the model to the newly available measurements, and (2) to correct the measurements when these are biased by making use of the model. The leading hypothesis is that this aim can be achieved by employing models that establish causality relationships between the state of the components and the measured variables together with mathematical optimization methods. The thesis demonstrates: how performance can be systematically modelled even though a compressor is installed at the industrial site, how the degradation of performance can be detected and quantified in real-time, and finally, how the effects of degradation on performance can be modelled and monitored while simultaneously detecting and correcting sensor faults. The methods have been successfully applied to a 10 MW centrifugal compressor. When monitoring and modelling the degradation of its performance, it was observed that the difference between the performance of the compressor in undegraded and degraded state depends on the operation conditions. The implication of this observation is that the state-of-the-art practice of scaling the manufacturer maps to obtain degraded maps can lead to misleading conclusions about the performance of the degraded compressor.

A numerical case study of a multistage compressor with relative supersonic rotors is presented. The purpose of the investigation was to determine the flow instability mechanism of the UEET compressor and its relation to the rotor shock structure in the relative velocity reference frame. The computational study was conducted with the NASA code ADPAC , utilizing the mixing-plane assumption for the boundary condition between adjacent, relatively-rotating blade rows. A steady, five-blade-row, numerical simulation using the Baldwin-Lomax turbulence model was performed, creating several constant speed lines. The results are presented, highlighting the role shock structure plays in the stability of the compressor. The shock structure in the downstream rotor isolates the upstream rotor from the exit conditions until the shock detaches from the leading edge. At this point the shock structure in the upstream rotor moves, changing the conditions for the downstream rotor. This continues with increasing pressure at the exit until the shock in the upstream rotor detaches from the leading edge. This event causes an instantaneous drop in the mass flow rate, initiating positive incident separation on the suction side of stator-two.
Master of Science

Thesis (M.S. in Mechanical Engineering)--Naval Postgraduate School, June 2008.
Thesis Advisor(s): Gannon, Anthony J. "June 2008." Description based on title screen as viewed on August 25, 2008. Includes bibliographical references (p. 29-30). Also available in print.

Thesis (Ph. D.) -- University of Maryland, College Park, 2008.
Thesis research directed by: Dept. of Mechanical Engineering. Title from t.p. of PDF. Includes bibliographical references. Published by UMI Dissertation Services, Ann Arbor, Mich. Also available in paper.

Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2014.
Cataloged from PDF version of thesis.
Includes bibliographical references (pages 135-136).
The unsteady pressure field experienced by a centrifugal compressor stage can be dominated by of the impeller-diffuser interaction. The energy of the unsteady field, under certain aerodynamic and structural conditions, is capable of forcing the rotating impeller blades to vibrate excessively to the point of failure, better known as a high cycle fatigue (HCF) failure. This thesis seeks to identify the physical mechanisms that set the forced response amplitude of an impeller due to the impeller-diffuser interaction. The centrifugal stage researched is comprised of a stationary discrete passage diffuser and an unshrouded rotating impeller with both main and splitter blades. The forced response of two splitter blade modes are computed for a variety of structural boundary conditions and unsteady loadings to elicit the driving physical mechanisms. The findings indicate that the forced response is enhanced when the excitation frequency matches a component's natural frequency, the characteristic wavelength of the unsteady loading matches that of the structural vibration mode, the resonance occurs at high speed, and when modal displacement exists at the impeller blade's trailing edge. The findings also suggest that modal coupling of blade and disk dominant modes leads to high sensitivity of the forced response to small variations in airfoil and disk backwall thickness. Identification of blade-disk couplings are described using a simplified SAFE (Singh's Advanced Frequency Evaluation) diagram. The forced response of taut strings, Bernoulli-Euler beams, and a two mass-spring system are also utilized to elicit how the physical mechanisms act on the impeller's forced response. The Bernoulli-Euler beam model suggests that a mismatch of the forcing wavelength to the structural wavelength by 50% will reduce the forced response amplitude by at least 75%. Finally, a decision tree is proposed to assess the relative resonant risk of impeller modes to the diffuser excitation by identifying which of the physical mechanisms may be the dominant driver of the forced response.
by Edward James Walton.
S.M.