gas turbine cooling studies

Gas Turbine Cooling Studies

Prof. David G. Bogard

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Turbulence and Turbine Cooling Laboratory The University of Texas at Austin

Acknowledgements: Sponsors: DOE National Energy Technology Laboratory

Office of Naval Research Pratt & Whitney GE Aviation Siemens

Presentation at UT-CEM Industrial Advisory Panel Meeting

Nov. 14 , 2017

Outline of presentation

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n  Background on turbine cooling. n  Explanation of film effectiveness. n  Representative studies:

q  Internal cross-flow effects on film cooling

q  Coolant thermal field measurement and prediction

n  New high speed facility.

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Heat transfer and cooling is important throughout the gas turbine engine

Schematic from Bunker (ASME Turbo Expo 2013)

Focus of this presentation

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Cooling flows in the combustor and turbine section


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Film cooling history in gas turbines

Turbine airfoil cooling:

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Internal cooling

Coolant Flow

Mainstream Flow

Film cooling

Evaluation of the performance of film cooling

7

Tc

η, hf

φ

𝒒"= 𝒉↓𝒇 ( 𝑻↓𝒂𝒘 − 𝑻↓𝒔𝒖𝒓𝒇𝒂𝒄𝒆 )

𝒒"= 𝒉↓𝒊 ( 𝑻↓𝒄 − 𝑻↓𝒊𝒏𝒕 𝒔𝒖𝒓𝒇𝒂𝒄𝒆 )

𝝋≡ 𝑻↓∞ − 𝑻↓𝒔𝒖𝒓𝒇𝒂𝒄𝒆 /𝑻↓∞ − 𝑻↓𝒄    𝜼≡ 𝑻↓∞ − 𝑻↓𝒂𝒘 /𝑻↓∞ − 𝑻↓𝒄  

Overall effec*veness:

Film cooling effec*veness:

Performance parameters:

Film cooling holes in gas turbine engines are often fed by an internal crossflow

November 20, 2017 8

Figure from Han et al (2000)

Internal crossflow has been found to reduce film cooling effectiveness

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Crossflow Feed – VRc = 0.35 Quiescent Plenum Feed

x/d 0 5 10 15 20

x/d 0 5 10 15 20

Data is from Wilkes et al. (2016) and Anderson et al (2015)

Mainstream Flow

Coolant Jets

Ple

num

Fee

d

M =

2.0

0 pt

04

x/d

z/d

510

1520

2530

3540

4550

-12-9-6-3036912

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

η

Motivation and Objectives

November 20, 2017 10

Film cooling holes are commonly fed by an internal crossflow, but the impact of the crossflow velocity is poorly understood, particularly for shaped holes.

For this study, the performance of 7-7-7 shaped holes was studied with specific objectives of:

•  Understanding how crossflow velocity impacts film cooling effectiveness over a wide range of conditions

•  Determining which parameters govern how internal crossflow impacts film cooling performance

•  Providing insight into improved film cooling design

Low Speed Recirculating Wind Tunnel Facility

November 20, 2017 11

Shaped hole geometry

Engine condition coolant density ratios are achieved by using liquid nitrogen to cool film coolant feed to a density ratio of DR = 2.0

Internal cross-flow reduced film effectiveness, with larger decrease with increasing VRc

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Sensitivity to VRc was small at lower VR, but substantial for VR > 1.

Spatially average film effectiveness is the average of 4 pitches over x/d = 5-20

Internal cross-flow shown to cause substantial reduction in film effectiveness!

The reduction in film effectiveness appeared to be due to jet asymmetry within the coolant hole.

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x/d

VR = 0.56 VR = 1.11 VR = 1.67

VRc =

0.2

VR

c = 0

.4

VRc =

0.6

0 5 10 15 20 25 0 5 10 15 20 25 0 5 10 15 20 25

The jet trajectories deviated noticeably from the centerline of the hole

This jet movement was the result of biasing within the diffuser

Model turbine vane test facility at the University of Texas

n  Liquid nitrogen cooled secondary flow: DR = 1.5

n  Mainstream turbulence 20% n  Reynolds number Rc = 7x105

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Wind Tunnel

Turbulence Grid

Flow Bleed

Wax Sprayer Location

Flow Bleed

Adjustable Wall

Mainstream

Test Section

Internal coolant flow

Study to compare computational predictions with experimentally measured thermal fields for coolant jets

n  Measurements were taken in a corner test section with a simulated 3 vane linear cascade of an 11.6x scale C3X vane.

Schematic of vane model

n  Low (k = 0.043 W/mK) and high (k = 1.0 W/mK) thermal conductivity vane models were tested, the high thermal conductivity model was designed to match the Biot number of an actual engine vane

n  2-D thermal fields were measured 0, 5 and 10 hole diameters downstream of a single row of coolant holes on the suction side of the vane

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Comparison of experimental and computational thermal fields for M=0.65

n  CFD predicts a much colder core temperature n  CFD thermal field predicted the jet core would be separated from the vane wall. n  The conducting wall has only a small effect on the over-flowing coolant jets. n  Experimental coolant jet has spread across the entire pitch by x/d = 10, CFD

coolant jet spread only 75% of the pitch

1 2 3z/d

M = 0.65, I = 0.35, DR = 1.2

z/d

y/d

0 1 2 3 40

0.5

1

1.5

2

2.5

0.1

0.2

0.3

0.4

0.5

0.6

0.70.8

0.9

1 2 3z/d

M = 0.65, I = 0.35, DR = 1.2

z/d

y/d

0 1 2 3 40

0.5

1

1.5

2

2.5

0.1

0.2

0.3

0.4

0.5

0.6

0.70.8

0.9

M = 1.1, I = 1, DR = 1.2

z/d

y/d

0 1 2 3 40

0.5

1

1.5

2

2.5

3

3.5

0.1

0.2

0.3

0.4

0.5

0.6

0.70.8

0.9

CONDUCTING Experimental Computa1onal ϑ

(e)

(f)

(g)

(h)

x/d

=10

x/d

=5

1 2 3z/d

1

2

0

y/d

M = 0.65, I = 0.35, DR = 1.2

z/d

y/d

0 1 2 3 40

0.5

1

1.5

2

2.5

0.1

0.2

0.3

0.4

0.5

0.6

0.70.8

0.9

1 2 3z/d

1

2

0

y/d

ADIABATIC Experimental Computa1onal

(a)

(b)

(c)

(d)

M = 0.65, I = 0.35, DR = 1.2

z/d

y/d

0 1 2 3 40

0.5

1

1.5

2

2.5

0.1

0.2

0.3

0.4

0.5

0.6

0.70.8

0.9(b)

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Comparison of experimental and computational thermal fields for M=1.11

n  CFD predicts a much colder core temperature. n  Significant coolant jet separation seen both experimentally and computationally. n  The CFD shows a split core due to counter rotating vortices– this was not observed

in the experiments. n  Again, the thermal effects on the high conductivity wall were minimal on the coolant

jets above the wall.

x/d

=10

x/d

=5

M = 1.1, I = 1, DR = 1.2

z/d

y/d

0 1 2 3 40

0.5

1

1.5

2

2.5

3

3.5

0.1

0.2

0.3

0.4

0.5

0.6

0.70.8

0.9

1 2 3z/d

1

2

0

y/d

M = 1.1, I = 1, DR = 1.2

z/d

y/d

0 1 2 3 40

0.5

1

1.5

2

2.5

0.1

0.2

0.3

0.4

0.5

0.6

0.70.8

0.9

1

2

0

y/d

1 2 3z/d

ADIABATIC Experimental Computa1onal

(a)

(b)

(c)

(d)

M = 1.1, I = 1, DR = 1.2

z/d

y/d

0 1 2 3 40

0.5

1

1.5

2

2.5

3

3.5

0.1

0.2

0.3

0.4

0.5

0.6

0.70.8

0.9

ϑ M = 1.1, I = 1, DR = 1.2

z/d

y/d

0 1 2 3 40

0.5

1

1.5

2

2.5

0.1

0.2

0.3

0.4

0.5

0.6

0.70.8

0.9

M = 1.1, I = 1, DR = 1.2

z/d

y/d

0 1 2 3 40

0.5

1

1.5

2

2.5

0.1

0.2

0.3

0.4

0.5

0.6

0.70.8

0.9

1 2 3z/d

1 2 3z/d

Experimental Computa1onal CONDUCTING

(e)

(f)

(g)

(h)

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Schematic of Univ. of Texas high speed wind tunnel facility

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Two wind tunnel facilities: flat plate test section and cascade test section.

New facility will operate in the transonic Mach number, matching gas turbine engine operating conditions.

Conclusions

n  Our lab has extensive experience in accurate simulation of gas turbine operating conditions in a laboratory environment. These simulations allow us to test large scale models to provide detailed spatial resolution of coolant flows. This is key to understanding the physical mechanisms involved.

n  This allows development of improved turbine cooling technologies.

n  It also allows development of and evaluation of improved computational models which are needed for engine designs.

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gas turbine cooling studies

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