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1
Piscopo, Giovanni. "Preliminary aerothermal design of axial compressors." Thesis, Cranfield University, 2013. http://dspace.lib.cranfield.ac.uk/handle/1826/7909.
Full textAbstract:
The research work disclosed in this publication is partially funded by the
Strategic Educational Pathways Scholarship Scheme (Malta). The
scholarship is part-financed by the European Union – European Social
Fund under Programme II – Cohesion Policy 2007-2013.This dissertation documents a compressor preliminary design study conducted by the author in fulfilment of his MSc thesis requirements. The compressor is intended for a new development engine within the 20Klbf thrust category, planned to be used on a short-haul aircraft, namely the ERJ-190. A market research suggests that there exists a definite opportunity for a commercially profitable engine within this thrust class. Furthermore, the proposed new engine is projected to outperform current production engines on critical issues such as fuel efficiency and operability. By and large, the objectives of this work have been achieved and a compressor design and layout is suggested, which matched or exceeded all the initial requirements. The quality of the results from this study are thought to be of sufficient detail to allow a further, more detailed development study to resolve some subtle pending issues. It is expected that, some compressor stages may have to be altered slightly during detailed design to augment their performance and ease of manufacture and assembly. Throughout this study, the importance of the compressor design figure of merits, pertaining to a short haul engine, has been outlined and their interaction on the design process is well documented. Furthermore, some rather unorthodox objectives such as compressor performance retention and reliability have been discussed. The author approached these subjects in an innovative way due to the limited non-proprietary knowledge available on these issues, especially considering their implications within preliminary design. Furthermore, the author developed and tested a new preliminary turbomachinery design code, named Turbodev, which can be used as an aid in future compressor design endeveours. Turbodev can handle most types of compressor layouts and generates an overall aerodynamic assessment of the turbomachinery performance. In conclusion; this documentation and the associated literature review aim to provide the reader with an overview of the work done and yield a better understanding of the decisions that face any design bureau when developing a new or modified engine component.
2
Lopes, Fernando de Oliveira. "Modelo computacional para projeto de compressores axiais." [s.n.], 2007. http://repositorio.unicamp.br/jspui/handle/REPOSIP/265338.
Full textAbstract:
Orientador: Jorge Isaias Llagostera BeltranDissertação (mestrado) - Universidade Estadual de Campinas, Faculdade de Engenharia Mecanica
Made available in DSpace on 2018-08-08T17:53:07Z (GMT). No. of bitstreams: 1 Lopes_FernandodeOliveira_M.pdf: 4254683 bytes, checksum: 3714a253c99b8967319409a00b69058a (MD5) Previous issue date: 2007
Resumo: Este trabalho apresenta o desenvolvimento de um programa computacional para modelagem inicial de compressores axiais de vários estágios pertencentes ao conjunto de turbinas a gás. O desenvolvimento do programa se baseia na metodologia adotada por Saravanamutto et al. (2001), faz uso da Primeira Lei da Termodinâmica para cálculo de potência consumida pelo compressor e da Segunda Lei da Termodinâmica para determinar o grau de irreversibilidade do sistema. O programa calcula a quantidade de estágios necessária para uma dada relação de pressão, a quantidade de palhetas por estágio e outros dados construtivos do compressor. O trabalho analisa a eficiência global de uma turbina a gás, avalia rendimento utilizando diferentes tipos combustíveis, estuda a influência da temperatura de entrada do ar no compressor, temperatura de entrada dos gases na turbina, e eficiência isentrópica do compressor e da turbina. Fatores que geram instabilidade no compressor são discutidos e algumas sugestões são apresentadas para evitar que compressores operem fora das condições iniciais. O trabalho apresenta procedimentos claros e detalhados para o préprojeto de um compressor de fluxo axial. Finalmente, o trabalho apresenta uma breve discussão sobre eficiência exergética de máquinas térmicas
Abstract: This work presents the developing of a computational program for designing axial compressors that hold multistage belonged gas turbine. The developing of the program is based on methodology adopted by Saravanamutto et al.(2001), it makes use of the First Law of Thermodynamic to calculate the power required by the axial compressor e the Second Law to calculate the level of irreversibilities. Beside of this the program presents the numbers of stages required for a given pressure ratio, the amount of blades per stage and other building parameters has been included to make a better analyze about the equipment. The work contains thermal efficiency analyzes of a gas turbines, where parameters such as fuels, temperature intlet turbine, environmental conditions, efficiency of the compressor and turbine are included. Other factors such as unstable conditions are discussed and solutions to avoid that axial compressors running in off design conditions. In summary the work provides a global view about thermal machines and how their parameters can influence both in the thermal and exergetic efficiency
Mestrado
Termica e Fluidos
Mestre em Engenharia Mecânica
3
Nucara, Pascal. "Design of gas turbine axial compressors for fuel flexibility." Thesis, University of Sussex, 2014. http://sro.sussex.ac.uk/id/eprint/48905/.
Full textAbstract:
Current gas turbine technology for power generation is generally optimised for natural gas. On the basis of current instabilities in natural gas price and supply, the use of alternative fuels, such as syngas, has recently gained high interest. Due to the different thermodynamic properties of syngas compared to natural gas the behaviour of existing gas turbine components may significantly change. From practical and economic points of view, it is generally considered that in order to meet the new fuel properties, the main effort should be put on the adaptation of conventional gas turbines in integrated gasification combined cycle (IGCC) plants rather than producing a new generation of gas turbine designs from scratch. In addition to the requirement of new combustion technologies, main critical issues are represented by the reduction of compressor surge margin and turbine blade overheating. Solutions might include thermodynamic cycle as well as turbine geometry modifications. The latter would be preferred in terms of power plant performance. The main aim of this thesis is to explore suitable solutions to be applied to gas turbine compressors in order to accommodate syngas combustion. Among others, the use of variable stator vanes (VSVs) and blade radial stacking line modifications are considered. These are investigated on reference geometries available in the public domain. A baseline compressor geometry representative of a conventional heavy-duty gas turbine fueled with natural gas is generated and modified according to the understating gained during this study. The re-designed machine is a result of the application of stator vanes re-staggering in the front stages and blade sweep in the rear stages in order to cope with compressor air supply control and critical flow separation regions respectively. The obtained results show that efficient and stable operation during power modulation can be achieved, while reducing the need of other modifications to the combined cycle plant. It was therefore concluded that the proposed option can be considered a viable option to satisfy some important technical and economic constraints imposed by the integration of an existing gas turbine within an IGCC plant.
4
Camp, Timothy Richard. "Aspects of the off-design performance of axial flow compressors." Thesis, University of Cambridge, 1995. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.387517.
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Robinson, Christopher J. "End-wall flows and blading design for axial flow compressors." Thesis, Cranfield University, 1991. http://dspace.lib.cranfield.ac.uk/handle/1826/6929.
Full textAbstract:
The flow in multistage axial flow compressors is
particularly complex in nature because of the proximity of
moving bladerows, the growth of end-wall boundary layers and
the presence of tip and seal leakages and secondary flow.
The problems associated with these phenomena are at their
most acute in the latter, subsonic stages of the core
compressor, where Reynolds numbers are modest and the
blading has low aspect ratio. Indeed, much of the
inefficiency of axial stages is believed to be associated
with the interaction between blading and end-wall flows.
The fact that the end-wall flow phenomena result in
conditions local to the blade which are quite different
from those over the major part of the annulus was
appreciated by many of the earliest workers in the axial
turbomachinery field. However, experiments on blading
designs aimed specifically at attacking the end-loss have
been sparse.
This thesis includes results from tests of conventional
and end-bent blading in a four-stage, low-speed, axial
compressor, built specifically for the task, at a scale
where high spatial measurement resolution could be readily
achieved within the flowpath. Two basic design styles are
considered: a zero a0 stage with DCA aerofoils and a
low-reaction controlled-diffusion design with cantilevered
stators.
The data gives insight into the flow phenomena present in
'buried' stages and has resulted in a much clearer
understanding of the behaviour of end-bent blading. A
3D Navier-Stokes solver was calibrated on the two
low-reaction stators and was found to give good agreement
with most aspects of the experimental results. An improved
design procedure is suggested based on the incorporation
of end-bends into the throughflow and iterative use of the
3D Navier-Stokes solver.
6
Ando, Victor Fujii. "Genetic algorithm for preliminary design optimisation of high-performance axial-flow compressors." Instituto Tecnológico de Aeronáutica, 2011. http://www.bd.bibl.ita.br/tde_busca/arquivo.php?codArquivo=1969.
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This work presents an approach to optimise the preliminary design of high-performance axial-flow compressors. The preliminary design within the Gas Turbine Group at ITA, is carried on with an in-house computational program based upon the streamline curvature method, using correlations from the literature to assess the losses. The choice of many parameters of the thermodynamic cycle and of geometries relies upon the expertise from the members of the Group. Nevertheless, it is still a laborious and time-consuming task, requiring successive trial and errors. Therefore, to support the compressor designer in the choice of some parameters, an optimisation program, named REMOGA, was written in FORTRAN language, allowing an easy integration with the programs developed by the Gas Turbine Group. The program is based upon a multi-objective genetic algorithm, with real codification and elitism. Then the REMOGA and the preliminary design program were integrated to design a 5-stage axial-flow compressor. Therefore, the stator air outlet angles, the temperature distribution and the hub-tip ratio were varied aiming at higher efficiencies and higher pressure ratios, but controlling the de Haller number and the camber angle. Thanks to the REMOGA, thousands of designs could be quickly evaluated. Finally, using a choice criterion, four solutions were selected for further analysis, revealing that the developed program was successful in finding more efficient and feasible compressor designs.
7
Birkenheier, David Andrew. "Non-uniform radial meanline method for off-design performance estimation of multistage axial compressors." Thesis, Massachusetts Institute of Technology, 2018. http://hdl.handle.net/1721.1/119062.
Full textAbstract:
Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2018.Cataloged from PDF version of thesis.
Includes bibliographical references (pages 111-112).
The increasing use of renewable energy sources necessitates power-generating gas turbines capable of frequently and rapidly starting up to supplement the energy supply when renewable sources alone cannot meet demand [1], [21. This makes the off-design performance of such gas turbines more important as they spend more of their operational life off the design point. Currently off-design performance cannot be estimated with high fidelity until late in the gas turbine compressor design process at which point the design is largely fixed and only limited changes can be made. This thesis presents a Non-Uniform Radial Meanline method for multistage axial compressor off-design performance estimation, capturing the transfer of radial flow non-uniformity and its impact on compressor blade row performance. This method enables the high-fidelity characterization of blade row performance and the stage matching of multistage compressors with non-uniformity effects included. A new representation of non-uniform radial flow profiles using orthonormal basis functions was developed to provide a compact representation suitable for inclusion in a one-dimensional performance estimation method. The link between radial flow non-uniformity and compressor blade row performance was characterized using three-dimensional embedded stage calculations. An initial implementation of the Non-Uniform Radial Meanline method was demonstrated for different compressor inlet non-uniformities. The computations show that the new approach provides an effective means of incorporating radial flow non-uniformity into a one-dimensional compressor performance estimation method.
by David Andrew Birkenheier.
S.M.
8
Iyengar, Vishwas. "A First Principles Based Methodology for Design of Axial Compressor Configurations." Diss., Georgia Institute of Technology, 2007. http://hdl.handle.net/1853/16163.
Full textAbstract:
Axial compressors are widely used in many aerodynamic applications. The design of an axial compressor configuration presents many challenges. Until recently, compressor design was done using 2-D viscous flow analyses that solve the flow field around cascades or in meridional planes or 3-D inviscid analyses. With the advent of modern computational methods it is now possible to analyze the 3-D viscous flow and accurately predict the performance of 3-D multistage compressors. It is necessary to retool the design methodologies to take advantage of the improved accuracy and physical fidelity of these advanced methods.
In this study, a first-principles based multi-objective technique for designing single stage compressors is described. The study accounts for stage aerodynamic characteristics, rotor-stator interactions and blade elastic deformations. A parametric representation of compressor blades that include leading and trailing edge camber line angles, thickness and camber distributions was used in this study A design of experiment approach is used to reduce the large combinations of design variables into a smaller subset. A response surface method is used to approximately map the output variables as a function of design variables. An optimized configuration is determined as the extremum of all extrema.
This method has been applied to a rotor-stator stage similar to NASA Stage 35. The study has two parts: a preliminary study where a limited number of design variables were used to give an understanding of the important design variables for subsequent use, and a comprehensive application of the methodology where a larger, more complete set of design variables are used. The extended methodology also attempts to minimize the acoustic fluctuations at the rotor-stator interface by considering a rotor-wake influence coefficient (RWIC). Results presented include performance map calculations at design and off-design speed along with a detailed visualization of the flow field at design and off-design conditions.
The present methodology provides a way to systematically screening through the plethora of design variables. By selecting the most influential design parameters and by optimizing the blade leading edge and trailing edge mean camber line angles, phenomenon s such as tip blockages, blade-to-blade shock structures and other loss mechanisms can be weakened or alleviated. It is found that these changes to the configuration can have a beneficial effect on total pressure ratio and stage adiabatic efficiency, thereby improving the performance of the axial compression system. Aeroacoustic benefits were found by minimizing the noise generating mechanisms associated with rotor wake-stator interactions. The new method presented is reliable, low time cost, and easily applicable to industry daily design optimization of turbomachinery blades.
9
Ramakdawala, Rizwan R. "Preliminary design code for an axial stage compressor." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2001. http://handle.dtic.mil/100.2/ADA397395.
Full textAbstract:
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.
10
Merchant, Ali A. (Ali Abbas). "Design and analysis of axial aspirated compressor stages." Thesis, Massachusetts Institute of Technology, 1999. http://hdl.handle.net/1721.1/9362.
Full textAbstract:
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.
11
Wisnia, Arnaud Irving. "A computational model for multistage axial compressor design." Thesis, Massachusetts Institute of Technology, 1998. http://hdl.handle.net/1721.1/50318.
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Bendali-Amor, M. "Secondary and endwall losses in an axial flow compressor." Thesis, University of Newcastle Upon Tyne, 1991. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.316159.
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Reid, William D. "Transonic axial compressor design case study and preparations for testing." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 1995. http://handle.dtic.mil/100.2/ADA306259.
Full textAbstract:
Thesis (M.S. in Aeronautical Engineering) Naval Postgraduate School, September 1995.Thesis advisor(s): Raymond P. Shreeve. "September 1995." Includes bibliographical references. Also available online.
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Fréchette, Luc G. (Luc Guy). "Implications of stability modeling for high-speed axial compressor design." Thesis, Massachusetts Institute of Technology, 1997. http://hdl.handle.net/1721.1/10721.
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Dorca, Luque Josep M. (Josep Maria) 1981. "Application of an energy-like stability metric for axial compressor design." Thesis, Massachusetts Institute of Technology, 2003. http://hdl.handle.net/1721.1/82800.
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Brand, Maximilian Lewis. "An improved blade passage model for estimating off-design axial compressor performance." Thesis, Massachusetts Institute of Technology, 2013. http://hdl.handle.net/1721.1/85765.
Full textAbstract:
Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2013.Cataloged from PDF version of thesis.
Includes bibliographical references (pages 127-129).
Accurate estimates of multistage axial compressor performance at off-design operating conditions are essential to the determination of key performance metrics of aircraft gas turbine engines, such as fuel burn, thrust output, and stable operating range. However, conventional RANS based CFD calculations of multistage axial compressors diverge at off-design operating conditions where large separation occurs and the stages are mismatched. This thesis demonstrates the feasibility of a body force based approach to capturing the three-dimensional flow field through a turbomachinery blade row at off-design conditions. A first principles based blade passage model is introduced which addresses the limitations of previous approaches. The inputs to the improved blade passage model are determined from three-dimensional, steady, single-passage RANS CFD calculations. In a first step towards modeling multistage configurations, the improved blade passage model is validated using a fan rotor test case. At the design operating conditions, the stagnation pressure rise coefficient and the work coefficient are both estimated within 5%, and the adiabatic efficiency is estimated within 1 percentage point over most of the span relative to single-passage RANS CFD simulations. At low mass flow operating conditions, where the single-passage RANS CFD diverges, the blade passage model and related body force representation are capable of computing the three-dimensional throughflow with separation and reversed flow. These results pave the way for future unsteady calculations to assess compressor stability and for multistage compressor simulations at off-design conditions.
by Maximilian Lewis Brand.
S.M.
17
Blanvillain, Emmanuel 1979. "Dynamic stability analysis of a multi-stage axial compressor with design implications." Thesis, Massachusetts Institute of Technology, 2003. http://hdl.handle.net/1721.1/82255.
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Mahmood, Syed Moez Hussain. "Optimization Capabilities for Axial Compressor Blades and Seal Teeth Cavity." University of Cincinnati / OhioLINK, 2016. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1458300148.
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Jones, James A. "A multidisciplinary algorithm for the 3-D design optimization of transonic axial compressor blades." Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2002. http://library.nps.navy.mil/uhtbin/hyperion-image/02Jun%5FJones%5FJames.pdf.
Full textAbstract:
Thesis (Ph. D. in Aeronautics and Astronautics)--Naval Postgraduate School, June 2002.Dissertation supervisor: Raymond P. Shreeve. Includes bibliographical references (p. 157-161). Also available online.
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Drayton, Scott. "Design, test, and evaluation of a transonic axial compressor rotor with splitter blades." Thesis, Monterey, California: Naval Postgraduate School, 2013. http://hdl.handle.net/10945/37616.
Full textAbstract:
Approved for public release; distribution is unlimitedA new design procedure was developed and documented that uses commercial-off-the-shelf software (MATLAB, SolidWorks, and ANSYS-CFX) for the geometric rendering and analysis of a transonic axial compressor rotor with splitter blades. Predictive numerical simulations were conducted and experimental data were collected at the NPS TPL utilizing the Transonic Compressor Rig. This study advanced the understanding of splitter blade geometry, placement, and performance benefits. In particular, it was determined that moving the splitter blade forward in the passage between the main blades, which was a departure from the trends demonstrated in the few available previous transonic axial compressor splitter blade studies, increased the mass flow range with no loss in overall performance. With a large 0.91 mm (0.036 in) tip clearance, to preserve the integrity of the rotor, the experimentally measured peak total-to-total pressure ratio was 1.69 and the peak total-to-total isentropic efficiency was 72 percent at 100 percent design speed. Additionally, a higher than predicted 7.5 percent mass flow rate range was experimentally measured, which would make for easier engine control if this concept were to be included in an actual gas turbine engine.
21
Lyes, Peter A. "Low speed axial compressor design and evaluation : High speed representation and endwall flow control studies." Thesis, Cranfield University, 1999. http://hdl.handle.net/1826/4251.
Full textAbstract:
This Thesis reports the design, build and test of two sets of blading for the
Cranfield University low speed research compressor. The first of these was a datum low
speed design based on the fourth stage of the DERA high speed research compressor
C 147. The emphasis of this datum design was on the high-to-low speed transformation
process and the evaluation of such a process through comparing detailed flow
measurements from both compressors.
Area traverse measurements in both the stationary and rotating frame of reference
were taken at Cranfield along with overall performance, blade surface static pressure and
flow visualisation measurements. These compare favourably with traverse and
performance measurements taken on C147 before commencement of the PhD work.
They show that despite the compromises made during the transformation process, due to
both geometric and aerodynamic considerations, both the primary and secondary flow
features can be successfully reproduced in the low speed environment.
The aim of the second design was to improve on the performance of the datum
blading through the use of advanced '3D' design concepts such as lean and sweep. The
blading used nominally the same blade sections as the datum, and parametric studies
were conducted into various lean/sweep configurations to try to optimise the blade
performance. The final blade geometry also incorporated leading edge recambering
towards the fixed endwalls of both the rotor and stator. The '3D' blading demonstrated a
1.5% increase in efficiency (over the datum blading) at design flow rising to around 3%
at near stall along with an improvement in stall margin and pressure rise characteristic.
The design work was completed using the TRANSCode flow solver for both the
blade-to-blade solutions (used in the SI-S2 datum design calculation) and the fully 3D
solutions (for the advanced design and post datum design appraisal). The 3D solutions
gave a reasonable representation of the mid-span and main 3D flow features but failed to
model the corner and tip clearance flow accurately. An interesting feature of the low
speed flowfield was the circumferential variation in total pressure observed at exit from
all rotors for both designs. This was not present at high speed and represents one of the
main differences between the high and low speed flow. Unsteady modelling of mid-
height sections from the first stage indicate that part of this variation is due to the
potential interaction of the rotor with the downstream stator while the remainder is due
to the wake structure from the upstream stator convecting through the rotor passage.
Finally, the implications for a high speed design based on the success of the 3D low
speed design are considered.
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Van, Rooij Michael P. C. "Development of a three-dimensional multistage inverse design method for aerodynamic matching of axial compressor blading." Related electronic resource: Current Research at SU : database of SU dissertations, recent titles available full text, 2008. http://wwwlib.umi.com/cr/syr/main.
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Jia, H.-X., G. Xi, L. Müller, R. Mailach, and K. Vogeler. "Effect of clocking on unsteady rotor blade loading in a low-speed axial compressor at design and off-design operating conditions." Sage, 2008. https://publish.fid-move.qucosa.de/id/qucosa%3A38439.
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This paper presents the results of stator clocking investigations at a design point and an operating point near the stability limit in a low-speed research compressor (LSRC). The unsteady flow field of the LSRC at several clocking configurations was investigated using a three-dimensional unsteady, viscous solver. The unsteady pressure on the rotor blades at midspan (MS) was measured using time-resolving piezoresistive miniature pressure transducers. The effect of clocking on the unsteady pressure fluctuation at MS on the rotor blades is discussed for different operating points. Based on the unsteady profile pressures, the blade pressure forces were calculated. The peak-to-peak amplitudes of the unsteady blade pressure forces are presented and analysed for different clocking positions at both the design point and the operating point near the stability limit of the compressor.
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Li, Yan Sheng. "Mixing in axial compressors." Thesis, University of Cambridge, 1990. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.334235.
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Grimshaw, Samuel David. "Bleed in axial compressors." Thesis, University of Cambridge, 2014. https://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.707970.
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McDougall, Neil Malcolm. "Stall inception in axial compressors." Thesis, University of Cambridge, 1988. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.237803.
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Khalsa, Amrit Singh. "Endwall blockage in axial compressors." Thesis, Massachusetts Institute of Technology, 1996. http://hdl.handle.net/1721.1/10826.
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Power, Bronwyn. "Aspirated compressors." Thesis, University of Cambridge, 2013. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.648363.
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Young, Anna Mollie. "Tip-clearance effects in axial compressors." Thesis, University of Cambridge, 2012. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.610603.
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Seitz, Peter Alexander. "Casing treatment for axial flow compressors." Thesis, University of Cambridge, 1999. https://www.repository.cam.ac.uk/handle/1810/251677.
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Storer, John Andrew. "Tip clearance flow in axial compressors." Thesis, University of Cambridge, 1991. https://www.repository.cam.ac.uk/handle/1810/251503.
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Leggett, John. "Detailed investigation of loss prediction of an axial compressor cascade at off-design conditions in the presence of incident free-stream disturbances using large eddy simulations." Thesis, University of Southampton, 2018. https://eprints.soton.ac.uk/422285/.
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The prediction of an axial compressor’s loss early on in the design phase is a valuable and important part of the design process. The work presented here focuses on assessing the accuracy of current prediction methods, Reynolds Averaged Navier Stokes (RANS), compared with highly accurate Large Eddy Simulations (LES). The simulations were performed at the challenging running conditions of engine relevant Mach (0.67) and Reynolds (300,000) numbers. The work looks at the effects of off-design incidence and the influence of different free-stream disturbances on loss prediction. From the highly accurate datasets produced by the LES the work is able to show how loss attribution varies under different conditions, and goes on to compare how well RANS captures these changes. It was found that overall loss trends are captured well by RANS but substantial differences exist when comparing individual loss sources, which are shown to vary significantly under different running conditions. The investigation into loss attribution is performed using the Denton (1993) loss breakdown as well as a novel application of the Miller (2013) mechanical work potential. In addition to the discovery of the variation in the sources of loss, the comparison between the loss analyses highlighted some of the limitations of the Denton loss breakdown, which was shown to have increasing error under large off-design incidence or in the presence of discrete disturbances. From the comparison of the loss breakdown analyses and LES and RANS flow field results, new insight into the characteristics, limitations and short comings of current modeling techniques have been found. The variation in the sources of loss under different running conditions was also discovered.
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Zaki, Mina Adel. "Physics based modeling of axial compressor stall." Diss., Atlanta, Ga. : Georgia Institute of Technology, 2009. http://hdl.handle.net/1853/31683.
Full textAbstract:
Thesis (Ph.D)--Aerospace Engineering, Georgia Institute of Technology, 2010.Committee Chair: Dr. Lakshmi N. Sankar; Committee Member: Dr. Alex Stein; Committee Member: Dr. J.V. R. Prasad; Committee Member: Dr. Richard Gaeta; Committee Member: Dr. Suresh Menon. Part of the SMARTech Electronic Thesis and Dissertation Collection.
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Wilson, Alexander George. "Stall and surge in axial flow compressors." Thesis, Cranfield University, 1996. http://dspace.lib.cranfield.ac.uk/handle/1826/10432.
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The objective of the work described in this thesis is twofold; to elucidate the nature of
stall and
surge in an axial flow aeroengine compressor, and to improve on current
computational stall modelling techniques. Particular attention is paid to
the initial stages of the stall/surge transient, and to the possibility of using active control
techniques to prevent or delay the onset of stall/surge.
A detailed
analysis is presented of measurements of the stalling behaviour of a Rolls-
Royce VIPER jet engine, showing a wide variety of stall inception and post-stall
behaviour. Stall transients are traced from disturbances through to stable
rotating stall or axisymmetic surge. The stall inception pattern at nearly all speeds is
shown to conform to the short circumferential length scale pattern described by Day
[1993a].
A multiple compressors in parallel stall model is developed using conventional stall
modelling techniques, but extended to include the
effects of the jet engine environment
The model is shown to
give a good representation of the overall stalling behaviour of
the
engine, although the details of the stall inception period are not accurately
predicted. A system identification technique is applied to the results of the model in
order to
develop a method of active control of stall/surge.
A new stall model is introduced and
developed, based on a time-accurate three
dimensional (but pitchwise averaged) solution of the viscous flow equations, with
bladerow performance represented by body forces. The flow in the annulus boundary
layers is calculated directly, and hence this new method is sufficiently complex to
model the initial localised disturbances that lead to
stall/surge. At the same time the
computational power required is compatible with application to long multistage
compressors.
35
Gallimore, Simon John. "Spanwise mixing in multi-stage axial compressors." Thesis, University of Cambridge, 1986. https://www.repository.cam.ac.uk/handle/1810/250879.
Full text
36
Thomas, Keegan Darrall. "Blade row and blockage modelling in an axial compressor throughflow code /." Link to the online version, 2005. http://hdl.handle.net/10019/1239.
Full text
37
Gill, M. E. "Surge prediction in multistage axial and centrifugal compressors." Thesis, Cranfield University, 1986. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.370630.
Full text
38
Paduano, James D. (James Donald). "Active control of rotating stall in axial compressors." Thesis, Massachusetts Institute of Technology, 1992. http://hdl.handle.net/1721.1/11630.
Full text
39
Meehan, Anthony. "Steady state response of an axial compression system to a constant heat input." Thesis, Georgia Institute of Technology, 1999. http://hdl.handle.net/1853/15975.
Full text
40
Li, Yan-Ling. "Numerical simulations of rotating stall in axial flow compressors." Thesis, University of Sussex, 2014. http://sro.sussex.ac.uk/id/eprint/47428/.
Full textAbstract:
Gas turbine compressor performance may encounter deterioration during service for various reasons such as damage by debris from the casing or foreign objects impacting on the blades, typically near the rotor's tip. Moreover, mal-schedule of Variable Stator Vanes (VSVs) during start-up may also result in performance deterioration and reduction in the surge margin. Ability to assess the effect of compressor deterioration using Computational Fluid Dynamics (CFD) is important at both design stage and in service. Compressor blade damage breaks the cyclic symmetry and the VSVs mal-schecule creates mis-match between stages together with geometric variations, thus computations are desirable to be performed using full annulus assemblies. Furthermore, downstream boundary conditions are also unknown during rotating stall or surge and simulations become difficult. This research presents unsteady time-accurate CFD analyses of compressor performance with tip curl blade damage in a single stage axial flow compressor and VSVs mal-schedule in a 3.5 stage axial flow compressor. Computations were per- formed near stall boundary to predict rotating stall characteristics. The primary objectives are to characterise the overall compressor performance and analyse the detailed flow behaviour. Computations for the nominal blade configurations were also performed for comparison purposes for both compressors. All unsteady simulations were performed at part speeds with a variable nozzle downstream representing an experimental throttle. For the blade damage study, two different degrees of damage for one blade and multiple damaged blades were investigated and compared with the results from the undamaged case. For the VSVs mal-schedule study, the first two stators were assumed to be variable and were used to create mal-schedule vane settings for the investigation. The effects of blade damage and VSVs mal-schedule on the aerodynamics performance and rotating stall characteristics for both compressor assemblies were investigated respectively and discussed in detail.
41
DiOrio, Austin Graf. "Small core axial compressors for high efficiency jet aircraft." Thesis, Massachusetts Institute of Technology, 2012. http://hdl.handle.net/1721.1/77107.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2012.Cataloged from department-submitted PDF version of thesis. This electronic version was submitted and approved by the author's academic department as part of an electronic thesis pilot project. The certified thesis is available in the Institute Archives and Special Collections.
Includes bibliographical references (p. 115-117).
This thesis quantifies mechanisms that limit efficiency in small core axial compressors, defined here as compressor exit corrected flow between 1.5 and 3.0 lbm/s. The first part of the thesis describes why a small engine core with high overall pressure ratio (OPR) is desirable for an efficient aircraft and shows that fuel burn can be reduced by up to 17% compared to current engines. The second part examines two specific effects: Reynolds number and tip clearance. At a core size of 1.5 lbm/s, Reynolds number may be as low as 160,000, resulting in reductions in stage efficiency up to 1.9% for blades designed for high Reynolds number flow. The calculations carried out indicate that blades optimized for this Reynolds number can increase stage efficiency by up to 1.6%. For small core compressors, non-dimensional tip clearances are increased, and it is estimated that tip clearances can be up to 4.5% clearance-to-span ratio at the last stage of a 1.5 lbm/s high pressure compressor. The efficiency penalty due to tip clearance is assessed computationally and a 1.6% decrease in polytropic efficiency is found for a 1% increase in gap-to-span ratio. At the above clearance, these efficiency penalties increase aircraft mission fuel burn by 3.4%, if current design guidelines are employed. This penalty, however, may be reduced to 0.4% if optimized blades and a smaller compressor radius than implied by geometric scaling, which allows reduced non-dimensional clearance, are implemented. Based on the results, it is suggested that experiments and computations should be directed at assessing: (i) the effects of clearance at values representative of these core sizes, and (ii) the effect of size on the ability to achieve a specific blade geometry and thus the impact on loss.
by Austin Graf DiOrio.
S.M.
42
Etchevers, Olivier. "Evaluation of rotating stall warning schemes for axial compressors." Thesis, Massachusetts Institute of Technology, 1992. http://hdl.handle.net/1721.1/12939.
Full text
43
Bae, Jinwoo W. "Active control of tip clearance flow in axial compressors." Thesis, Massachusetts Institute of Technology, 2001. http://hdl.handle.net/1721.1/8705.
Full textAbstract:
Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2001.Includes bibliographical references.
Control of compressor tip clearance flows is explored in a linear cascade using three types of fluidic actuators; Normal Synthetic Jet (NSJ; unsteady jet normal to the mean flow with zero net mass flux), Directed Synthetic Jet (DSJ; injection roughly aligned with the mean flow), and Steady Directed Jet (SDJ), mounted on the casing wall. The objective is to affect the following measures: (1) reduction of tip leakage flow rate, (2) mixing enhancement between tip leakage and core flow, and (3) increase in streamwise momentum of the flow in the endwall region. The measurements show that the NSJ provides mixing enhancement only, or both mixing enhancement and leakage flow reduction, depending on its pitchwise location. The DSJ and SDJ actuators provide streamwise momentum enhancement with a consequent reduction of clearance-related blockage. The blockage reduction associated with the use of NSJ is sensitive to actuator frequency, whereas that with the use of DSJ is not. For a given actuation amplitude, DSJ and SDJ are about twice as effective as NSJ in reducing clearance-related blockage. Further the DSJ and SDJ can eliminate clearance-related blockage with a time-averaged momentum flux roughly 16% of the momentum flux of the leakage flow.
(cont.) However, achieving overall gain in efficiency appears to be hard; the decrease in loss is only about 30% of the expended flow power from the present SDJ actuator, which is the best among the actuators considered. Guidelines for improving the efficiency of the directed jet actuation are presented. Time-resolved measurements show periodic unsteadiness of the tip clearance vortex with the peak frequency corresponding to the optimum condition for blockage reduction with the NSJ. A physical explanation of the source of the observed periodic unsteadiness is suggested based on trailing vortex instability theory. Observations of the time scale for the unsteadiness from different compressor geometries and flow conditions are shown to scale with a reduced frequency based on convective time through the blade passage.
by Jinwoo Bae.
Ph.D.
44
White, N. M. "Optimising stator vane settings in multistage axial flow compressors." Thesis, Cranfield University, 2002. http://dspace.lib.cranfield.ac.uk/handle/1826/10756.
Full textAbstract:
There is a common
requirement in the process, oil and gas turbine industries for high
performance axial flow compressors operating over a wide range of mass flow rate and
rotational
speed at high efficiency. The trends have been for higher blade loadings
(greater pressure rise per stage) and higher efficiency which are increasingly achieved
through sophisticated Computational Fluid Dynamics designs. These trends, however,
tend to
mitigate against stable operating flow range (reduced surge margin), which can
often lead to
performance compromises. The objective of this work is to investigate the
possibility of using alternative means to gain ow range by better use of variable
geometry, which may permit design objectives to be better achieved. Variable
geometry of the type envisaged is already often employed to overcome part-speed
operating problems, but it proposed here that there may be additional benefits from their
more
intelligent control.
The
operation of axial compressors with a wide operating range is limited by
instabilities, which cause a full breakdown of the flow, which is surge. These
instabilities, which are caused by high incidence and subsequent stalling of stages occur
due to different
phenomena at part and full speed operation. The problem at part-speed
is that the front
stages are often heavily stalled and rear stages choked, whereas at high
speeds, the front stages are operating close to choke and the rear stages tend to be
stalling. Optimisation of the design to full load conditions can often provide part-speed
problems and to achieve the acceptable performance, variable geometry over the front
region of the compressor is sometimes used to modify the flow angles and avoid stage
stall and
subsequent surge. To-date, such variable settings follow some schedule
established
by analysis and experiment whereas this work presents a methodology of
setting blade rows using an optimisation procedure and investigates the likelihood of
performance benefits being obtained by a control technique which reacts° to these
changing conditions.
The construction of the numerical method
presented in this thesis was done with an
emphasis upon its intended contribution towards a eventual online control application.
Therefore, a practical approach has been employed in the development of the
compressor modelling techniques used in the work. Specifically, a highly empirical
one-dimensional
performance prediction code was constructed, employing successful
correlations taken from the literature. This was
coupled to a surge prediction method
that has been shown in the
past to function more than satisfactorily in a multistage
environment.
Finally, the predicted stage and overall performance (including the surge
point) characteristics were passed to a optimisation program, which allowed these
simulated conditions to be
investigated.
It is
hoped that the work presented has illustrated the potential (from a aerodynamic
performance point of view) of such a control technique to offer additional freedom in
the
operation of a multistage axial flow compressor. Moreover, the numerical
modelling techniques have been developed enough to envisage (at least in part) their
simple integration within a practical system. Clearly, some further investigations are
required to take this work forward and the next logical step would be to improve the
empirical rules with which the blade performance is predicted. A experimental
programme would also be of great advantage, for example in the study of how the
deviation
angle for a given blade row varies over time (operating hours) in a real
machine due to
ageing and fouling. This would allow better estimates of the stage work
during long term operation so that the optimiser could adapt to the slowly degrading
performance of the blades. Finally, it is important to verify the simulated results with
measured data, taken at the same optimal stator vane settings as given by the program.
This must be carried out before it can be
applied to a real application, although a limited
study of this nature is presented in chapter 6.
45
White, Nicholas M. "Optimising stator vane settings in multistage axial flow compressors." Thesis, Cranfield University, 2002. http://dspace.lib.cranfield.ac.uk/handle/1826/10756.
Full textAbstract:
There is a common requirement in the process, oil and gas turbine industries for high performance axial flow compressors operating over a wide range of mass flow rate and rotational speed at high efficiency. The trends have been for higher blade loadings (greater pressure rise per stage) and higher efficiency which are increasingly achieved through sophisticated Computational Fluid Dynamics designs. These trends, however, tend to mitigate against stable operating flow range (reduced surge margin), which can often lead to performance compromises. The objective of this work is to investigate the possibility of using alternative means to gain ow range by better use of variable geometry, which may permit design objectives to be better achieved. Variable geometry of the type envisaged is already often employed to overcome part-speed operating problems, but it proposed here that there may be additional benefits from their more intelligent control. The operation of axial compressors with a wide operating range is limited by instabilities, which cause a full breakdown of the flow, which is surge. These instabilities, which are caused by high incidence and subsequent stalling of stages occur due to different phenomena at part and full speed operation. The problem at part-speed is that the front stages are often heavily stalled and rear stages choked, whereas at high speeds, the front stages are operating close to choke and the rear stages tend to be stalling. Optimisation of the design to full load conditions can often provide part-speed problems and to achieve the acceptable performance, variable geometry over the front region of the compressor is sometimes used to modify the flow angles and avoid stage stall and subsequent surge. To-date, such variable settings follow some schedule established by analysis and experiment whereas this work presents a methodology of setting blade rows using an optimisation procedure and investigates the likelihood of performance benefits being obtained by a control technique which reacts° to these changing conditions. The construction of the numerical method presented in this thesis was done with an emphasis upon its intended contribution towards a eventual online control application. Therefore, a practical approach has been employed in the development of the compressor modelling techniques used in the work. Specifically, a highly empirical one-dimensional performance prediction code was constructed, employing successful correlations taken from the literature. This was coupled to a surge prediction method that has been shown in the past to function more than satisfactorily in a multistage environment. Finally, the predicted stage and overall performance (including the surge point) characteristics were passed to a optimisation program, which allowed these simulated conditions to be investigated. It is hoped that the work presented has illustrated the potential (from a aerodynamic performance point of view) of such a control technique to offer additional freedom in the operation of a multistage axial flow compressor. Moreover, the numerical modelling techniques have been developed enough to envisage (at least in part) their simple integration within a practical system. Clearly, some further investigations are required to take this work forward and the next logical step would be to improve the empirical rules with which the blade performance is predicted. A experimental programme would also be of great advantage, for example in the study of how the deviation angle for a given blade row varies over time (operating hours) in a real machine due to ageing and fouling. This would allow better estimates of the stage work during long term operation so that the optimiser could adapt to the slowly degrading performance of the blades. Finally, it is important to verify the simulated results with measured data, taken at the same optimal stator vane settings as given by the program. This must be carried out before it can be applied to a real application, although a limited study of this nature is presented in chapter 6.
46
Bloch, Gregory S. "A wide-range axial-flow compressor stage performance model." Thesis, This resource online, 1991. http://scholar.lib.vt.edu/theses/available/etd-08182009-040326/.
Full text
47
Baker, Jonathan D. "Analysis of the sensitivity of multi-stage axial compressors to fouling at various stages." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2002. http://library.nps.navy.mil/uhtbin/hyperion-image/02Sep%5FBaker.pdf.
Full text
48
Gill, Andrew. "A comparison between stall prediction models for axial flow compressors." Thesis, Stellenbosch : Stellenbosch University, 2006. http://hdl.handle.net/10019.1/18702.
Full textAbstract:
Thesis (MScEng)--Stellenbosch University, 2006.ENGLISH ABSTRACT: The Stellenbosch University Compressor Code (SUCC) has been developed for the purpose of predicting the performance of axial flow compressors by means of axisymmetric inviscid throughflow methods with boundary layer blockage and empirical blade row loss models. This thesis describes the process of the implementation and verification of a number of stall prediction criteria in the SUCC. In addition, it was considered desirable to determine how certain factors influence the accuracy of the stall prediction criteria, namely the nature of the computational grid, the choice of throughflow method used, and the use of a boundary layer blockage model and a radial mixing model. The stall prediction criteria implemented were the di®usion factor limit criterion, de Haller's criterion, Aungier's blade row criterion, Aungier's boundary layer separation criterion, Dunham's, Aungier's and the static-to-static stability criteria. The compressors used as test cases were the Rofanco 3-stage low speed compressor, the NACA 10-stage subsonic compressor, and the NACA 5-stage and 8-stage transonic compressors. Accurate boundary layer blockage modelling was found to be of great importance in the prediction of the onset of stall, and that the matrix throughflow Method provided slightly better accuracy than the streamline curvature method as implemented in the SUCC by the author. The ideal computational grid was found to have many streamlines and a small number of quasi-orthogonals which do not occur inside blade rows. Radial mixing modelling improved the stability of both the matrix throughflow and streamline curvature methods without significantly affecting the accuracy of the stall prediction criteria. De Haller's criterion was over-conservative in estimating the stall line for transonic conditions, but more useful in subsonic conditions. Aungier's blade row criterion provided accurate results on all but the Rofanco compressor. The diffusion factor criterion provided over- optimistic predictions on all machines, but was less inaccurate than de Haller's criterion on the NACA 5-stage transsonic machine near design conditions. The stability methods performed uniformly and equally badly, supporting the claims of other researchers that they are of limited usefulness with throughflow simulations. Aungier's boundary layer separation method failed to predict stall entirely, although this could reflect a shortcoming of the boundary layer blockage model.
AFRIKAANSE OPSOMMING: Die Stellenbosch University Compressor Code (SUCC) is ontwikkel om die prestasie van aksiaalvloei kompressors te voorspel met behulp van aksisimmetriese nie-viskeuse deurvloeimetodes met grenslaagblokkasie en empiriese modelle vir die verliese binne lemrye. Hierdie tesis beskryf die proses waarmee sekere staakvoorspellingsmetodes in die SUCC geïmplementeer en geverifieer is. Dit was ook nodig om die effek van sekere faktore, naamlik die vorm van die berekeningsrooster, die keuse van deurvloeimetode en die gebruik van `n grenslaagblokkasiemodel en radiale vloeivermengingsmodel op die akuraatheid van die staakvoorspellingsmetodes te bepaal. Die staakvoorspellingsmetodes wat geïmplementeer is, is die diffusie faktor beperking metode, de Haller se metode, Aungier se lemrymetode, Aungier se grenslaagmetode en die Dunham, Aungier en die statiese-tot-statiese stabiliteitsmetodes. Die kompressors wat gebruik is om die metodes te toets is die Rofanco 3-stadium lae-spoed kompressor, die NACA 10-stadium subsoniese kompressor en die NACA 5- en 8-stadium transsoniese kompressors. Daar is vasgestel dat akkurate grenslaagblokkasie modelle van groot belang was om `n akkurate aanduiding van die begin van staking te voorspel, en dat, vir die SUCC, die Matriks Deurvloei Metode oor die algemeen 'n bietjie meer akkuraat as die Stroomlyn Kromming Metode is. Daar is ook vasgestel dat die beste berekeningsrooster een is wat baie stroomlyne, en die kleinste moontlike getal quasi-ortogonale het, wat nie binne lemrye geplaas mag word nie. Die numeriese stabiliteit van beide die Matriks Deurvloei en die Stroomlyn Kromming Metode verbeter deur gebruik te maak van radiale vloeivermengingsmodelle, sonder om die akkuraatheid van voorspellings te benadeel. De Haller se metode was oorkonserwatief waar dit gebruik is om die staak-lyn vir transsoniese vloei toestande, maar meer nuttig in die subsoniese vloei gebied. Aungier se lemrymetode het akkurate resultate gelewer vir alle kompressors getoets, behalwe die Rofanco. Die diffusie faktor metode was oor die algemeen minder akuraat as Aungier se metode, maar meer akkuraat as de Haller se metode vir transsoniese toestande. Die stabiliteitsmetodes het almal ewe swak gevaar. Dit stem ooreen met die bevindings van vorige navorsing, wat bewys het dat hierdie metodes nie toepaslik is vir simulasies wat deurvloeimetodes gebruik nie. Aungier se grenslaagmetode het ook baie swak gevaar. Waarskynlik is dit as gevolg van tekortkomings in die grenslaagblokkasiemodel.
49
Pantelidis, Konstantinos. "Reynolds number effects on the aerodynamics of compact axial compressors." Thesis, University of Cambridge, 2018. https://www.repository.cam.ac.uk/handle/1810/284940.
Full textAbstract:
An axial compressor for a domestic appliance can be designed to be smaller than an equivalent centrifugal compressor. However, the performance of such a compact axial compression system is limited by increased viscous losses and reduced flow turning at low Reynolds numbers ($Re$). In domestic appliance compressors, $Re$ is typically in the range $10^4$ - $10^5$. Although the aerodynamics of isolated aerofoils operating at these $Re$ have been studied extensively, the flow fields within low $Re$ axial compressors have not been investigated in detail. This dissertation aims to develop an improved understanding of loss variation at low $Re$ and to explore how the losses can be reduced through design changes. Experiments on a 5 times scaled-up single stage axial compressor have been conducted across a range of $Re$ of $10^4$ - $10^5$. The flow field has been characterised using detailed area traverses with a miniaturised five-hole probe at the rotor inlet, rotor exit and stator exit and a miniature hot-wire at the rotor exit. The probe was specifically designed and calibrated for the scale of the experiments and methods to improve the accuracy of the measurements have been applied including a probe geometry correction. The traverse experiments were performed at the design operating condition ($\phi=0.55$ and $Re= 6\times10^4$) and at a condition close to stall for a datum stage design, a stage with an improved stator design and two stators with compound lean. It was found that losses in the rotor were greater than the stator losses across the whole range of $Re$. As expected, the loss decreased with increasing $Re$ for both the stator and rotor. The losses were also increased by three-dimensional flow, with typical loss coefficients at the hub and tip of the blade rows in the range of $20-30\%$. A major contributor to the rotor loss was an unexpected hub separation that increased in size as $Re$ was reduced. At higher $Re$, the major loss sources were the rotor tip leakage, the stator wake and the stator hub separation. The results indicate that an improved stator design that accounts for the actual, measured, rotor exit flow field at low $Re$ could reduce the $Re$ at which blade row losses start to rise dramatically as well as reduce the loss across all $Re$. The improved stator design was better matched to the radial distribution of rotor exit flow angle, which led to a decrease in stator loss across all $Re$. For all stator designs, however, the measured stage stall margin was identical at all $Re$. This, along with the increase in velocity deficit in the rotor tip region at off-design indicates that stall occurred in the rotor and was neither $Re$ nor stator design dependent. The introduction of compound lean to the the stator design had the expected result of decreasing the endwall corner separation loss and increasing midspan losses. The experiments have shown that there was a loss increase in both the midspan and casing region much greater than the corresponding decrease in the stator hub. Also the mass flow redistribution in the experiments was larger that the redistribution predicted by the CFD. Three-dimensional RANS computations at low $Re$ of the same designs as experimentally studied were also conducted in order to investigate the predictive accuracy of industry standard CFD. The simulation results predicted the overall loss distribution but overestimated the end-wall losses and failed to capture the drop in stage performance at low $Re$. The differences with the experiments were caused by the inherent limitations of a fully turbulent solver that cannot reproduce transitional flow-features. Similarly to the experiments, there was no stall margin dependency on $Re$ in the simulations. This thesis has shown that with axial compressors designed specifically for low $Re$, the $Re$ at which the losses start increasing exponentially can be shifted from $10\times10^4$ to $ 4\times10^4$. The loss increase is predominantly caused by the rotor hub corner separation.
50
Lavrich, Philip Lewis. "Time resolved measurements of rotating stall in axial flow compressors." Thesis, Massachusetts Institute of Technology, 1988. http://hdl.handle.net/1721.1/14567.
Full text