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TECHNICAL PAPERS

Effect of Blade Passage Surface Heat Extraction on Axial Compressor Performance

[+] Author and Article Information
P. N. Shah1

Gas Turbine Laboratory, Massachusetts Institute of Technology, Cambridge, MA 02139parthiv@alum.mit.edu

C. S. Tan

Gas Turbine Laboratory, Massachusetts Institute of Technology, Cambridge, MA 02139choon@mit.edu

1

Author to whom all correspondence should be addressed.

J. Turbomach 129(3), 457-467 (Feb 01, 2005) (11 pages) doi:10.1115/1.2372776 History: Received October 01, 2004; Revised February 01, 2005

Axial compressor performance with heat extraction via blade passage surfaces (compressor cooling) is compared to its adiabatic counterpart, using computational experiments and mean line modeling. For a multistage compressor with an adiabatic design point, results at fixed corrected rotor speed indicate that, if available, compressor cooling would (i) raise the overall pressure ratio (at a given corrected flow), (ii) raise the maximum mass flow capability, (iii) raise the efficiency, defined as the ratio of isentropic work for a given pressure ratio to actual shaft work, and (iv) provide rear stage choking relief at low corrected speed. In addition, it is found that, if available, cooling in the front stages is better than in the rear stages. This is primarily a thermodynamic effect that results from the fact that, for a given gas, the compression work required to achieve a given pressure ratio decreases as the gas becomes colder. Heat transfer considerations indicate that the engineering challenges lie in achieving high enough heat transfer rates to provide a significant impact to the compressor’s performance.

Copyright © 2007 by American Society of Mechanical Engineers
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References

Figures

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Figure 7

Eight stage compressor layout. The first stage is also analyzed as a single stage fan.

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Figure 8

Single stage compressor map, with and without cooling. The solid line is adiabatic; the dashed line is q*=−0.0025 per blade row.

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Figure 9

A single stage compressor map, with and without different effects from a cooled cascade performance. The solid line with circles is adiabatic; the dashed line with squares shows the cooling effects of a change in deviation only; a solid line with plusses shows cooling effects of change in ω only; a dashed line with triangles shows both effects. For cooled cases, q*=−0.0025 per blade row.

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Figure 10

Cooling schemes studied in eight stage compressor layout. q* values are per blade row.

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Figure 11

Eight stage efficiency map, with and without cooling. A dash-dotted line is adiabatic; a solid line is q*=−0.01 in the last two stages; a dashed line is q*=−0.0025 in all stages; a solid line with circles is q*=−0.01 in first two stages. q* values are per blade row.

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Figure 12

Eight stage compressor map, with and without cooling. The dash-dotted line is adiabatic; the solid line is q*=−0.01 in the last two stages; the dashed line is q*=−0.0025 in all stages; the solid line with circles is q*=−0.01 in first two stages. q* values are per blade row.

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Figure 13

Rotor 35 geometry and boundary conditions (single passage periodic) for nonadiabatic case. Filled static temperature contours on wall boundaries (hub, casing, and blade surfaces) show a BC of 100K on the blade and middle portion of the outer casing surface. (The positive x axis is in the downstream axial direction).

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Figure 15

Rotor 35 efficiency map

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Figure 16

Blade surface nondimensional heat transfer per unit solidity versus the wall temperature at various turbulent Reynolds numbers, generated using Reynolds analogy over a flat plate

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Figure 1

Selected flight path. The heavy, broken line represents low Mach number trajectory. The heavy, solid line is ρu2∕2=28.4kPa, representing a high Mach number trajectory. Faint broken lines are isocontours of ρu2∕2 in kPa. Faint dash-dotted lines are isocontours of specific energy, in km.

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Figure 2

Temperature, T versus entropy, s, for isentropic compression (A–B), isentropic compression with isobaric pre-cooling (A–C–D), and interspersed isobaric cooling and isentropic compression in N steps (A–E–⋯–F).

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Figure 3

Total pressure reduction coefficient, ω, versus the flow inlet angle at low (0.4) and high (0.8) inlet Mach number, for cascade 2. Solid lines are adiabatic wall BC; dashed lines are Twall∕Tt,in=0.5.

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Figure 4

Nondimensional cooling, q* versus the flow inlet angle at low (0.4) and high (0.8) inlet Mach number, for cascade 2, cooled case, Twall∕Tt,in=0.5

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Figure 5

Cascade 2, computed entropy generation versus incidence at low (0.4) and high (0.8) inlet Mach number using net entropy flux method (solid lines given by Eq. 6) and direct volume integration (dashed lines given by Eq. 7). Adiabatic and cooled (Twall∕Tt,in=0.5) cases shown.

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Figure 6

Cascade 2, viscous (black), and thermal (gray) dissipation versus incidence at low (0.4) and high (0.8) inlet Mach number. Plots on the left are adiabatic wall BC; plots on the right are Twall∕Tt,in=0.5.

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