Shock Losses in Transonic Compressor Blade Rows

Miller, Genevieve R.; Lewis, George W.; Hartmann, Melvin J.

doi:10.1115/1.3673182

Cited by 71 publications

(37 citation statements)

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“…F,F rv ,F Qv ,F zv = blade force, viscous body force components per unit mass in the r, 6, and z directions H, h = total enthalpy, static enthalpy / =rothalpy defined by I=h + w 2 /2-(r 2 Q 2 )/2 ( = direction of computation in the meridional plane m = meridional streamline direction P 0 ,P s ,P t = reference static, and total pressures q = total absolute velocity defined by # = t v,v',v w, w', z X p radial, radius of curvature direction entropy total temperature, static temperature time absolute velocity, fluctuation, time average relative velocity, fluctuation, time average axial coordinate mean relative flow angle defined by blade lean angle in ?-0 plane defined by tane=-r(d8/de) angle between the direction of computation (0 and the radial direction, ratio of specific heats blade thickness blockage factor density Presented as Paper 79-1515 at the AIAA 12th Fluid and Plasma Dynamics Conference, Williamsburg, Va., July 23-25, 1979; submitted Aug. 14,1979; revision received Sept. 11,1980 T HE objective of the work being described here is to provide a logically consistent method for predicting the axisymmetric mean flow in turbomachinery using all available information about the blade-to-blade flow variations, which are regarded as a perturbation on this axisymmetric mean. This is a logical step beyond currently available design systems, such as that given in Ref.…”

Section: Nomenclaturementioning

confidence: 99%

Blade-to-Blade Flow Effects on Mean Flow in Transonic Compressors

Sehra

Kerrebrock

1981

AIAA Journal

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This paper describes a method of introducing nonaxisymmetric effects into the axisymmetric flowfield computation for a highly-loaded compressor. The approach is centered around modeling of three important nonaxisymmetric phenomena that control the mean momentum and energy transfer processes. Apparent stresses are introduced into the mean flow momentum equations by pitchwise averaging. The concept of apparent entropy resulting from the production of fluctuation energy is introduced. An expression for mean rothalpy variation along the streamline is derived by pitchwise averaging of the energy equation. Mean-flow equations suitable for the streamline-curvature computational scheme are developed which include these three effects. An assessment of the importance of these effects and experimental verifications of the computational approach have been achieved by use of time resolved experimental measurements of the flow in the MIT Slowdown Compressor. The revised streamline-curvature procedure has been used to predict the axisymmetric (peripheral mean) flow. The agreement of predicted and measured rotor outlet tangential velocity is excellent. The results show that three-dimensional inviscid effects do not modify significantly the mean flow, and that the direct effect of apparent stresses on the mean flow is also small. The major deviations of the actual flow from the inviscid flow are due to the variations of mean rothalpy and the production of apparent entropy, both effects which are introduced in turbomachine aerodynamics for the first time here. Nomenclature F,F rv ,F Qv ,F zv = blade force, viscous body force components per unit mass in the r, 6, and z directions H, h = total enthalpy, static enthalpy / =rothalpy defined by I=h + w 2 /2-(r 2 Q 2 )/2 ( = direction of computation in the meridional plane m = meridional streamline direction P 0 ,P s ,P t = reference static, and total pressures q = total absolute velocity defined by # = t v,v',v w, w', z X p radial, radius of curvature direction entropy total temperature, static temperature time absolute velocity, fluctuation, time average relative velocity, fluctuation, time average axial coordinate mean relative flow angle defined by blade lean angle in ?-0 plane defined by tane=-r(d8/de) angle between the direction of computation (0 and the radial direction, ratio of specific heats blade thickness blockage factor density Presented as Paper 79-1515 at

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Section: Nomenclaturementioning

confidence: 99%

Blade-to-Blade Flow Effects on Mean Flow in Transonic Compressors

Sehra

Kerrebrock

1981

AIAA Journal

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“…www.intechopen.com Shock correction applies only to supersonic blades and can be calculated reasonably well using the two-dimensional shock loss model provided by Miller et al, 1961, to which the reader is referred for details. Secondary losses are caused both by friction on the endwalls and secondary vorticity located on planes perpendicular to the shaft axis.…”

Section: Summary Of Compressor Losses At Design and Off-design Conditmentioning

confidence: 99%

“…A work by Schobeiri, 1996 confirms this. A step further is required when inlet supersonic flows have to be dealt with (as in the case of transonic stages), for which the predictive models given by Miller et al, 1961, is by far the most used.…”

Section: Aerodynamic Design Problem Formulationmentioning

confidence: 99%

Advances in Aerodynamic Design of Gas Turbines Compressors

Benini¹

2010

Gas Turbines

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“…Shock losses and Mach number effect is modelled using normal shock and isentropic relations developed by Miller et al (25) . The effects of secondary loss are considered in the last two terms of Eq.…”

Section: Validationmentioning

confidence: 99%

Performance Analysis of Axial Flow Compressor and Part Load Consideration in a Gas Turbine Application

Asli

Tousi

2013

JTST

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A performance analysis of a three stage axial flow compressor with part load consideration in gas turbine application regarding surge margin is presented in this paper. A three stage compressor was designed and numerical simulation as well as a modified stream line curvature analysis was applied to validate the design model. These two analyses verified the design which was based on a test case model and showed an acceptable accuracy. Performance was predicted by introducing a modified one dimensional method for the compressor. A typical gas turbine with specific components was considered to evaluate the effect of compressor early design parameters on its behavior at part loads. Four compressors were designed with change in reaction ratio and flow coefficient as the input data of the design process. The analysis showed that an increase of 8% and decrease of 5% in reaction ratio and flow coefficient respectively, compared with the base design model, would lead to getting away the running line from the surge line by an amount of 5% approximately and working in a safer margin as well as higher efficiency working range, so the energy conversion can be done efficiently.

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Shock Losses in Transonic Compressor Blade Rows

Cited by 71 publications

References 1 publication

Blade-to-Blade Flow Effects on Mean Flow in Transonic Compressors

Blade-to-Blade Flow Effects on Mean Flow in Transonic Compressors

Advances in Aerodynamic Design of Gas Turbines Compressors

Performance Analysis of Axial Flow Compressor and Part Load Consideration in a Gas Turbine Application

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