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Relevant bibliographies by topics / Aircraft Engine Compressor Blade

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Academic literature on the topic 'Aircraft Engine Compressor Blade'

Author: Grafiati

Published: 4 June 2021

Last updated: 6 September 2023

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Journal articles on the topic "Aircraft Engine Compressor Blade"

1

Gilge, Philipp, Andreas Kellersmann, Jens Friedrichs, and Jörg R. Seume. "Surface roughness of real operationally used compressor blade and blisk." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 233, no. 14 (2019): 5321–30. http://dx.doi.org/10.1177/0954410019843438.

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Deterioration of axial compressors is in general a major concern in aircraft engine maintenance. Among other effects, roughness in high-pressure compressor reduces the pressure rise and thus efficiency, thereby increasing the specific fuel consumption of an engine. Therefore, it is important to improve the understanding of roughness on compressor blading and their impact on compressor performance. To investigate the surface roughness of rotor blades of a compressors, different stages of an axial high-pressure compressor and a first-stage blisk (BLade–Integrated–dISK) of a regional aircraft eng

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2

Potapov, V. A., and A. A. Sanko. "Performance simulation of multi-stage axial-flow compressor of turbo-shaft engine with account for erosive wear nonlinearity of its blades." Civil Aviation High Technologies 23, no. 5 (2020): 39–53. http://dx.doi.org/10.26467/2079-0619-2020-23-5-39-53.

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The construction and useful practice of gas-turbine engine diagnosis systems depend largely on the availability of the engine mathematical models and its certain components in their structure. Utilization of multi-stage axial flow compressor performance with account for erosive wear of its parts during the operation fundamentally raises possibilities of such systems as erosive wear of flow channel, blade rings of impellers and vane rings of multi-stage compressor is a common cause of preschedule gas-turbine engine detaching from an aircraft. As evidenced by various contributions presented in t

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3

ILIE, Marcel, and Augustin SEMENESCU. "AERODYNAMIC STUDIES OF AIRCRAFT ENGINE TURBINE STAGE." ANNALS OF THE ACADEMY OF ROMANIAN SCIENTISTS Series on ENGINEERING SCIENCES 14, no. 2 (2022): 5–18. http://dx.doi.org/10.56082/annalsarscieng.2022.2.5.

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The present research concerns the aerodynamic computational studies of stator-rotor turbine stage. The computational studies are carried out using the large-eddy simulation approach. In the aircraft engine compressor/turbine stage blade-vortex interactions occur. The present study aims at the understanding the blade-vortex interaction mechanism and its impact on the aerodynamics of rotor-stator compressor/turbine stages. The computational studies are carried out in a rotating frame of reference, for high-Reynolds number flow, Re = 1.3x105. The analysis reveals that the blade-vortex interaction

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4

Obrocki, Wojciech, Amadeusz Setkowicz, Maciej Masłyk, and Jan Sieniawski. "Influence of aircraft engine compressor blades leading edge damage on their fatigue strength." Mechanik 91, no. 3 (2018): 205–9. http://dx.doi.org/10.17814/mechanik.2018.3.35.

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Article presents the research results of aircraft compressor blade damage length and its position influence on fatigue strength under high number cycles conditions. The criteria for blade damage detection classification and test research methodology were developed. Designed and tested the instrumentation for compressor blades fatigue tests. Fluorescent method was used to determine the source of fatigue cracking initiation and its propagation direction during fatigue test.

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Kvasha, Yu A., and N. A. Zinevych. "Aerodynamic improvement of an aircraft gas-turbine engine fan." Technical mechanics 2021, no. 3 (2021): 23–29. http://dx.doi.org/10.15407/itm2021.03.023.

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This work is concerned with the development of approaches to the aerodynamic improvement of axial-flow compressors for gas-turbine engines. The aim of this work is the aerodynamic improvement of an aircraft gas-turbine engine two-stage fan by numerical simulation of 3D turbulent gas flows. The approach used in this study features: varying the spatial shape of the fan blades for the first- and the second-stage impeller by varying the profile angle along the blade height; formulating quality criteria as the mean integral values of the power characteristics of each impeller of the fan over the op

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Lee, Bok-Won, Jungjun Suh, Hongchul Lee, and Tae-gu Kim. "Investigations on fretting fatigue in aircraft engine compressor blade." Engineering Failure Analysis 18, no. 7 (2011): 1900–1908. http://dx.doi.org/10.1016/j.engfailanal.2011.07.021.

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7

Ngoret, Joshua Kimtai, and Venkata Parasuram Kommula. "Comprehending Occurrence of Premature Failure in Compressor Turbine (CT) Blades for Short-Haul Aircraft Fleet." International Journal of Engineering Research in Africa 42 (April 2019): 10–23. http://dx.doi.org/10.4028/www.scientific.net/jera.42.10.

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This paper presents results from modeling of Compressor Turbine (CT) blades for short-haul aircraft fleet occasioned by thermo-mechanical stresses in order to comprehend the occurrence of premature failure. A 3D PT6A-114A engine high pressure (HP) CT blade geometrical model was developed in commercial CAD-SolidWorks, then imported to ANSYS 15.0 environment for finite element analysis (FEA). The CT blade was investigated for transient thermal stresses from heat generated by the combustors and static structural stresses from rotational velocities of the engine which account for 80% of inertial f

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8

Boyko, Ludmila, Vadym Datsenko, Aleksandr Dyomin, and Nataliya Pizhankova. "Devising a method for calculating the turboshaft gas turbine engine performance involving a blade-by-blade description of the multi-stage compressor in a two-dimensional setting." Eastern-European Journal of Enterprise Technologies 4, no. 8(112) (2021): 59–66. http://dx.doi.org/10.15587/1729-4061.2021.238538.

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The design and adjustment of modern gas turbine engines significantly rely on the use of numerical research methods. This paper reports a method devised for calculating the thermogasdynamic parameters and characteristics of a turboshaft gas turbine engine. The special feature of a given method is a two-dimensional blade-by-blade description of the compressor in the engine system. Underlying the calculation method is a nonlinear mathematical model that makes it possible to describe the established processes occurring in individual nodes and in the engine in general. To build a mathematical mode

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9

MacLeod, J. D., and J. C. G. Laflamme. "Compressor Coating Effects on Gas Turbine Engine Performance." Journal of Engineering for Gas Turbines and Power 113, no. 4 (1991): 530–34. http://dx.doi.org/10.1115/1.2906273.

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In an attempt to increase the time between maintenance actions and to improve performance retention of turboprop engines installed in transport and maritime patrol aircraft, the Canadian Department of National Defence is evaluating an erosion and corrosion-resistant blade coating, for use on compressors. As coatings could appreciably alter engine performance by virtue of their application thickness and surface quality, the National Research Council of Canada was asked to quantify any performance changes that could occur. A project was initiated, utilizing a new Allison T56 turboprop engine, to

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10

Xu, Wen Chen, De Bin Shan, Wei Feng Zhang, and Yan Lu. "Precision Forging Process Analysis and Preform Optimization of Small Compressor Blade of Titanium Alloy." Applied Mechanics and Materials 364 (August 2013): 493–99. http://dx.doi.org/10.4028/www.scientific.net/amm.364.493.

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Small compressor blades of titanium alloy are important mechanical structural components of advanced aircraft engines. In the present study, the preheating scheme and the forging process of the small compressor blade were modeled by the commercial software DEFORM3D and the forging experiments were conducted on a 10,000KN screw press. The results show that the reasonable preheating temperature for the small blade should be controlled in 960-970°C. The geometric shape and size of the blade preform influenced not only overlap and underfilling defects but also the dimensional precision of small co

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Dissertations / Theses on the topic "Aircraft Engine Compressor Blade"

1

Yarlagadda, Santosh. "Performance Analysis of J85 Turbojet Engine Matching Thrust with Reduced Inlet Pressure to the Compressor." University of Toledo / OhioLINK, 2010. http://rave.ohiolink.edu/etdc/view?acc_num=toledo1271367584.

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Borgoltz, Aurelien. "Modifications of Coherent Structures in Fan Blade Wakes for Broadband Noise Reduction." Diss., Virginia Tech, 2007. http://hdl.handle.net/10919/29619.

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The effects of trailing edge flow control on the wakes of a linear cascade of idealized fan blades was investigated experiments with a view to the likely effects on broadband aircraft engine interaction noise. Single and three-component hotwire velocity measurements were made downstream of the cascade for a chord Reynolds number of 390,000 and a Mach number of 0.07. Measurements of the two-point velocity correlation were used extensively to evaluate the impact of various flow control strategies on the organization of the coherent structures of the wakes and their potential to generate noise.

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Geiger, Derek Henry. "Comparative Analysis of Serrated Trailing Edge Designs on Idealized Aircraft Engine Fan Blades for Noise Reduction." Thesis, Virginia Tech, 2004. http://hdl.handle.net/10919/40542.

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The effects of serrated trailing edge designs, designed for noise reduction, on the flow-field downstream of an idealized aircraft engine fan blade row were investigated in detail. The measurements were performed in the Virginia Tech low speed linear cascade tunnel on one set of baseline GE-Rotor-B blades and four sets of GE-Rotor-B blades with serrated trailing edges. The four serrated blade sets consisted of two different serration sizes (1.27 cm and 2.54 cm) and for each different serration size a second set of blades with added trailing edge camber. The cascade row consisted of 8 GE-Rotor-

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Parthasarathy, Nikhil Kaushik. "An efficient algorithm for blade loss simulations applied to a high-order rotor dynamics problem." Thesis, Texas A&M University, 2003. http://hdl.handle.net/1969.1/189.

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In this thesis, a novel approach is presented for blade loss simulation of an aircraft gas turbine rotor mounted on rolling element bearings with squeeze film dampers, seal rub and enclosed in a flexible housing. The modal truncation augmentation (MTA) method provides an efficient tool for modeling this large order system with localized nonlinearities in the ball bearings. The gas turbine engine, which is composed of the power turbine and gas generator rotors, is modeled with 38 lumped masses. A nonlinear angular contact bearing model is employed, which has ball and race degrees of freedom and

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Gongol, Jakub. "Návrh malého proudového motoru do 1kN tahu." Master's thesis, Vysoké učení technické v Brně. Fakulta strojního inženýrství, 2013. http://www.nusl.cz/ntk/nusl-230963.

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This work will be focused on issue of a jet engine. The thesis will be divided into search retrieval part and computational part. In the search retrieval part it will focus on different configurations of jet engines as well as areas of their use. The main part of the thesis will however focus on a calculations where a turbine, compressor and an exhaust nozzle will be designed in order to give a thrust of approximately 1kN. Next step will be determination of an engine charcteristic that will give us a preview on how the engine performance will look like in off-design modes.

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Subramanya, S. "Prediction of Physical Behavior of Rotating Blades under Tip-Rub Impact using Numerical Modeling." Thesis, 2013. http://hdl.handle.net/2005/3083.

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Rotating blades, which are the most critical components of any turbo-machinery, need to be designed to withstand forced vibrations due to accidental tip rub impact against inner surface of casing. These vibrations are typically dependent on operating conditions and geometric parameters. In the current study, a rotor test rig with a maximum tip speed capability of 144 km/hr has been developed for studying the dynamic behavior of representative jet engine compressor blades actuated by the closure of clearance between the tip of a given rotating blade and a sector of the inner lining of the casin

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7

Horng, Jyi-Yeong, and 洪吉永. "Diffusion Bonding and Hot Twist Forming of the Aircraft Engine Fan Blade." Thesis, 1999. http://ndltd.ncl.edu.tw/handle/21302625648179834356.

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碩士<br>國立臺灣大學<br>機械工程學研究所<br>87<br>This thesis is aimed at studying the diffusion bonding and hot twist forming processes for producing aircraft engine hollow fan blades. It is divided into two main parts. The first part is for the diffusion bonding process that includes the experimental studies and numerical simulations. Several die and blade specimens are designed and manufactured for carrying out the experiments. Finite element methods (FEM) are performed to obtain stress distribution on the bonding interface. The simulation result is compared with the experiment to identify the effect of th

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Books on the topic "Aircraft Engine Compressor Blade"

1

Kelly, Carney, Gallardo Vicente, and NASA Glenn Research Center, eds. Simulation of aircraft engine blade-out structural dynamics. National Aeronautics and Space Administration, Glenn Research Center, 2001.

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2

Kelly, Carney, Gallardo V. C, and NASA Glenn Research Center, eds. Simulation of aircraft engine blade-out structural dynamics. National Aeronautics and Space Administration, Glenn Research Center, 2001.

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3

S, Reddy E., and United States. National Aeronautics and Space Administration., eds. Root damage analysis of aircraft engine blade subject to ice impact. National Aeronautics and Space Administration, 1992.

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4

L, Mattern Duane, Le Dzu K, and United States. National Aeronautics and Space Administration., eds. Comparisons of rig and engine dynamic events in the compressor of an axi-centrifugal turboshaft engine. National Aeronautics and Space Administration, 1996.

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5

L, Mattern Duane, Le Dzu K, and United States. National Aeronautics and Space Administration., eds. Comparisons of rig and engine dynamic events in the compressor of an axi-centrifugal turboshaft engine. National Aeronautics and Space Administration, 1996.

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6

Heidelberg, Laurence J. Advanced turboprop wing installation effects measured by unsteady blade pressure and noise. National Aeronautics and Space Administration, 1987.

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7

United States. General Accounting Office., ed. Protest Of Air Force Contract Award For Aircraft Engine Blade Sets... 155177, B-261680... U.S. GAO... September 8, 1995. s.n., 1997.

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8

Analysis of aircraft engine blade subject to ice impact. National Aeronautics and Space Administration, 1991.

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9

Root damage analysis of aircraft engine blade subject to ice impact. National Aeronautics and Space Administration, 1992.

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10

Structural tailoring of aircraft engine blade subject to ice impact constraints. National Aeronautics and Space Administration, 1993.

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Book chapters on the topic "Aircraft Engine Compressor Blade"

1

Adams, Maurice L. "Aircraft Engine Compressor Blade Tip Rubs." In Rotating Machinery Research and Development Test Rigs. CRC Press, 2017. http://dx.doi.org/10.1201/9781315116723-15.

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2

Decher, Reiner. "Bypass and Other Engines." In The Vortex and The Jet. Springer Singapore, 2022. http://dx.doi.org/10.1007/978-981-16-8028-1_11.

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AbstractThe modern fan engine has a propeller-like component whose aerodynamic performance differs from that of a propeller at the blade level but serves the same function. The blading aerodynamics differs from that of a compressor in notable ways.

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3

Filipkovskij, S. V., V. S. Chigrin, A. A. Sobolev, and E. T. Vasilevskij. "Simulation of Aircraft Engine Dynamic Effect on Aircraft Wing Caused by a Fan Blade-Off." In Integrated Computer Technologies in Mechanical Engineering - 2022. Springer Nature Switzerland, 2023. http://dx.doi.org/10.1007/978-3-031-36201-9_35.

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4

Van Der Merwe, Shade Rouxzeta, Daniel Ogochukwu Okanigbe, Dawood Ahmed Desai, and Glen Snedden. "A Review on Impact Resistance of Partially Filled 3D Printed Titanium Matrix Composite Designed Aircraft Turbine Engine Fan Blade." In The Minerals, Metals & Materials Series. Springer International Publishing, 2022. http://dx.doi.org/10.1007/978-3-030-92381-5_63.

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Van Der Merwe, Shade Rouxzeta, Daniel Ogochukwu Okanigbe, Dawood Ahmed Desai, and Glen Campbell Snedden. "Aircraft Engine Fan Blade Design: Impact Tolerance Prediction of Partially Filled 3D Printed Aluminum, Titanium, and PEEK-Filled Waste Metal Dusts." In Resource Recovery and Recycling from Waste Metal Dust. Springer International Publishing, 2023. http://dx.doi.org/10.1007/978-3-031-22492-8_10.

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6

Martynenko, Gennadii, Natalia Smetankina, Volodymyr Martynenko, Vyacheslav Merculov, and Mykola Kostin. "Influence of Using Different Material Models of an Aircraft Gas Turbine Engine Fan Blade and a Bird when Simulating the Dynamics of a Collision Process in Flight." In Integrated Computer Technologies in Mechanical Engineering - 2022. Springer Nature Switzerland, 2023. http://dx.doi.org/10.1007/978-3-031-36201-9_33.

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7

"Failure of a Compressor Blade in an Aircraft Engine." In Failure Analysis of Engineering Structures. ASM International, 2005. http://dx.doi.org/10.31399/asm.tb.faesmch.t51270141.

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"Failure of a Compressor Blade in an Aircraft Engine." In Failure Analysis of Engineering Structures. ASM International, 2005. http://dx.doi.org/10.31399/asm.tb.faesmch.t51270150.

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9

"Failure of a First-Stage Compressor Blade in an Aircraft Engine." In Failure Analysis of Engineering Structures. ASM International, 2005. http://dx.doi.org/10.31399/asm.tb.faesmch.t51270118.

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"Failure of a First-Stage Compressor Blade in an Aircraft Engine." In Failure Analysis of Engineering Structures. ASM International, 2005. http://dx.doi.org/10.31399/asm.tb.faesmch.t51270128.

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Conference papers on the topic "Aircraft Engine Compressor Blade"

1

Brandes, Tim, Christian Koch, and Stephan Staudacher. "Estimation of Aircraft Engine Flight Mission Severity Caused by Erosion." In ASME Turbo Expo 2020: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2020. http://dx.doi.org/10.1115/gt2020-16297.

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Abstract More and more attention is being devoted to assessing severity of the engine operation for a high number of flights in a minimum of time. Compressor erosion is one of the physical phenomena contributing to this severity. Hence, an effective method is developed which allows a general judgment of the severity of engine operation with regards to compressor erosion. The shortening of the camber line at blade leading edge is selected as the parameter describing the degree of severity. The particle impingement conditions experienced by compressor blades throughout a flight mission are compu

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2

Drewczynski, Marcin, Romuald Rzadkowski, and Zofia Ostrowska. "Dynamic Stress Analysis of a Blade in a Partially Blocked Engine Inlet." In ASME Turbo Expo 2014: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2014. http://dx.doi.org/10.1115/gt2014-26011.

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One of the greatest dangers to properly working aircraft engines, apart from human error, is the ingestion foreign objects into the compressor. These objects may be birds, dust, hail, volcanic ash, ice on wings but also aircraft parts or simply garbage. Modern jet engines can suffer major damage from even small objects being sucked into the engine. We want to show how foreign object debris affects the dynamic stress level of a rotor blade. A comparison between an undisturbed engine inlet and one with an ingested foreign object is carried out. The analysis will focus on the first stage compress

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3

Batailly, Alain, Markus B. Meingast, Mathias Legrand, and Jean-Philippe Ousty. "Rotor-Stator Interaction Scenarios for the Centrifugal Compressor of a Helicopter Engine." In ASME 2013 International Design Engineering Technical Conferences and Computers and Information in Engineering Conference. American Society of Mechanical Engineers, 2013. http://dx.doi.org/10.1115/detc2013-12211.

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The present paper deals with interaction phenomena that may arise in the centrifugal compressor of helicopter engines when structural contacts occur between the blade-tips and the surrounding casing. Those phenomena may lead to increased levels of vibration and are currently under investigation for axial compressors of aircraft engines for which blade failures were observed. The growing understanding of these phenomena allows for more challenging numerical simulations with increased geometrical complexity such as centrifugal compressors. The simulation of blade-tip/casing contacts is carried o

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Batailly, Alain, Mathias Legrand, Antoine Millecamps, and François Garcin. "High-Pressure Compressor Blade Dynamics Under Aerodynamic and Blade-Tip Unilateral Contact Forcings." In ASME Turbo Expo 2014: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2014. http://dx.doi.org/10.1115/gt2014-25675.

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Recent studies focused on the numerical prediction of structural instabilities that may arise in rotating components of an aircraft engine. These instabilities are commonly classified into two categories: those induced by aerodynamic phenomena (such as the pressure applied on the blade by the incoming air flow) and those related to structural phenomena (such as potential blade/casing contacts). Based on an existing numerical strategy for the analysis of rotor/stator interactions induced by unilateral contacts between rotating and static components, this paper aims at combining both types of in

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Döring, Felix, Stephan Staudacher, and Christian Koch. "Predicting the Temporal Progression of Aircraft Engine Compressor Performance Deterioration due to Particle Deposition." In ASME Turbo Expo 2017: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2017. http://dx.doi.org/10.1115/gt2017-63544.

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Aircraft engine performance deterioration due to particle deposition on compressor blading and end walls gradually progresses with increasing time of operation. Deposition effects can be mitigated by on-wing maintenance actions. Application of condition based maintenance strategies in order to minimize operating costs requires high-grade physical deterioration models. In previous work, a model and experimental setup were developed to quantify both magnitudes and timescales of deposition effects on blade row performance as a function of engine operating time. The model and experimental data pub

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Anbarasan, Selwyn, S. Esakki Muthu, Hardik Roy, P. Udayanan, and Girish K. Degaonkar. "Residual Life Estimation of Axial Compressor Blade of a Turbo-Shaft Engine." In ASME 2014 Gas Turbine India Conference. American Society of Mechanical Engineers, 2014. http://dx.doi.org/10.1115/gtindia2014-8241.

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In aeronautical industry, flight safety is the first and foremost concern. Structural failure in aero engine aids to high risk in flight safety and human lives. High Cycle Fatigue (HCF) failure accounts for forty percent of blade structural failure. All critical components in the aero engine are life limited components and are replaced when its prescribed life is reached. Earlier components are designed as per safe-life design philosophy. Ninety percent of the critical components are retired utilizing less than fifty percent of its safe-life capability. Extending the life of component reduces

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Rao, J. S., Narayan Rangarajan, Rejin Ratnakar, R. Rzadkowski, M. Soliński, and I. Vorobiev. "Life Calculation of First Stage Compressor Blade of a Trainer Aircraft." In ASME Turbo Expo 2012: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2012. http://dx.doi.org/10.1115/gt2012-68070.

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This paper is concerned with life estimation of a compressor rotor blade of an engine at the Air Force Institute of Technology in Warsaw used in a trainer aircraft. A bird strike is simulated by two or three blade passage blocks in the incoming flow and the pressure field is obtained from a CFD code. For the blade the Campbell diagram is prepared and the critical speeds are identified. The alternating pressures corresponding to the critical speed are obtained from an FFT. A nonlinear damping model is estimated using Lazan’s hysteresis law; the equivalent viscous damping model is determined as

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8

Methel, Jeanne, William J. Gooding, John C. Fabian, Nicole L. Key, and Mark Whitlock. "The Development of a Low Specific Speed Centrifugal Compressor Research Facility." In ASME Turbo Expo 2016: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2016. http://dx.doi.org/10.1115/gt2016-56683.

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To achieve aggressive specific fuel consumption goals, aircraft engines are tending toward higher overall pressure ratios and higher bypass ratios for turbofans. As sizes decrease to meet these requirements, centrifugal compressors become a viable option as the last stage of the high pressure compressor. The last stages of an axial compressor in a small core engine face reduced efficiency due to the relatively large tip clearances with respect to blade height, and therefore, it may be more appropriate to finish the final compression stage with a low specific speed centrifugal compressor. A new

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9

Batailly, Alain, and Mathias Legrand. "Conjectural Bifurcation Analysis of an Aircraft Engine Blade Undergoing 3D Unilateral Contact Constraints." In ASME Turbo Expo 2014: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2014. http://dx.doi.org/10.1115/gt2014-25674.

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Prediction of rotor/stator interaction phenomena between a blade-tip and the surrounding abradable coating deposited on the casing has seen recent promising numerical developments that revealed consistency with several experimental set-up. In particular, the location of critical rotational frequencies, damaged blade areas as well as the wear pattern along the casing circumference were accurately predicted for an interaction scenario involving a low-pressure compressor blade and the surrounding abradable coating deposited on a perfectly rigid casing. The structural behaviour of the blade in the

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10

Rzadkowski, Romuald, and Artur Maurin. "Multi-Stage Coupled Forced Response of Aircraft Engine Compressor and Turbine Bladed Discs." In ASME Turbo Expo 2016: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2016. http://dx.doi.org/10.1115/gt2016-57643.

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Considered here is the effect of multistage coupling on the dynamics of an aircraft engine rotor with eight mistuned bladed discs on a drum-disc shaft during foreign object ingestion (FOI). In the dynamic model, each disc had a different number of rotor blades. Free and forced vibrations were examined using finite element models of single rotating blades, bladed discs and an entire rotor with bladed discs. Calculations of the mode shapes of flexible mistuned bladed disc-shaft assemblies took into account simultaneous excitations of the first and second stages of the compressor and the turbine

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