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TECHNICAL PAPERS

Improving Aerodynamic Matching of Axial Compressor Blading Using a Three-Dimensional Multistage Inverse Design Method

[+] Author and Article Information
M. P. C. van Rooij1

 Syracuse University, Syracuse, NY 13244

T. Q. Dang

 Syracuse University, Syracuse, NY 13244

L. M. Larosiliere2

U.S. Army Research Laboratory, NASA Glenn Research Center, Cleveland, OH 44135

1

Present address: Siemens Power Generation Industrial Applications.

2

Present address: Concepts NREC.

J. Turbomach 129(1), 108-118 (Feb 01, 2005) (11 pages) doi:10.1115/1.2372773 History: Received October 01, 2004; Revised February 01, 2005

Current turbomachinery design systems increasingly rely on multistage CFD as a means to diagnose designs and assess performance potential. However, design weaknesses attributed to improper stage matching are addressed using often ineffective strategies involving a costly iterative loop between blading modification, revision of design intent, and further evaluation of aerodynamic performance. A scheme is proposed herein which greatly simplifies the design point blade row matching process. It is based on a three-dimensional viscous inverse method that has been extended to allow blading analysis and design in a multi-blade row environment. For computational expediency, blade row coupling is achieved through an averaging-plane approximation. To limit computational time, the inverse method was parallelized. The proposed method allows improvement of design point blade row matching by direct regulation of the circulation capacity of the blading within a multistage environment. During the design calculation, blade shapes are adjusted to account for inflow and outflow conditions while producing a prescribed pressure loading. Thus, it is computationally ensured that the intended pressure-loading distribution is consistent with the derived blading geometry operating in a multiblade row environment that accounts for certain blade row interactions. The viability of the method is demonstrated in design exercises involving the rotors of a 2.5 stage, highly loaded compressor. Individually redesigned rotors display mismatching when run in the 2.5 stage, evident as a deviation from design intent. However, simultaneous redesign of the rotors in their multistage environment produces the design intent, indicating that aerodynamic matching has been achieved.

FIGURES IN THIS ARTICLE
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Copyright © 2007 by American Society of Mechanical Engineers
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Figures

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Figure 1

(a) Current design system and (b) proposed design system incorporating 3D inverse blading design

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Figure 2

Comparison between measured data and CFD simulations of NASA rotor 37 using INV3D and ADPAC : (a) performance map of total pressure ratio versus mass flow and (b) exit spanwise profile of total-to-total adiabatic efficiency at nominal operation

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Figure 3

Meridional flow path for the 2.5-stage compressor used in the design examples: Inlet guide vane (IGV), rotor 1 (R1), stator 1(S1), rotor 2 (R2) and stator 2 (S2); rotor 1: 26 blades, average solidity=2.16, aspect ratio=0.83; rotor 2: 56 blades, average solidity=1.96, aspect ratio=0.86. Also indicated, are multistage computational averaging planes.

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Figure 4

Diagnosis (multistage analysis) of flow structure induced by original geometry at design point in terms of contours of relative Mach number at 3%, 50%, and 100% (rotor blade tip) span. IGV and exit region are not shown for clarity.

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Figure 5

Predicted pressure-loading distribution for original R1 at design point: (a) multistage analysis of original design, (b) isolated analysis of original R1 with throughflow conditions, and (c) prescribed design intent for redesign exercise

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Figure 6

Predicted pressure-loading distribution for original R2 at design point: (a) multistage analysis of original design, (b) isolated analysis of original R2 with throughflow conditions, and (c) prescribed design intent for redesign exercise

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Figure 7

Isolated analysis of R1 and R2 (case A) at design point in terms of midspan contours of relative Mach number

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Figure 8

Diagnosis of flow structure induced by design case A geometry at design point in terms of contours of relative Mach number at 3%, 50%, and 100% (rotor blade tip) span. IGV and exit region are not shown for clarity.

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Figure 9

Diagnosis of flow structure induced by design case B geometry at design point in terms of contours of relative Mach number at 3%, 50%, and 100% (rotor blade tip) span. IGV and exit region are not shown for clarity.

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Figure 10

Multistage analysis of design speed throttling (varying backpressure, pb) for case A in terms of spanwise profiles of axisymmetric-averaged rotor exit cumulative total pressure

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Figure 11

Multistage analysis of design speed throttling for case B in terms of spanwise profiles of axisymmetric-averaged rotor exit cumulative total pressure

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Figure 12

Multistage analysis of design speed throttling for case A (left) and case B (right) in terms of relative Mach number contour at midspan for different backpressures

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Figure 13

Multistage analysis of design speed throttling for cases A and B. Shown are throttling characteristics of the respective rotors 1 and rotors 2 within the 2.5 stage.

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Figure 14

Multistage analysis of design speed throttling for cases A and B. Shown are the throttling characteristics of the complete 2.5 stages.

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