A-1㻌感圧塗料による遷音速並進振動翼列の翼面上非定常圧力

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A-2

A Data Connection Method for Fluid-Structure Interaction Analysis of Turbine Blades

*Tadashi TANUMA (Teikyo Univ.), Hiroshi OKUDA, Gaku HASHIMOTO (Tokyo Univ.) and Kumi AKIYAMA (Office Kei)

ABSTRACT

Aerodynamic optimum designs of stator blades are already introduced in many designs of actual operating

commercial steam turbine units. However, aerodynamic optimum designs of rotating blades are still difficult due

to high centrifugal force and vibration stress on rotating blades. This paper presents a data connection method

for fluid-structure interaction analysis of turbine rotating blades to enhance FSI analysis for turbine designs.

Key words: Steam turbines, Long blades, Fluid-Structure Interaction Analysis, unsteady analysis

Fig.1 2) (CFD)

4)

FINETM/Turbo (NUMECA )

(CSD:

Computational Structural Dynamics) FrontISTR5)

FSI

Fig.1

2)3)

Fig.1 Flow chart of turbine blade FSI methodology

Geometry generation of deformed blades in operation

Steady FEA with centrifugal force

Aerodynamic AnalysisCFD

Structural AnalysisCSD (Computational Structural Dynamics)

CFD mesh generation

Unsteady 3D CFD of L-0 stator and rotating blades

Dynamic FEA for Detailed Model

Damping evaluation

Model steam turbine tests Wheel box tests

Interpolation of tractions on blade surfaces

Blade and stage geometry design CSD mesh

generation


7 / 298

Fig.1

Fig.2 Traction interpolation method

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

1.1

-0.6 -0.4 -0.2 0 0.2 0.4 0.6

Stat

ic p

ress

ure

rela

tive

x coordinate relative

Fig.3 CFD to FEA data transformation: Static Pressures

using Simple Method

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

1.1

-0.6 -0.4 -0.2 0 0.2 0.4 0.6

Stat

ic p

ress

ure

rela

tive


Fig.4 CFD to FEA data transformation: Static Pressures

using Modified Method

Pi

P i1

P i4

P i3

P i2

CFD mesh

CSD mesh

1l 2l

3l4l

nTP kikni , ikikikpT

n

ni

ni l

P

lP14

1

1


8 / 298

-1

-0.8

-0.6

-0.4

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

-0.6 -0.4 -0.2 0 0.2 0.4 0.6

Tra

ctio

n y

rela

tive


Fig.5 CFD to FEA data transformation: Traction y using

Modified Method

Fig.3

1) CFD

, 62, 9, pp.

22 -29, 2007.

Tanuma, T., Okuda, H., Hashimoto, G., Yamamoto, S.,

Shibukawa, Okuno, K., Saeki, H., Tsukuda, T.,

“Aerodynamic and Structural Numerical Investigation of

Unsteady Flow Effects on Last Stage Blades,” Proc. of

ASME Turbo Expo 2015, GT2015-43848, 2015.

3) Gaku Hashimoto, Hiroshi Okuda, Tadashi Tanuma,

Yoshinari Fukui, Kenichi Okuno, Tomohiko Tsukuda and

Naoki Shibukawa, 2014, Large-Scale One-Way FSI

Analysis for High-Performance Blade Design, JSME

Computational Mechanics Division Conference paper

No.7, 2014.

4) Yamamoto, S. and Daiguji H., "Higher-Order-Accurate

Upwind Schemes for Solving the Compressible Euler and

Navier-Stokes Equations", Computers and Fluids, 22-2/3,

pp. 259-270, 1993.

5) II

FrontSTR , 2008.


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Statistical Sensitivity Study of Mistuning Effect on Transonic Fan Flutter

*Atsushi TATEISHI, Toshinori WATANABE, Takehiro HIMENO (The Univ. of Tokyo), Mizuho AOTSUKA and Takeshi MUROOKA (IHI)

ABSTRACT

This paper presents a detailed investigation of the effect of mistuning on transonic fan flutter, which is

conducted as part of comprehensive efforts to find reasons for the mismatch in the flutter boundary in our past

numerical prediction. The flutter boundary of a full-annulus fan assembly with different mistuning levels is

statistically evaluated based on eigenvalue analysis and Monte-Carlo simulation. Nominal levels of mistuning

due to manufacturing tolerance have little effect to the flutter boundary because the decline in aerodynamic

damping is very steep. Therefore, the accuracy associated with the computational fluid dynamics is likely to

have caused the mismatch in the flutter boundary. Detailed observations of modal properties show that the mode

shape in flutter modes can be highly deviated from pure traveling wave mode, even if the level of mistuning is

nominal. For largely mistuned cases, highly-localized, single-blade dominant modes appear due to the

.

Key words: Fan, Aeroelasticity, Stall flutter, Mistuning, Monte-Carlo simulation

1-5)

6-7)

(1) 3-5,7-9)

(2) 8-9)

(3) 9)

(4) 10-13)

CFD

IHI

CEFS1 14)

(1F)


16 / 298

Fig. 1 Table 1 18

64

CEFS1 80%N 82.5 N

N 1F

Fig. 27)

75% 77.5%

Fig. 2 FSI Point

(1)

(2)

(3)

15)

(1)

(3)

16)

𝑗𝛿𝑗, 𝛿𝑀,𝑗

𝜀, 𝜀𝑀 ,

1F

𝒒 ̈+ (𝛬 + 𝛥𝛬)𝒒 = 𝑀−1(𝐴𝒒 + 𝐵𝒒)̇ (1)

𝒒 = [𝑞1 ⋯ 𝑞𝑁𝑏]𝑇 𝛬 = 𝜔02𝐼 ,

𝛥𝛬 = 𝜔02

⎣⎢⎡

2𝛿1 + 𝛿12 ⋱2𝛿𝑁𝑏 + 𝛿𝑁𝑏

2 ⎦⎥⎤,

𝑀 =⎣⎢⎡

1 + 𝛿𝑀,1 ⋱

1 + 𝛿𝑀,𝑁𝑏⎦⎥⎤

𝐴 =⎣⎢⎢⎡

𝑎1 𝑎2 ⋯ 𝑎𝑁𝑏𝑎𝑁𝑏 𝑎1 ⋯ 𝑎𝑁𝑏−1⋮ ⋱ ⋮

𝑎2 ⋯ 𝑎𝑁𝑏 𝑎1 ⎦⎥⎥⎤

, 𝐵 =⎣⎢⎢⎡ 𝑏1 𝑏2 ⋯ 𝑏𝑁𝑏𝑏𝑁𝑏 𝑏1 ⋯ 𝑏𝑁𝑏−1

⋮ ⋱ ⋮𝑏2 ⋯ 𝑏𝑁𝑏 𝑏1 ⎦

⎥⎥⎤

Eq. (1)

𝑁𝑏 = 18𝛬, 𝐴, 𝐵𝜔0

𝛥𝛬 𝑀

1

𝜆,𝝍

Fig. 2 Mismatch of flutter boundary on low shaft speeds in our past numerical prediction 7)

Fig. 1 CEFS1 fan

Table 1 Specifications of CEFS1 fan rotor Number of blades 18

Aspect ratio 1.6

Relative Mach number at tip 1.4

Blade material Ti-6Al-4V

Range of flutter region and structural mode shape

80%N, 82.5%N(1F mode)

(a) Full-annulus view(b) 1F mode shape

(80% speed)


17 / 298

𝜆 [ 𝝍𝜆𝝍] = [ 𝑂 𝐼

−𝛬 − 𝛥𝛬 + 𝑀−1𝐴 𝑀−1𝐵] [ 𝝍𝜆𝝍] (2)

𝜆 𝜇AE𝑓AE Eq. (3) 𝝍

𝜇AE = −Re(𝜆) 𝑓AE = |Im(𝜆)|/2𝜋𝜇AE > 0

𝜇AE < 0

𝛬 𝐴, 𝐵

7)

(MC)

Eq. (2)

MC

𝑁fan = 10000

Eq. (4)

1F

𝜇min

𝜇min = 1𝑁fan

∑ min(𝜇AE,1, ⋯ , 𝜇AE,𝑁𝑏)𝑁fan𝑖=1 𝑖

𝜇min

𝜇min = 0𝜇min

”frequency mistuning”

May 17), Beirow 18), Schnell 19)

Table 2

1F

May 22

5

0.47 0.17%

0.2

May 17)

17

Cov(𝛿, 𝛿𝑀)/√𝑉 (𝛿)𝑉 (𝛿𝑀) = −0.25

Table 3, Table 4

1F

𝜀 = 0.2, 0.5%𝜀 = 1.0, 1.5%

May

𝜀𝑀 = 15%, 𝜀𝑀 = 30%

3

Table 2 Standard deviation of blade-alone frequency in past studies (1F mode)

Author Model STDEV [%]

May 17) UHBR Fan 0.47 (all)

0.17 (clean)

Beirow 18) HPC blisk (1st stage) 0.13

Schnell 19) Counter-rotating fan 0.25

Table 3 Test cases for frequency mistuningNo. STDEV[%] Assumed situation 1 𝜀 = 0.2% Nominal level 2 𝜀 = 0.5% Nominal level 3 𝜀 = 1.0% Large mistuning 4 𝜀 = 1.5% Very large mistuning

Table 4 Test cases for mass mistuning

No. STDEV[%] Assumed situation 1 𝜀𝑀 = 15% Nominal level 2 𝜀𝑀 = 30% Large deviation


18 / 298

Fig. 3

𝜀𝑀 , 𝜀 1F

𝜇min80% C Fig. 3 (a)

𝜀𝑀 𝜀

Fig. 3 (b) 𝜀 𝜇min

𝜀𝑀 = 0

Fig. 4

𝜀 = 0.2%

Fig. 4 Change in averaged minimum damping rate due to frequency mistuning on each rotational speed. The “F” and “S” show experimental stability limits due to surge and flutter, respectively.

Fig. 3 Effects of mass and frequency mistuning on the statistics of damping rate in the flutter mode (80%N, OP: C)(a) Cumulative probability (b) Change in averaged minimum damping rate

μ

[1/s

]m

in

μ

[1/s

]m

in

Fig. 5 Flutter boundary obtained from averaged minimum damping rate


19 / 298

72.5%

75.0%

~ 82.5%

𝜀 =0.5% 𝜀 = 1.0%

75%

Fig. 5 Fig. 4

𝜀 = 1.0%50%

𝜀 = 1.0%75.0%,

77.5% B

𝜇min = 00.925, 0.862

75.0%,

77.5%

80% C

Fig. 6

Traveling Wave Mode (TWM)

𝜀 = 0.2%

𝜀 = 0.5%

𝜀 = 1.0, 1.5%

Fig. 6 Histogram of aeroelastic eigenvalues in the 1F mode family (80%N, OP: C)

(a) 𝜀 = 0.2% (b) 𝜀 = 0.5% (c) 𝜀 = 1.0% (d) 𝜀 = 1.5%

230

Stable

Unstable

Merging modesMerged

Single-bladedominant mode

0

20

40

60

80

240 250 260 230 240 250 260 230 240 250 260 230 240 250 260

Single-bladedominant mode

f [Hz]

μ [

1/s

]

modes

AE

AEf [Hz]

AE f [Hz]AE

f [Hz]AE


20 / 298

𝜇AE = 25 [1/s]

1

Eq. (1) 𝑞10

(Single-Blade Dominant Mode,

SBDM)

Fig. 7

Im(𝜆) > 014 1

2

0

𝜀 = 1.0, 1.5%

”mode localization” 12)

𝜀 = 1.5% 2 SBDM

mode localization

TWM TWM

𝒔 TWM 𝑊H

𝒒 = 𝑊𝒔 𝒔 = [𝑠1 ⋯ 𝑠𝑁𝑏]𝑇𝑊𝑗𝑘 = exp {2π𝑖(𝑗−1)(𝑘−1)

𝑁𝑏 } 𝑊𝑗𝑘−1 = 𝑊𝑗𝑘𝐻/𝑁𝑏

Fig. 8 𝜀 = 0.5%

Fig. 7 Examples of aeroelastic mode shapes in the 1F mode family with different mistuning levels (14 modes from the lowest damping, real part, only aeroelastic modes for Im(𝜆) > 0 are shown)

Fig. 8 Typical aeroelastice modeshape of the flutter mode (80%N, OP: C, 𝜀 = 0.5%),

Solid bars: fan No. 57, thin lines: other 20 fans


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2

𝜃𝑖+1 − 𝜃𝑖 |𝑞𝑖|TWM |𝑠𝑖|

20

Fig. 8 -10 80 [deg]

TWM |𝑠i|40 [deg]

TWM

40 60 [deg]

TWM

Fig. 9 𝜀 = 1.5%Fig. 7

4~10

1~3 10~11

4~10

TWM 40 [deg]

TWM 𝜀 =0.5%

SBDM

Fig. 10 11

TWM

(1)

(2)

(3)

(4)

Fig. 10 Typical aeroelastice modeshape of the single-blade dominant mode (80%N, OP: C, 𝜀 = 1.5%),

solid bars: fan No. 720, thin lines: other 20 fans

Fig. 9 Typical aeroelastice modeshape of the flutter mode (80%N, OP: C, 𝜀 = 1.5%),

solid bars: fan No. 720, thin lines: other 20 fans


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JSPS 14J10312

IHI CEFS1

1) Platzer, M. F. and Carta, F. O. E., “AGARD Manual on Aeroelasticity in Axial-Flow Turbomachines Volume 1: Unsteady Turbomachinery Aerodynamics”, AGARDo- graph No. 298, Vol. 1, 1988

2) Jeffers, J. D. and Meece, C. E., “F100 Fan Stall Flutter Problem Review and Solution”, Journal of Aircraft Vol. 12, No. 4, pp.350-357, Apl. 1975

3) Stargardter, H., “Subsonic/Transonic Stall Flutter Study Final Report”, NASA-CR-165256, Pratt and Whitney Aircraft Group, Jun, 1979

4) Isomura, K. and Giles, M. B., “A Numerical Study ofFlutter in a Transonic Fan”, Journal of Turbomachinery, No. 120 (3), pp.500-507, Jul. 1998

5) Kontos, K., Weir, D., and Ross, D., “Quiet High Speed Fan II (QHSF II): Final Report”. NASA CR-2012- 217451, Aug. 2012

6) “”,

Vol. 44, No. 4, pp.282-291, Jul. 2016 7) ”

FSI ” Vol. 44, No. 4, pp.292-301, Jul. 2016

8) Vahdati, M., Smith, N. H. S., and Zhao, F., “Influence of Intake on Fan Blade Flutter”, Journal of Turbomachinery, No. 137 (8), 081002, Aug. 2015

9) Vahdati, M. and Cumpsty, N., “Aeroelastic Instability in Transonic Fans”, Journal of Engineering for Gas Turbine and Power, Vol. 138 (2), 022604, Sep. 2015

10) Pierre, C., and Murthy, D. V., “Aeroelastic Modal Characteristic of Mistuned Blade Assemblies: Mode Localization and Loss of Eigenstructure”, AIAA Journal, Vol. 30, No. 10, pp. 2483-2496, Oct. 1992

11) Srinivasan, A. V., “Flutter and Resonant Vibration Characteristics of Engine Blades”, Journal of Engineering for Gas Turbine and Power, Vol. 119 (4), pp.742-775, Oct. 1997

12) Castanier, M. P. and Pierre, C., “Modeling and Analysis of Mistuned Bladed Disk Vibration: Status and Emerging Directions”, Journal of Propulsion and Power, Vol. 22, No. 2, pp. 384-396, Mar.-Apl. 2006

13) Salles, L. and Vahdati, M., “Comparison of Two Numerical Algorithms for Computing the Effects of Mistuning of Fan Flutter”, ASME Paper GT2016-57324, 2016

14) Murooka, T., Goto, S., Mizuta, I., and Kodama, H. “New

Concept Design and Development of an Advanced Transonic Fan Rotor”, IGTC2007 Tokyo, TS-053, 2007

15) Kielb, R. E., Feiner, D. M., Griffin, J. H., and Miyakozawa, T., “Flutter of Mistuned Bladed Disks and Blisks with Aerodynamic and FMM Structural Coupling”, ASME Paper GT2004-54315, 2004

16) Kielb, R. E., Hall, K. C., Hong, E., and Pai, S. S., Probabilistic Flutter Analysis of a Mistuned Bladed

Disk”, ASME Paper GT2006-90847, 2006 17) May, M., “Model Updating for the Aeroelastic ROM of a

Modern Blisk”, Proceedings of the 13th International Symposium on Unsteady Aerodynamics, Aeroacoustics and Aeroelasticity of Turbomachines, I13-S13-7, 2012

18) Beirow, B., Giersch, T., Kuhhorn, A., and Nipkau, J., Forced Response Analysis of a Mistuned Compressor

Blisk”, Journal of Engineering for Gas Turbine and Power, Vol. 136 (6), 062507, Jun. 2014

19) Schnell, R., Lengyel-Kampmann, T., and Nicke, E., “On the Impact of Geometric Variability on Fan Aerodynamic Performance, Unsteady Blade Row Interaction, and Its Mechanical Characteristics”, Journal of Turbomachinery, Vol. 136 (9), 091005, Sep. 2014


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A-6

Effect of Tip Clearance Gap and Wall Boundary Layer Thickness to Performance of Transonic Compressor

*Yasunori SAKUMA, Toshinori WATANABE and Takehiro HIMENO (Univ. of Tokyo)

ABSTRACT

Numerical analysis has been conducted on transonic compressor NASA Rotor 37 with 9 different tip

clearance gaps while varying the wall boundary layer thickness of the incoming flow. Overall performance,

basic flow characteristics, and near casing phenomena has been carefully observed. According to the result, it

has been shown that NASA Rotor 37 can be lead to numerical instability through two different mechanisms

depending on the size of the tip clearance gap and the wall boundary layer thickness. Stall margin became the

maximum with the tip clearance gap at which the stall mechanism switched from one to another.

Key words: Compressor, Tip Clearance, Stall Inception, Numerical Analysis

hbase

h

HLE ctip

h / HLE


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k-ω

2.106

1.48

17,188

0.356

252

36

Total pressure ratio

Relative Mach number at tip

Rotation speed [rpm]

Tip clearance gap [mm]

Blade tip radius [mm]

Number of rotor blades

h

h [mm] h / ctip [%]h / hbase h / HLE [%]

0.120

0.240

0.361

0.481

0.721

0.961

1.202

1.442

0.000

0.158

0.317

0.475

0.633

0.950

1.266

1.583

1.900

0.000

0.25

0.50

0.75

1.00

1.50

2.00

2.50

3.00

0.00

0.089

0.178

0.267

0.356

0.534

0.712

0.890

1.068

0.000

025CL

000CL

050CL

075CL

100CL

150CL

200CL

250CL

300CL

Station 1

Leading Edge

Shroud

Hub

Station 2Station 1a Station 3 Station 4Station 4a

Trailing Edge

30% span

50% span

70% span

90% span

95% span

Laser Anemometer

Blade-to-blade

Survey Planes

Aero Performance DataLaser Anemometer Spanwise SurveyAero Performance Data

Axial Position, x

-4.19 -5% chord 10% chord 4.57 10.16 10.67x [cm]

1 1a 2 3 4a 4Station

* x = 0 at intersection of hub and leading edge

Rad

ial

Po

siti

on

, r


30 / 298

1≤+y

1

4

t

t

P

P=π ( )1

1

t1t4

1

14

−

−=

−

TT

PP tt γγ

η

1tP 1tT 4tP 4tT

corrm designm

reft

reft

corrPP

TTmm

1

1=

Pref = 101,325 [Pa] Tref = 288.15 [K]

1tP

1tP

SM

Outlet

InletFlow

0.995

0.986

1.000

1BL

0BL

2BL

reft PP1Mass averaged total pressure

−=designcorr

stallcorr

stall

design

m

mSM

_

_1π

π

stall design


31 / 298


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LE

TE

LE

TE

LE

TE

LE

TE

M

1.9

1.8

1.7

1.6

1.5

1.4

1.3

1.2

1.1

1

0.9

0.8

0.7

0.6

0.5

0.4

0.3

0.2

0.1

LE

TE

LE

TE

LE

TE

LE

TE

M

1.9

1.8

1.7

1.6

1.5

1.4

1.3

1.2

1.1

1

0.9

0.8

0.7

0.6

0.5

0.4

0.3

0.2

0.1

LE

TE

LE

TE

LE

TE

LE

TE

M

1.9

1.8

1.7

1.6

1.5

1.4

1.3

1.2

1.1

1

0.9

0.8

0.7

0.6

0.5

0.4

0.3

0.2

0.1

TCm

m


33 / 298

LE

TE

LE

TE

LE

TE

xτ

60

57

54

51

48

45

42

39

36

33

30

27

24

21

18

15

12

9

6

3

TC

CL

TC

TC

AA

mm

100

_⋅

&&


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A-12

Behaviors of Tip Leakage Flows in Transonic Centrifugal Compressor with Splitter Blade at Low Flow Rate Condition

*Masanao KANEKO(Hosei Univ.) and Hoshio TSUJITA(Hosei Univ.)

ABSTRACT

In this study, the individual influence of the tip leakage flows from the main and the splitter blades on the flow

behavior and the loss generation in the transonic centrifugal compressor at low flow rate was investigated by

performing the numerical analysis at four calculation conditions, which were prescribed by the presence or the

absence of the tip clearances of the main and the splitter blades. The computed results including those for the

design condition clarified that the tip leakage vortex from the leading edge of the main blade increased the

incidence angle to the splitter blade, and consequently increased the loss. On the other hand, the presence of the

tip clearance of the splitter blade decreased the loss by reducing the incidence angle to the splitter blade. The

former influence was enhanced but the latter was reduced by the decrease of the flow rate.

Key words: Centrifugal compressor, Loss generation, Shock wave, Tip leakage flow, Transonic compressor

1

Krain1)

50,000 rpm

1.7 6.1

2.55 kg/s

Shroud

(TC TC-M TC-S NTC)

TC

0.5 mm 0.3 mm TC-M

TC-S

Shroud

NTC

CFD STAR-CD Ver.4.18

SIMPLE

Table 1 Dimensions of test impeller

Inlet hub diameter of main blade 1m,h (mm) 60

Inlet tip diameter of main blade 1m,t (mm) 156

Outlet diameter 2 (mm) 224

Number of blades / (Main/Splitter) 13/13

Inlet blade angle at main blade tip 1m,t (deg.) 63.5

Outlet blade angle 2 (deg.) 38

Outlet blade height 2 (mm) 10.2


55 / 298

MARS(Monotone Advection and Reconstruction

Scheme)

Spalart-Allmaras

1.0×10-4

Tip 0.64

/ 2=1.89

1

1

380

39 37 +

1.0

TC

Shroud

50,000 rpm

2 80% span

=1.0

(1)

3

4

5

* *=1.3

*

(2)

(TC TC-M

TC-S)

6 5

7

TC

TC2)

2(a) 3(a)

(SWm1)

4(a)

SWm1 / 0.1

5(a)

(Vt,m-ss1)

6(a) SWm1 Vt,m-ss1

7(a)

Shroud SWm1

(LMR1)

Vt,m-ss1

7(a)

(Vt,s-ss1) 4(b) 5(a)

Vt,s-ss1

4(a)

/ 0.1

Main blade

Splitter blade

2

250

1

1

1

Fig. 1 Computational grid


56 / 298

TC 3

TC 3 SWm1

2 3 2

7

(LMR2) TC TC-M 80%

span

TC-S NTC 80% span

3 TC TC-M

SWm1

TC-S

NTC

2 (TC TC-M)

LMR 2

SWm1

LMR2=1.0

SWm1 SWm1 SWm1 SWm1

SWs1

LMR2 LMR2 LMR2

SWm1 SWm1 SWm1

SWs1

SWm1

(a) TC (b) TC-M (c) TC-S (d) NTC Fig. 2 Relative Mach number and static pressure coefficient distributions on blade-to-blade surface at 80% span

( =0.05)

( =0.01)

Rot

atio

n

Axial

(a) TC (b) TC-M (c) TC-S

(a) Main blade (b) Splitter blade

Fig. 3 Limited streamline and static pressure coefficient distribution on blade surfaces

Fig. 4 Streamwise distribution of blade loading coefficient near blade tip

-0.50

0.00

0.50

1.00

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80 0.90 1.00

LE TE

-0.50

0.00

0.50

1.00

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80 0.90 1.00

LE TETC

TC-M

TC-S

NTC

(d) NTC


57 / 298

SWm1

LMR 2

7

TC

TC-M NTC TC-S

TC

TC-M NTC TC-S

4(b) TC-M TC

TC

TC-S TC

Shroud 7

TC-S LMR2

TC LMR2 Vt,m-ss1

TC-S

NTC TC

NTC Shroud LMR2

Vt,m-ss1VVVVVVVVVVVVVVVV

Vt,s-ss1 Splitter blade

Main blade

Vt,m-ps1

Vt,m-ss1VVVVVVVVVVVVVVVVV

Vt,s-ss1

(a) TC

(d) NTC Fig. 5 Entropy function * distribution and behaviors of tip leakage flows

(c) TC-S

(b) TC-M

Splitter blade

Main blade

Vt,m-ss1VVVVVVVVVVVVVVVVVVVVVVV

Vt,s-ss1

e

Vt,m-ps1

Vt,m-ss1VVVVVVVVVVVVVVVVVVVVV

Vt,s-ss1

(a) TC

(b) TC-M

(c) TC-S

[1/s]

Fig. 6 Behaviors of tip leakage vortices and static pressure coefficient distribution on blade-to-blade surface near blade tip


58 / 298

TC LMR2 Vt,m-ss1

TC

Vt,m-ss1

Vt,m-ss1

TC TC-M

NTC TC-S

7

2 TC

TC-M TC

7(a) (b) TC-M

5(a) (b) TC Vt,s-ss1

TC-S TC

2(a) (c) 7(a)

(c) TC TC-S

NTC TC

Vt,s-ss1

(SWs1) 2

7 (a) (d)

Shroud

3(d)

NTC SWs1

TC

TC-M NTC TC-S

4(b) 7

5(a) (c) Vt,s-ss1 TC-S

TC Vt,m-ss1

Vt,m-ss1 Vt,m-ss1

4(a) TC

SWm1

2 4(a)

TC

TC-M NTC

TC-M

(Vt,m-ps1)

5(b)

TC-M Vt,m-ps1

5

SWm1

=1.0LMR2

LMR1

( =0.05)

Rot

atio

n

Axial

Fig. 7 Relative Mach number distribution on blade-to-blade surface near blade tip and streamline and

static pressure coefficient distribution around leading edge of splitter blade

LMR2

LMR1

LMR2 LMR2

(a) TC (b) TC-M (c) TC-S (d) NTC


59 / 298

2 7 LMR 2 TC

TC-M SWm1 TC-S NTC

Shroud

3 5(a) (b)

TC TC-M Shroud

Vt,m-ss1

SWm1 5

6 (a) (b)

Vt,m-ss1

5

5 7 TC Vt,s-ss1

5(a)

TC-M 5(b) Vt,m-ps1

TC-S TC Vt,s-ss1

5(c) TC Vt,m-ss1

4(b)

Vt,s-ss1

Vt,m-ss1 Vt,s-ss1

NTC Shroud SWs1

3(d) 5(d)

5

SWm1

Shroud

Vt,m-ss1

SWm1

Vt,m-ss1

Vt,s-ss1

Vt,m-ss1

Vt,m-ss1

Vt,s-ss1

SWs1 Vt,m-ps1

Shroud

Eisenlohr, G., Krain, H., Richter, FA., Valentin, T.: Investigations of the Flow Through a High Pressure Ratio Centrifugal Impeller, Proceedings of ASME Turbo Expo 2002, Amsterdam, the Netherlands, GT-2002-30394 (2002).Kaneko, M., Tsujita, H.: Influences of Tip Leakage Flows on Flow Behavior in Transonic Centrifugal Compressor with Splitter Blade at Design Condition, IGTC 2015, Tokyo, Japan, 174 (2015).


60 / 298

A-13

Effects of Blade Tip Clearance on Hub-Corner Separation in a Stator Cascade of a Transonic Axial Compressor

* Seishiro SAITO (Kyushu Univ.), Yuki TAMURA (Kyushu Univ.), Masato FURUKAWA (Kyushu Univ.), Kazutoyo YAMADA (Kyushu Univ.),

Akinori MATUOKA (KHI) and Naoyuki NIWA (KHI)

ABSTRACT

The paper presents the effects of blade tip clearance on hub-corner separation in a stator cascade of a multi-

stage transonic axial compressor. In this study, the effects have been numerically investigated for three clearance

cases: uniform clearance, partial clearance, and no clearance. The numerical analysis shows that the leakage vortex

due to the stator blade tip clearance is the main cause for loss near the hub in the case of uniform clearance. In the

case of no clearance, low-energy fluid of the secondary flow on the hub accumulates on the suction side of the

blade, producing the separation vortex near the hub corner. In the partial clearance case, the separation vortex

occurring near the filled clearance section and the leakage vortex from the rear tip clearance produce a huge loss,

generating very complicated flow field.

Key words: Transonic Axial Compressor, RANS, Visual Data Mining, Blade Tip Clearance, Hub Corner

Separation

1

(VSV)

Yamada (1)

(2)

(3)

(4)

2

3

RANS(5)

2

2


61 / 298

20

30 104

1

AutoGrid5 ver9.10

NUMECA

H-J-O

1

520

2 410 2

340

y+<1

3 2

3

Navier-Stokes

k- (6)

RANS

cell-centered

SHUS Simple High-resolution Upwind

Scheme (7) MUSCL

Gauss

(a) Leading edge side

(b) Trailing edge side

Fig. 1 Closeup view of partial clearance

Fig. 2 Shape of blade tip clearance

Fig. 3 Spanwise distributions of

total pressure loss coefficient at 1st stator exit

Fig. 4 Spanwise distributions of axial velocity at 1st

stator exit


62 / 298

Euler

MFGS Matrix Free Gauss-Seidel

3 4

(1)

Pt1 Pt2

Pt0 ps0

50%

3 30% 50%

30%

15%

3%

4

15%

5%

5%

10%

10% 20%

30%

2

10% 30%

30%

5%

Fig. 5 Limiting streamlines on suction surface and vortex cores in the 1st stator (in uniform clearance case)

Fig. 6 Distribution of entropy function

(in uniform clearance case)


63 / 298

5

(5)

(2)

u

5

30%

5

2

6

1

0 0

t t

t t

T PS

T P (3)

Tt, Pt

Tt0, Pt0

5

6 30%

3

3%

10% 50%

50%

Fig. 7 Limiting streamlines on suction surface and vortex cores in the 1st stator (in no clearance case)


(in no clearance case)


64 / 298

7

20%

7

8

10 30%

Fig. 9 Limiting streamlines on suction surface and vortex cores in the 1st stator (in partial clearance case)

Fig. 10 Limiting streamlines on suction surface of 1st stator (in partial clearance case)


(in partial clearance case)


65 / 298

50%

3 10% 20%

7

8 50%

9

10

11

2

10%

30%

30%

8

30%

30%

70% 90%

5 9

11 50%

90%

3 RANS

(1)


66 / 298

(2)

(3)

(2)

30

(1) K.Yamada, et al., Large-scale DES Analysis of Stall Inception Process in a Multi-Stage Axial Flow Compressor, ASME PaperNo. GT2016-57104 (2016).

(2) Gbadebo, S. A., Cumpsty N. A. and Hynes, T. P., Three-Dimensional Separations in Axial Compressors, ASME Journal of Turbomachinery, Vol. 127 (2005), pp. 331-339.

(3) Gbadebo, S. A., Cumpsty N. A. and Hynes, T. P., Interaction of Tip Clearance Flow and Three-Dimensional Separations in Axial Compressors, ASME Journal of Turbomachinery, Vol. 129 (2007), pp. 679-685.

(4) Lei, V. M., Spakovszky, Z. S., and Greitzer, E. M., A Criterion for Axial Compressor Hub-Corner Stall, ASME Journal of Turbomachinery, 130, (2008), pp. 031006 10.

(5) 23 91 (2003),

206-213. (6) Wilcox, D. C., Simulation of Transition with a Two-

Equation Turbulence Model, AIAA Journal, Vol.32, No.2 (1994), pp.247-255.

(7) Shima, E. and Jounouchi, T., Role of CFD in Aeronautical Engineering (No.14) -AUSM Type Upwind Schemes, the 14th NAL Symposium on Aircraft Computational Aerodynamics, NAL SP-34 (1997), pp.7-12.


67 / 298

A-14

Verification of Control Effect of Tip Leakage Flow in Centrifugal Compressor by Ring-type Plasma Actuators

* Toshiyuki OHASHI, Hitoshi ISHIKAWA, Takehiko SEGAWA and Takayuki MATSUNUMA

ABSTRACT

Velocity and pressure distributions near the outlet in a centrifugal compressor system were measured by PIV

and pressure scanner for evaluating control effects of the tip leakage flow by means of a ring-type plasma

actuator. A copper wire coated with silicone was embedded in a casing surface facing tips of the impeller. Under

a rotating speed of impeller = 1,500 rpm, dielectric barrier discharge (DBD) plasma was formed in between

their clearances (s = 1.2 mm) by applying the fixed peak-to-peak voltage Vp-p = 11.7 kV with frequencies fp = 5,

10, 20, 30 and 40 kHz. In the case of fp = 5 kHz, velocity in the outside of impeller hub increased in comparison

with that without flow control, although it tends to decrease according to higher fp. Static pressure near the hub

side of diffuser was also enlarged further negative value for the particular case of fp = 5 kHz.

Key words: Tip leakage flow, Centrifugal compressor, Ring-type plasma actuator, Active flow control, PIV

(1),(2)

PA(3),(4)

PA

(5),(6)

PA (7)

PA 1

DBD

2 3

PIV

PA(8)

3

2

Shaft (Metal)

DBD Plasma

Tip Leakage Flow

Tip Clearance

Metal Wire Coated with Insulation Material

Casing

Centrifugal Impeller(Aluminum Alloy)

Main Flow

Power Supply

RotationGND

Fig. 1 Schematic drawing of ring-type plasma

actuator for centrifugal compressor


68 / 298

3

PA 1

2,000 cc

d1 = 42.4 mm

d2 = 59 mm

7,760

rpm DC m-link , CPH50

D1 = 44.4 mm

h2 = 5.2 mm 3

c = 1 mm

0.1 mm

2 mm

1.5 mm

1.3 mm 0.45 mm

PSI , PG1040F

DBD

2

65 mm

95 mm 125 mm 4

12 Esterline NetScanner

9116 500 Hz

PA peak-to-peak Vp-p

fp Vp-p fp

Table. 1 Specifications of experimental apparatus

Impeller

Number of blade (n) 12 blade Inlet diameter (d1) 42.4 mm Outlet diameter (d2) 59 mm Outlet blade height (h1) 5.0 mm

Casing Suction diameter (D1) 44.4 mm Diffuser height (h2) 5.2 mm Groove depth for PA 1.5 mm

PA

Tip clearance (s) 1.2 mm

Wire outer diameter 1.3 mm Thickness of insulator 0.45 mm

(b) Cross section

65 mm

95 mm

Pressure Scanner

CCDCamera

Double PulseNd:YAG Laser

Static Pressure Hole 12

mm)

Suction

125 mm

PA

HV-RFPower Supply

FlowVisualization

(PIV)

r

z

(a) Top view

Impeller

h2

Fig. 3 Experimental setup

x

y

PA off PA on

Fig. 2 Main result in the previous study(8)


69 / 298

Vp-p

(8) Vp-p

= 11.7 kV = 1,500 rpm

PA

s

= 1.2 mm 4 fp = 5

kHz

PA Vp-p = 11.7 kV

fp = 5 kHz 40 kHz

PIV

r = 30 mm 60 mm

h = 5.2 mm 2

Nd:YAG

Litron Lasers , NANO S 30-30 PIV, 30 mJ/pulse

3 PA

Flame Straddle 205 410

3.75 Hz PIV

205

5 5(a) PA

5(b) fp = 5 kHz PA

5(c) (f)

5

r = 33.9 mm

u 6 fp = 10, 20, 30,

40 kHz PA

fp = 40 kHz

14.9%

fp = 5 kHz

5.1%

PIV

500 Hz

90 s fp = 0 PA , 5, 10, 20, 30,

40 kHz

7

65 mm

95 mm 125 mm

PA

fp = 5 kHz

65 mm 95mm

-6000

-4000

-2000

0

2000

4000

6000

0 50 100 150 200

V [

V]

t [ s](b) 20 kHz

-6000

-4000

-2000

0

2000

4000

6000

0 50 100 150 200

V [

V]

t [ s](a) 5 kHz

-6000

-4000

-2000

0

2000

4000

6000

0 50 100 150 200

V [

V]

t [ s](c) 40 kHz

Fig. 4 Voltage form variations


70 / 298

PA

PA

P 8(a) PA

Cr = 22

nF ,V-Q

CPA

VPA

QPA

QPA = Qr CPAVPA =

CrVr QPA Cr

VPA VPA QPA

x y

8(a)

1

PA

9 PIV

fp

x mm

ym

m

510152025308

10

12

z m

m

2

4

35 40 45 50 55r mm

0

(a) PA Off

x mm

ym

m

510152025308

10

12

z m

m

2

4

35 40 45 50 55r mm

0

(f) 40 kHz

x mm

ym

m

510152025308

10

12

z m

m

0

2

4

35 40 45 50 55r mm

(e) 30 kHz

x mm

ym

m

510152025308

10

12

z m

m

0

2

4

35 40 45 50 55r mm

(c) 10 kHz

x mm

ym

m

510152025308

10

12

0

z m

m

2

4

35 40 45 50 55r mm

(d) 20 kHz

x mm

ym

m

510152025308

10

12

z m

m

0

2

4

35 40 45 50 55r mm

(b) 5 kHz

4.5

3.6

2.7

1.8

0.9

0

U [m/s]

Fig. 5 Time-averaged velocity distributions near the

impeller exit for various fp. Peak-to-peak voltage was

fixed at Vp-p = 11.7 kV

0.0 1.0 2.0 3.0 4.0 5.08

9

10

11

12

13

14off5 kHz10 kHz20 kHz30 kHz40 kHz

u [m/s]

z [m

m]

0

1

2

3

4

5

6

Fig. 6 Averaged u profiles at r = 33.9 mm

-6.0

-5.0

-4.0

-3.0

-2.0

-1.0

0.0

0 10 20 30 40

125mm 95 mm 65 mm

P [

Pa]

f p[kHz]

Fig. 7 Static pressure as a function of fp


71 / 298

4

PA

9

fp = 5kHz P

10 W

= 1,500 rpm

PA

PA

PA

peak-to-peak

= 1,500 rpm

peak-to-peak Vp-p = 11.7 kV

fp

= 5 kHz 10 kHz 20 kHz 30 kHz 40 kHz

(1) fp = 10, 20, 30, 40 kHz PA

fp = 5 kHz

5.1%

(2) 65 mm 95mm

fp = 5 kHz

fp = 5 kHz P 10 W

JSPS (B) 2689040

RA

-6000 -4000 -2000 0 2000 4000 6000

-3x10-7

-2x10-7

-1x10-7

0

1x10-7

2x10-7

3x10-7

V [V]

Q[C

]

Vp-p = 11.7 kVfp-p = 20 kHz

Capacitor for measuring electrical charge (Cr )

VPA

Vr

A

PA (CPA)

+QPA

-QPA

+Qr

-Qr

QPA = Qr = CrVr

Oscilloscope

(a) Circuit diagram

(b) V-Q Lissajous figure

Fig. 8 Evaluation of power consumption

Power

0

10

20

30

40

50

60

70

0 10 20 30 40 50

P [

W]

f p [kHz]

Fig . 9 Power consumption as a function of fp


72 / 298

(1) , , , ,

,

, (B ), 79 801 ,

pp. 900-916 (2013).

(2) , , ,

,

, Vol. 49, No. 1, pp. 69-74 (2012).

(3) Corke, T. C., Enloe, C. L., Wilkinson, S. P.,

Dielectric Barrier Discharge Plasma Actuators for

Flow Control, Annual Review of Fluid

Mechanics, Vol. 42, pp. 505-529 (2010).

(4) , ,

,

, Vol. 29, No. 4, pp. 243-250

(2010).

(5) Van Ness II, D. K., Corke, T. C., Morris, S. C.,

Tip Clearance Flow Visualization of a Turbine

Blade Cascade with Active and Passive Flow

Control, ASME Paper, GT2008-50703 (2008).

(6) Saddoughi, S., Bennett, G., Boespflug, M.,

Puterbaugh, S. L., and Wadia, A. R.,

Experimental Investigation of Tip Clearance

Flow in a Transonic Compressor with and without

Plasma Actuators, ASME Paper, GT2014-25294

(2014).

(7) Matsunuma, T., Segawa, T., Active Control of

Tip Leakage Flow for Low-Pressure Turbine by

Ring-Type Plasma Actuators, AIAA Paper,

2013-2726 (2013).

(8) , , ,

,

2015 , S0530402 (2015).


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44 2016.10

B-2

Studies on Heat Transfer Coefficient and Flow Field through a Complex Cooling Channel for Turbine Blade (Influence of the channel entrance inflow condition)

Ken-ichi FUNAZAKI, Sekai TERABE(Iwate Univ.), Masahide KAZARI, Takeshi HORIUCHI(KHI)

ABSTRACT

This paper deals with experimental and numerical analysis on the heat transfer and the flow field in a realistic serpentine

internal cooling channel of turbine blade. This study aims at obtaining design guidance of the internal cooling channel

through the investigation of heat transfer coefficient with the focus on the flow condition at the channel entrance. A

transient method using Thermochromic Liquid Crystal is employed to measure the surface heat transfer distribution inside

the channel. Numerical analysis using ANSYS CFX is also carried out to enhance the knowledge of the flow field.

Key words: Internal cooling channel, Serpentine channel, Heat transfer, Flow field, CFD

1.

2.

2.1


103 / 298

Fig. 1 Experimental set up

Fig.2 Test model

2.2

3

PS

SS

3 T1-T7,P1-P7

3

60°

2

Fig. 3 Flow channel inside the test model

3.

3.1 (5)

1st

pass, 2nd pass, 3rd pass

T1,T3,T5,T7

90°(6)

3 A-A

D(=33.4[mm])

ReD(=25,000)

4

3.2

(Hue)

RGB Hue (5).

4


104 / 298

Hue

5 -Hue

6

Hue S

1

3

Fig. 4 Calibration device

(a)PS

(b)SS

Fig. 5 Calibration curve: (a) PS, (b) SS

3.3

sabbas(7)

14 11

13 14

4

4.1

6

7

ANSYS ICEM CFD

0.067D

y+<1

ANSYS CFX Ver.14 1

(1) q(x,y)

Tw(x,y) Tg(x,y)

T1,T3,T5,T7

1mm

Fig. 6 Computational domain

Fig. 7 Computational grid


105 / 298

Table. 1 Computational conditions

5

5.1

8-10

8 1st pass

0-0.7

0.7-1

9 2nd pass

10

3rd pass

Fig. 8 Comparison of heat transfer coefficient on PS extracted

along the passage centerlines (1st pass)

Fig. 9 Comparison of heat transfer coefficient on PS extracted

along the passage centerlines (2nd pass)

Fig. 10 Comparison of heat transfer coefficient on PS

extracted along the passage centerlines (3rd pass)

11-13

11 1st pass

0.7-1

12

2nd pass

13 3rd pass

Fig. 11 Comparison of heat transfer coefficient on SS

extracted along the passage centerlines (1st pass)


106 / 298


extracted along the passage centerlines (2nd pass)


extracted along the passage centerlines (3rd pass)

14 15

2nd pass

Fig. 14 Heat transfer coefficient distributions

(PS_with_honeycomb)


(SS_with_honeycomb)

16

17

2nd pass

2nd pass

Fig.16 Heat transfer coefficient distributions

(PS_CFD)


(SS_CFD)

5.2

18 19 3

1st pass X/D=3 X/D=13 Q

Hn

18 YZ

(i),(ii)

(iii) (iii) (i)

(ii)

X

19

X


107 / 298

1st pass

Fig.18 Distribution and shape of vortex core (near SS)

Fig.19 Distribution and shape of vortex core (near PS)

PIV

6

1st pass

1) J. C. HAN. and J. S. PARK, INT. J. Heat Mass Transfer,Vol.31,No.1,pp.183-195,(1988)

2) CFD 39

,2011,pp.131-13

3) Christopher LeBlanc, Srinath V. Ekkad, Tony Lambert and Veera Rajendran Detailed Heat Transfer Distributions in Engine Similar Cooling Channels for a Turbine Rotor Blade with Different Rib OrientationsASME Turbo Expo GT2011 45254,(2011)

4) Pearce, R. , Ireland, P. T. Mc Gilvray, M., and Romero, E. omputational Study of the effect of Inlet Velocity

Profile and Rib Orientation on Heat Transfer in Rotating Ribbed Radial Turbine Cooling Passages E Turbo Expo 2016,GT2016-57832,(2016)

5)

Vol.26 No.101, 1998, pp.76

6) 58-548 B 1992 1015 7) Saabas et al., Application of the Transient Test Technique

to Measure Local Heat Transfer Coefficients Associated with Augmented Airfoil Cooling Passages, ASME Paper 87-GT-212, 1987


108 / 298

B-3

PSP

ABSTRACT

This paper describes several attempts to optimize double flow control devices (DFCD), which was invented by some of the authors for

achieving better film cooling performance of gas turbine cooling holes. The device, which will be referred to as DFCD, is a pair of

protrusions with the shape of hemi-spheroid attached to the turbine blade surface just upstream of each of cooling holes. It has been revealed

in the previous studies that the optimized devices are able to improve film cooling effectiveness dramatically. However DFCD optimization

has only been done under low density- ratio condition, therefore there is room for the improvement of device shape and configuration for

the use at high density ratio conditions. In order to investigate the film effectiveness at a high density ratio, we use CFD-based Taguchi

Methods to optimize DFCD shape prior to the experiment. In the experimental study, we apply CO2 as the cooling air and measure film

cooling effectiveness using PSP technique for DR=1.53.

Key words: Film Cooling, Flow Control, Taguchi-Method, PSP

TIT

CRVP

CRVP

DFCD

DFCD DFCD

DBV

PSP DFCD

CFD

DFCD PSP

CFD PSP

Fig.1

150mm 250mm

CO2 CO2

ABS

PSP

UVLED

CCD

Fig.2

PSP


109 / 298

Fig.1 Wind Tunnel for Film cooling Investigations

Fig.2 PSP Excitation View

Fig.3

Base Block Hole Block

Hole Block

Round Hole(RH) RH

Fig.4 d=10mm

=30deg. L=6d

3d 3

6d DFCD

DFCD DFCD

DFCD

DFCD

DFCD

Fig.3 Test Model

Fig.4 Round Hole Geometry

DFCDFig.5 DFCD Flow Model

Counter rotating vortex pair (CRVP)

CRVP lift-offDFCD CRVP

Device Based VorticesDBV

DFCDDBV DFCD

Fig.5 Flow Model of DFCD

Fig.6 PSP

(PSP: Pressure Sensitive Paint)

PSP

PSP

CO2

PSP


110 / 298

(1)

Fig.6 Schematic View of PSP measurement

T C Po2

air mix CO2

Charbonnier 2)

(2)

MW Po2

ref air, gas

CO2

air

(Po2;air)/(Po2;ref) 1

PSP

PSP

PSP

PSP (3) Stern-

Volmer

PSP

Fig.7 Stern-Volmer

I P PSP A B C

ref

dark

PSP

(PtTFPP)

PtBS

PSP UVLED (

LEDH60 400nm)

CCD (BITLAN BU-51LN)

PSP

( LV0630 630nm )

S/N

0.05s f1.4

60

Fig.7 Stern-volmer Plot

PSP

CCD

PSP Fig.8

Fig.8 Data Procession of Film cooling Investigation

PSP

Iref CO2

I, Idark 3

I Iref Idark

PSP

airO

mixOairOwallwall

P

PP

CCCC

TTTT

2

22

22

1/

/1

11

;;

;;

22

22

refogaso

refoairo

PP

PPMW

2

refrefdark

darkref

pp

Cpp

BAII

II


111 / 298

(1) DFCD

(2)

(ex. The height of DFCD =3mm, 4mm, 5mm) (3) (4) (2)

DFCD (5) (4) DFCD

CFD (6) (5) CFD

SN

(7) CFD

Round hole 2.2DFCD

A G Table 1Table 2

Kawabata 1)

DFCDFig.9

L1818 DFCD

Table 1 Control factors

Table 2

Fig.9 DFCD geometry and control factor

DFCD

DFCD

BR=1.5 3.0

DFCDBR=1.0 DFCD

(4) SN SN

Y 0 x/d 20 -3.0 z/d 3.0i (Case01

Case18) n n=2Y w/o DFCD

(5)

25deg.,

20deg.,

111log10/

iii YYn

NS

,/

,´

,oDFCDw

i

i Y

YY

ANSYS CFX 14.0(ANSYS Inc.)RANS

SST

Fig.101

21d ANSYS ICEM CFD ver.15.0

DFCD

1600

2d Density


112 / 298

Fig.10 Computational Domain

36 (= :18× :2)Fig.11

D F SN

SN

A3-B1-C2-D3-E3-F1-G1(Optimal)

Fig.11 Response graphs for major effect

d

Re=6000 (6)

BR=0.5,1.0,1.5

(6) U 2

(Air) (CO2)

DR 2 )=1.53 MW=1.519

(RH) DFCD (Optimal)

DFCD

BR=1.0 BR BR

BR=0.5

1.0 1.5

-3.0 z/d 3.0, 0 x/d 20

Fig.12 BR=0.5

RH

Jet

DFCD

CRVP DFCD

Mix-out

Fig.15

x/d=10 RH

Fig.16

DFCD 20

Fig.12 Film Effectiveness Distributions at BR=0.5

[(1)Only Round Hole, (2)Round Hole + DFCD]

BR=1.0

Fig.13 RH BR=0.5

DFCD DBV CRVP

x/d=3.0 RH

1.2


[(1)Only Round Hole, (2)Round Hole + DFCD]

BR BR=1.5

Fig.14 RH BR

UU

BR 22


113 / 298

CRVP

DFCD

x/d 2

2.7

BR DFCD

.

PSP

RH

BR=1.0 Schmidt 3) Pedersen 4) Hui 5)

Fig.17 Pedersen

Re


[ (1)Only Round Hole, (2)Round Hole + DFCD]

Fig.15 Span wise-Averaged Film Effectiveness

Fig.16 Area-averaged film effectiveness

Fig.17 Comparison of RH centerline-effectiveness to published data

PSP

DFCD

1)PSP

2) DFCD

BR=1.0 BR=1.5

2.7

BR=0.5

DFCD

PIV

1) Hirokazu Kawabata, Ken-ichi Funazaki, Ryota Nakata,

2014, -Plate Film Cooling Performance

by Double Flow Control Device : Part - Optimization of

device shape and arrangement by experiment-And CFD-based

Taguchi Method ASME, GT2014-26070

2) Experimental and Numerical

Study of the Thermal Performance of Film Cooled Turbine

Platform Proc. of ASME Turbo Expo

3)Cooling With Compound Angle Holes: Adiabatic

Journal of Turbomachinery, vol. 118, Oct. 1996, p. 807. 4) Pedersen, D. differences between the mainstream and the secondary fluid measured by the heat- Journal of heat transfer, 1977. 5) Wenwu Zhou. , Blake Johnson. , Hui Hu. , 2015,

AIAA 53rd Aerospace Science Meeting


114 / 298

B-4

Endwall Secondary Flow Control in Low Pressure Turbine Effect of Devices

*Ryota SATO(Iwate Univ.), Kenichi FUNAZAKI(Iwate Univ.), Teruyuki SHIOTA, Daichi MURAKAMI, Hirohumi OGURA, Kodai ITO(Iwate Univ.), Juo HURUKAWA(IHI)

ABSTRACT

The purpose of this paper is to propose a new method of secondary flow control on end-wall region for End-

wall loss reduction in Low Pressure Turbine. In this paper, the effect of dimple devices on endwall is investigated

through combined experimental and CFD studies in a low speed linear cascade. Detailed comparisons are made

on the cases with or without dimple devices on the End-wall. The results show that total pressure loss of near

endwall regions can be reduced because the vorttices generated by the dimple devices suppress the advections of

the cross-flow to the blade suction side, eventually reducing the magnitude of stream-wise vorticities of passage

vortex and counter vortex.

Key words: Low Pressure Turbine, End-wall, Secondary Flow, Cross-flow, Vortex, Control


115 / 298

2 Dimple

1 Dimple

EWD 1 EWD_Triple

3 Dimple

2,3 Dimple

Base EWD

EWD EWD_Triple


116 / 298

EXP results Base vs EWD


117 / 298


118 / 298

EWD

EWD_Triple


119 / 298

B-5

Application of CHT Analysis to Development of Cooled Turbine Blade for Industrial Gas Turbine

*Takeshi HORIUCHI, Tomoki TANIGUCHI, Ryozo TANAKA, Masanori RYU and Masahide KAZARI (Kawasaki Heavy Industries, Ltd.)

ABSTRACT

Advanced cooled turbine blades which have high cooling effectiveness are essential as well as rise of turbine

inlet temperature to improve thermal efficiency of industrial gas turbines. For this development, we need to

understand metal temperature distribution on the cooled turbine blades in detail. For this purpose, applicability

of CHT (Conjugate Heat Transfer) analysis to the cooled turbine blade has been investigated in terms of

prediction of metal temperature by comparing with measurement results by pyrometer. This paper describes the

details of CHT analysis method and calculation results with measurement temperature by pyrometer.

Key words: Cooled turbine blades, Metal temperature, Prediction, Conjugate heat transfer, Pyrometer

3

1

CHT Conjugate Heat

Transfer

CHT

CHT

MARK , C3X

TBC

CHT

CHT

CHT

Fig.1 CHT analysis


120 / 298

2 30MW

“L30A”

GGT 1

1 30.1MW

40.1%

PT 2 3

GGT 1 3

PT 1 3

Fig.2 L30A gas turbine

Table 1 Characteristic feature of L30A

230.1MW 40.1% 24.9

470

ISO 13A

CHT Conjugate Heat Transfer

2

CFD FEM

STAR-CCM+ Ver.10.04

CD-adapco 3

CHT

Fig.3 Conjugate approach

4

1

STAR-CCM+

1330 1100

230

y+ y+ < 3

y+ 1


121 / 298

a) Overview

b) Mesh on blade surface

Fig.4 Computational domain and mesh system

FINE/Turbo NUMECA

CFD CHT

SST k-

1

( )1exp

,2

51

−⋅

⋅⋅=

−

T

C

CTq

λ

λελ 1

C1, C2 1 2

q

T

Rotadata

2

8

Table 2 Major specifications of pyrometer system

600 – 1200±0.5%RD

2mm 75mmInGaAs 1.6 m 1MHz

5 CHT

6

CHT

A F

4.2

Fig.5 Metal temperature distribution by CHT analysis

A

B

C

D

E F


122 / 298

Fig.6 Metal temperature distribution by pyrometer

7 Hub Mean Tip

CHT

8 10

0

LE PS

SS

SS

8 Hub -0.8

-0.4 0.1 0.2 0.7 0.9

20 55

9 Mean LE

0.2

PS

25 60

SS 30 60

10 Tip LE

0.2

PS

25 50

SS 55

20 60

2mm

CHT

Fig.7 Definition of radius location for evaluation

Fig.8 Metal temperature distribution at Hub

Fig.9 Metal temperature distribution at Mean

LEPS SS

LEPS SS

A

B

C

D

E

F


123 / 298

Fig.10 Metal temperature distribution at Tip

CHT

A C D F

11 LE

A

Mean

Tip

A 30%

12

LE

LE

A

Fig.11 Results of CHT analysis for leading edge

Fig.12 Radial temperature profile at main flow inlet

and streamlines from LE-film cooling holes

13

C

C

Fig.13 Close-up of trailing edge’s cutback near hub

14 PS CHT

D

D

30 50%

LEPS SS

A


124 / 298

D

50% F

Fig.14 Results of CHT analysis for trailing edge side

CHT

CHT

CHT

CHT

K. Kusterer, et al, “CONJUGATE CALCULATIONS FOR A FILM-COOLED BLADE UNDER DIFFERENT OPERATING CONDITIONS”, ASME Paper GT2004-53719D. E. Bohn, C. Tümmers, “NUMERICAL 3-D CONJUGATE FLOW AND HEAT TRANSFER INVESTIGATION OF A TRANSONIC CONVECTION-COOLED THERMAL BARRIER COATED TURBINE GUIDE VANE WITH REDUCED COOLING FLUID MASS FLOW”, ASME PaperGT2003-38431J. Luo, E. H. Razinsky, “CONJUGATE HEAT TRANSFER ANALYSIS OF A COOLED TURBINE VANE USING THE V2F TURBULENCE MODEL”, ASME Paper GT2006-91109G. Nowak, W. Wróblewski, “APPLICATION OF CONJUGATE HEAT TRANSFER FOR COOLING OPTIMAIZATION OF A TURBINE AIRFOIL”, ASME Paper GT2009-59818Kuo-San. Ho, et al, “CONJUGATE HEAT TRANSFER ANALYSIS FOR GAS TURBINE COOLED BLADE”, ASME Paper GT2014-25952K. Tsukamoto, et al, “CONJUGATE HEAT TRANSFER ANALYSIS IN AN ACTUAL GAS TURBINE ROTOR BLADE IN COMPARISON WITH PYROMETER DATA”, ASME Paper GT2014-26962Kuo-San. Ho, et al, “CONJUGATE HEAT TRANSFER ANALYSIS FOR GAS TURBINE FILM-COOLED BLADE”, ASME Paper GT2016-56688T. Taniguchi, et al, “APPLICATION OF AN OPTICAL PYROMETER TO NEWLY DEVELOPED INDUSTRIAL GAS TURBINE”, ASME Paper GT2012-68679Bohn, D., Krüger, U., and Kusterer, K ., 2001, “Conjugate Heat Transfer : An Advanced Computational Method for the Cooling Design of Modern Gas Turbine Blades and Vanes”, Heat Transfer in Gas Turbines, eds. B. Sundén and M. Faghri, pp.58-108, WIT Press, Southampton, UKT. Taniguchi, et al,“TEMPERATURE MEASUREMENT OF HIGH SPEED ROTATING TURBINE BLADES USING A PYROMETER”, ASME Paper GT2006-90247R. Tanaka, et al, “CONTINUING IMPROVEMENTS OF 20MW-CLASS GT KAWASAKI L20A”, ASME Paper GT2006-90394

D


125 / 298

44 2016.10

B-7

A Surrogate Model for Turbine Blade Temperature Estimation

*Eiji SAKAI (CRIEPI) and Toshihiko TAKAHASHI (CRIEPI)

For the development of easy-to-use temperature estimation method for turbine blade, polynomial chaos

expansion, which is a computationally efficient method for performing uncertainty analysis on large complicated

models, is applied to 3-D turbine blade conjugate heat transfer simulations. Main flow temperature, main flow

rate, cooling flow temperature and outlet pressure are selected as uncertain parameters. Surrogate models for

blade temperature estimation are obtained from 15 sample simulation runs. The estimated temperature by the

surrogate models correspond well with the conjugate heat transfer simulations, and the surrogate models are

found to be useful for blade temperature estimation in response to a change of gas turbine operation conditions.

Key words: Surrogate Model, Temperature Estimation, Turbine Blade, Polynomial Chaos Expansion

1),2 PCE

PCE

1100

PCE

Fluent ver12

3 NS

SA

SIMPLE

Fig.1 Turbine Blade and Computational Grid

Inlet

Outlet

Periodic

Periodic

Cooling Holes


126 / 298

Fig.1 1

12

230

140 90

1 y+<1

Xi Tg

Gg Tc Gc

Pout Gc

Gg 2.0

4 Xi

Table 1 Xi

Xi i

Xi= + i

3600rpm 76

5%

Table 1 and

i Xi

1 Tg, K 1300 80

2 Tc, K 590 30

3 Gg, kg/s 2.8 0.3

4 Pout, Pa 621500 75660

Table 2 Sample Points

Run Tg, K Tc, K Gg, kg/s Pout, Pa

Run0 1300 590 2.80 621500

Run1 1300 642 3.32 752500

Run2 1300 538 2.28 490500

Run3 1439 590 3.32 490500

Run4 1439 642 2.28 621500

Run5 1439 538 2.80 752500

Run6 1161 590 2.28 752500

Run7 1439 642 2.80 490500

Run8 1161 538 3.32 621500

Run9 1439 642 3.32 752500

Run10 1161 538 2.28 490500

Run11 1300 590 3.32 752500

Run12 1439 642 2.80 621500

Run13 1161 590 2.28 621500

Run14 1300 538 2.80 490500

Table 3 Stochastic Basis Functions, i

i Order, p i

0 p=0 1

1

2

3

4

p=1

5

6

7

8

9

10

11

12

13

14

p=2

PCE 2 P 1=15

Table 2

T

1 fi

i

Table 3 Table 2

3 i

= i i

Tos

Tis Tcs 50

Ts

fi

Table 4

Run15 18 4

i = i i

Run19 Run20 50%

Run21

Fig. 2 Tos

Tis


127 / 298

Fig.

3 Fig.4 Tcs

50% Ts

Fig.4

0 1 PS

SS

Table 4 Validation Points

Run Tg, K Tc, K Gg, kg/s Pout, Pa

Run15 1359 660 3.02 798100

Run16 1487 612 3.50 677600

Run17 1241 520 2.58 444900

Run18 1113 568 2.10 565300

Run19 1369 603 3.05 675400

Run20 1275 578 2.6 561400

Run21 929 525 2.2 379600

Fig. 2 Comparison of Tos and Tis

Fig. 3 Comparison of Tcs

Fig.4 Comparison of Ts

Fig.2 4

Run21

Fig. 5 Run21 Table 1

Run21

Run21

Run21

Run21

Fig. 5 Probability of Run21

600 800 1000 1200

800

1000

1200

Tem

p, K

(S

urro

gate

Mod

el)

Temp, K (CHT)

Run16

Run15

Run19

Run20Run17

Run18

Run21

Run21

Run18

Run16

Run15

Run19Run20

Run17

TosTis

Tis

Tos

800 10000

50

100

Tcs, K

Spa

n, %

CHTSurrogate Model

Run21

Run18

Run17 Run20

Run19

Run15

Run16

1 0 1

800

1000

1200

x/ST

s, K

LETE TEPS SS

CHTSurrogate Model

Run21

Run18

Run17

Run20

Run19

Run15

Run16

1000 1200 1400 1600

Pro

babi

lity

500 600 700

Pro

babi

lity

2 3 4

Pro

babi

lity

2 4 6 8 10105

Pro

babi

lity


128 / 298

Fig. 6 Sensitivity Analysis

Tg Tc

Gg Pout

Fig. 6 S S

Fig.6 Tcs20, Tcs50, Tcs80

20%, 50%, 80%

Fig.6

Tg Tos

Tcs

Tis

Tos Tcs

Tg Tc

Tcs Tc Tg

1100

PCE

1) Stefanou, G.: The Stochastic Finite Element Method: Past, Present and Future, Comput. MethodsAppl. Mech. Engrg., Vol.198, No.25 (2013) pp.1031-1051.

2) Kato, H., Mashiko, H., Tokuyama, Y., Funazaki, K., and Takida, J. : Robust Aerodynamic Shape Optimization of Supersonic Turbine Using Non-Intrusive Polynomial Chaos Expansion, 9th World Congress on Structural and Multidisciplinary Optimization, (2011).

0

40

80

Tos

Tis

Tcs20 Tcs50 Tcs80

Main Flow Temperature, TgCooling Flow Temperature, TcMain Flow Rate, GgOutlet Pressure, Pout

S, K


129 / 298

B-8

CO2

Flow visualization experiment of super critical CO2 turbine bladecooling passage

*Hisashi MATSUDA, Fumio OTOMO, Tatsuro UCHIDA, Junichi TOMINAGA,Hideyuki MAEDA and Satoru SEKINE (Toshiba Corp.)

ABSTRACT

Flow visualization experiments for super critical CO2 turbine blade cooling passage were examined using a large scale cooling flow passage model. Complex configuration of the cooling passage was exactly modeled by combination of plexiglass pipes with stereo lithography molding. Flow visualization experiments clarified that there is no measure stagnated flow through the cooling flow passage. PIV measurements were also carried out and maximum velocity of each cooling passage was obtained accurately and balance of flow volume of each cooling passage was evaluated. The supercritical CO2 turbine blade cooling passage designed was validated.

Key words: CO2 turbine, Cooling blade, Flow visualization, PIV

CO2

CO2

CO2

CCS Carbon Capture and

Storage

CO2

CO21-4)

CO2

CCS CO2

CH4

1,150

30MPa

1,150

30MPa

(CO2 H2O)

CO2

CO2

1/30

CO2

EOR Enhanced

Oil Recovery

CO2

Fig.1 CO2

CO2

1.5 2.01)

CO2

1

CO2 3 No.1

3 3 (No.4 6 ) 5


130 / 298

(No.7 11 )

1)

CO2

( )

PIV

CO2

CFD

CO2

ANSYS CFX v150

SST 2000

y+=3

CO2

Fig.2

Fig.3,4 Fig.2

Fig.3,4

Fig.3

Fig.4

CFD

CO2

Re

CO2 Re

10 4 5


131 / 298

CO2

Fig.2 PIV

20mm 5mm

PIV

Fig.5

Fig.6

3.2 PIV

1mm 15

10mm

30Nl/min Re 8000

PIV

Photron FASTCAM SA-Z

Lee Laser

532nm, 100W

Labsmith LV880

PIV Koncerto

1024×1024 pixels 1kHz

1000

Fig.7

PIV

PIV

PIV

PIV Fig.8

11(a,b) Fig.8 No.1 (

20mm, 160mm )

Fig.9 No.4 (


132 / 298

20mm 160mm )

Fig.10 No.8 ( 8mm

160mm ) Fig.11 No.11 (

5mm 150mm )

Fig.a

Fig.b Fig.b PIV

1000

r D

(1974)5)

3mm No.11 2.5mm

PIV

Fig.10 11

8mm 5mm PIV

(32 32 16 16

) PIV

Fig.9 No.4

PIV

No.1 3 No.4 6

20mm PIV

No.7 11


133 / 298

No.1 3 No.4

6 No.7 11

1/7 6)

No.7 No.11

Table 1 5

0.9%

PIV

PIV

CO2

Fig .12 CFD

No.1 No.3 No.4 No.6

No.7

No.11 PIV CFD

CFD

4

CO2

ANSYS CFX v150

PIV

8mm 5mm ,

PIV

PIV

CO2

41

4) Y.Nakamura, Y.Iwai1, M.Itoh, Y.Morisawa, T.Sasaki,

D.Cusano, M. Harris and E. Petersen (2015),

Table 1 Measured maximum velocity and integrated flow

volume by 1/7 order row for flow passages No7-No.11


134 / 298


135 / 298


136 / 298


137 / 298


138 / 298


139 / 298

B-12

A Method of Creep Lifetime Estimation of Superalloy SpecimensPrepared from an In-Service Gas Turbine Blade

*Yohei MATSUI, Mitsutoshi OKADA, Toshihiko TAKAHASHI and Terutaka FUJIOKA(CRIEPI)

ABSTRACT

Creep lifetime estimation of a gas turbine blade is conducted by the iso-stress method. Three types of

specimens with different shapes and dimensions are prepared from an in-service blade, and creep tests are

carried out using them at high temperatures. To estimate the creep lifetime of the blade, the reference

temperature is calculated by heat transfer numerical analysis which combines convection heat transfer analysis

around the blade with heat conduction analysis in blade structure. As a result, larger differences in creep lifetime

among the three types of specimens are caused at higher test temperatures. On the basis of the result, the effects

of specimen shape and test atmosphere on creep lifetime estimation are examined. And tests at lower

temperatures with longer rupture life are desired to be performed to obtain life evaluation with smaller scatters.

Key words: Iso-Stress Method, Creep, Gas Turbine, Superalloy, Size Effect


140 / 298

1

1

Ni GTD-111

3

4mm 4

2mm 2

2mm 1mm

3

2

1 910 930

Fig.1 The geometry of the blade

Fig.2 The mean temperature in the

cross-sectional surfaces of the blade

Fig.3 The shape and the size of the

specimens

(c)flat plate specimen

(b)

(a)

(c)


141 / 298

950 970 990 5

137MPa 8) 4

2 Ar

Table.1 Chemical composition of GTD-1117

Fig.4 The SEM images of the specimens after

the creep tests(1223K 137MPa) (a)

specimen,(c)flat plate

specimen

(c) 5

(b) 5

(a) 5


142 / 298

Fig.5 The results of the creep tests


143 / 298

B-13

Temperature Estimation Method for a Gas Turbine Combustor Component

*Mitsutoshi OKADA, Toshihiko TAKAHASHI, Susumu YAMADA and Tomoharu FUJII (CRIEPI)

ABSTRACT

Temperature estimation methods for a gas turbine combustor component are developed in terms of

microstructural changes and computational fluid dynamics (CFD) for life assessment. The growth of thermally

grown oxide (TGO) in thermal barrier coating is observed, and the change in Vickers hardness is also grasped by

carrying out high-temperature heating test using the specimens prepared from the in-service combustor. On the

basis of them, temperature estimation equations are derived. The temperature estimation based on TGO thickness

is suitable for higher-temperature region while the one based on the hardness is for lower-temperature. Conjugate

heat transfer simulation, coupling heat conduction calculation in solid with CFD calculation of heat transfer in

fluid, is also conducted to estimate the temperature distribution of the combustor component. The simulation result

agrees well with the estimation using the TGO thickness and hardness.

Key words: Compressor, Combustor, Turbine, Control, Maintenance

1,2)

5)

6)

7)

GT

1

(Alloy A) TBC


144 / 298

1

Table 1 Material and operation data of combustor

component

1 Alloy A

Alloy B

1

Alloy C 1

TBC

TBC

1(b)

2 FE-SEM

2(a) B

2(b) A

560 m

90 m

APS

(a) Shape of combustor component

(b) Positions for temperature estimation

(Upper: u, Side: s, Down: d)

Fig. 1 Structure of combustor component

2(b) A

TGO

APS

EPMA A

TGO Al

Co Ni Cr Al

8)

TGO TGO

B TGO

3 TGO

TGO SEM

Air intake

inlet

Outlet

Pipes for cooling steam

Acoustic Liner


145 / 298

1000 3

TGO 10

30 TGO

TGO TGO/

TGO

A TGO

B 1 m

(a) B1-u

(b) A1-d (at the vicinity of outlet)

Fig. 2 Morphologies of interface between topcoat and

bondcoat

Fig. 3 Distribution of Outer TGO thickness

Fig. 4 Relationship between outer TGO thickness and

square root of test time

3

TGO

B

20mm

10mm

60mm

873K 973K 1073K 1123K 1173K 1223K

1273K

9000h

TGO

4.9N 10

500 m 5

200 m

4

TGO

TGO

B TGO

0.3 m 1123K

TGO Al

Topcoat

Bondcoat20 m

Outer TGO

Outer TGO

20 mInternal oxide

Topcoat

Bondcoat

Out

er T

GO

thic

knes

s [

m]

A1A2B1C1D1E1 C2

Alloy B Alloy CAlloy A

PositionOutletInlet

UpperSideDown

Square root of test time [h1/2]

Cha

nge

in o

uter

TG

O th

ickn

ess

[m

]

1073K1123K1173K1223K


146 / 298

1273K Al

TGO

873K 973K

6000h

1073K 1223K

TGO

dTGO

t (h)

(1)

kTGO TGO

TGO

Al

2(b TGO 5 m

5 kTGO

kTGO

(2)

R 8.31J/(mol K) T

K

Fig. 5 Arrhenius plot of growth rate constant of outer

TGO thickness

Fig. 6 Relationship between Vickers hardness and test

time

Fig. 7 Relationship between Vickers hardness and

Larson-Miller parameter

6 h (Hv) t

B

208Hv 873K 973K

1073K

9)

Alloy A M23C6

7)

7 h LMP

LMP l

(3)

Gro

wth

rat

e co

nsta

nt o

f ou

ter T

GO

thci

knes

s[lo

g ek T

GO]

Temperature [1/Kx104]

Vic

ker’s

hard

ness

[Hv]

Test time [h]

873K973K1073K1123K1173K1223K1273K

Vic

kers

har

dnes

s [H

v]

Larson-Miller Parameter LMPT(20+log10t)/103


147 / 298

LMP

l=22

4)

h LMP l

2

Table 2 Equations for the relationship between Vickers

hardness and LMP

l 23.623.6<l 26

26<l

TGO dTGO

(3)

2

2

LMP

LMP (3)

4 6

TGO

TGO

A1 TGO

8

4 6

TGO

TGO

TGO

Fig. 8 Result of estimated temperature in the vicinity of

internal wall

Fig. 9 Numerical mesh (Surface mesh)

9 1

Alloy A Alloy

B

1

CFD (TBC

)

TBC

Est

imat

ed te

mpe

ratu

re [K

]

A1A2B1C1D1E1 C2

Alloy B Alloy CAlloy A

PositionOutletInlet

Upper Side DownOuter TGO and hardnessNumerical simulation


148 / 298

1

FLUENT(Ver.12)

CFD Reynolds

Navier-Stokes

RANS Reynolds Averaged

Navier-Stokes Simulation

Spalart & Allmaras10)

0.85 0

CFD

SIMPLE

Discrete Ordinate

(DO )11)

8

1

10

8

TBC

TGO

Fig.10 Local metal temperature on the inner wall

1)

Vol. 45, No. 6 (1996) pp. 699-704.2)

Vol. 56, No. 8 (2007) pp. 757-763.

3) CoCAMP-ISIJ Vol. 9

(1996) pp. 1267.

4)

Vol. 38, No. 2 (2010) pp.102-

108.

5) Heidmann, J.D., Kassab, A.J. and Steinthorsson, E.:Conjugate Heat Transfer Effects on a Reakistic Film-cooled Turbine Vane, ASME Paper GT2003-38553 (2003).

6) Dhiman, S., and Yavuzkurt, S.: Film cooling calculations with an iterative conjugate heat transfer approach using empirical heat transfer coefficient corrections, ASME paper GT2010-22958 (2010).

7) Okada, M., Takahashi, T., Yamada, S., Ozeki, T., Fujii, T.: Development of temperature estimation method for a gas turbine transition piece, ASME paper GT2016-56182(2016).

8)

Vol. 68 No. 6 (2004)pp.372-380.

9)

10) Spalart, P. R. and Allmaras, S. R.: A one-equation

turbulence model for aerodynamic flows, AIAA paper 92-

0439, 1992.

11) Modest, M. F.: Radiative Heat Transfer, McGraw-Hill, 1993.

1173

673

K


149 / 298


150 / 298


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44 2016.10

B-15

SUS316L

Strengths and Microstructure of SUS316L Fabricated by Selective Laser Melting

*Kazuki AKINO and Koji KAKEHI(TMU Univ.)

ABSTRACT

Recently, AM (Additive Manufacturing) process which can produce highly complex components is been

gaining significant attention. Ti alloys, Ni alloys, and Co-Cr alloys are the subjects of recent work.

Ni-based superalloys have phases such as and phase, and Ti alloys and Co-Cr alloys are the

multiphase alloys which have phase transformation, thus it is difficult to clarify the influential factors of

AM process on strengths of these alloys. In this study, we used SUS316L stainless steel which is a

single-phase solid-solution alloy and does not have precipitates, in order to clarify characteristic

influential factors of AM compared with a conventional material. It was found that the coarse columnar

grains grew up along the built direction and the columnar cell structure of dislocations which are induced

during the AM process. AM specimens showed higher tensile and creep strengths compared with the

conventional (hot working) material because of the high-density of dislocations.

Key words: additive manufacturing, austenite stainless steel, solid-solution alloy, built direction, molten

pool boundary

3D1) Ti Ni

Co-Cr

Ti 2) Co-Cr3) Ni

4, 5)

unexpected -Phase Formation 6)

SUS316L

LPW LPW316L (

10- SLM 3D

SLM280

Fig. 1 50 mm

AM Additive Manufacturing

3.1 mm

0 °

90 °

Fig. 1

Z Z plane X X plane

Y Y plane

PF Plastic Forming

20 mm

1080 °C SHT Solution

Heat Treatment


153 / 298

Fig. 1 Cutting AM specimen for microstructural

observation.

SHT

as-built SHT

SHT Ar

1080 °C 1 h

4.63×10-4 s-1 24 °C

600 °C

600 °C 235 MPa 260 MPa

OM Optical

Microscope SEM/EBSD

IPF Inverse Pole Figure Map

TEM

Fig.

2 Fig. 2

Fig. 2 Laser scanning trace on Z plane of as-built

specimen.

X Y Z IPF

Fig. 3 IPF Fig. 4

Fig. 3 IPF maps of as-built specimen.

Fig. 4 IPF map of PF(SHT) specimen.

Fig. 3 X

Y

Z

Fig. 4

24 °C 600 °C

Fig. 5 Fig. 6

specimen


154 / 298

(a) As-built specimens

(b) SHT specimens

Fig. 5 Stress-strain curves at 24 °C.

(a) As-built specimens.

(b) SHT specimens.

Fig. 6 Stress-strain curves at 600 °C.

24 °C

600 °C 90 ° 0 °

24 °C 90 °

90 °

0.2 %

600 °C 90 °

0 °

11 %

0 ° 24 °C

600 °C 2

600 °C

90 °

0 ° Fig. 6(b)

90 °

600 °C Fig. 8

600 °C 0.2 % 165 MPa

235 MPa7)

(a) Comparison between AM and PF specimens.

(b) Influence of built direction on creep curves.

Fig. 7 Creep curves of AM and PF specimens at

600 °C.


155 / 298

235 MPa 0 °

500 h 7) 45.12

h 50.62 %

260 MPa 0 ° 610 h

12.39 % 90 ° 692 h

10.20 % 90 °

TEM Fig.8

Fig. 8 TEM microstructures of (a) AM as-built, (b)

AM(SHT), and (c) PF(SHT) specimens.

Fig. 8(a)

Fig.

8(b) Fig. 8(c)

SUS316L 0 °

90 °

SUS316L

8)

106-7 K/s 9)

Nb Mo

Ni IN718

90 °10)

316L

90 °

Fig. 9

X

MPB Molten Pool Boundary 11)

Fig. 9 OM image of defect due to lack of fusion on X

plane of as-built specimen.

1.19 /mm2

MPB

0 °

MPB11) 600 C

90

0

Fig. 6(b) 90


156 / 298

0.2 %

MicroCT 5)

12)

Fig. 8

Fig. 8(c) 90 °

0.2 %

0 °

Fig.

7(a)

90 °

0.2 %

24 °C

0 °

0 °

0 °

(i)

0 ° 90 °

(ii)

(iii) 90 °

1) K. Kakehi: Material Stage, Technical Information Institute. 15 (2015) 45-52.

2) B. Vrancken, L. Thijs, J. Kruth and J. V. Humbeeck: Journal of Alloys and Compounds. 541 (2012) 177-185.

3) S.H. Sun, Y. Koizumi, S. Kurosu, Y. P. Li, H. Matsumoto and A. Chiba: Acta Materialia. 64 (2014) 154-168.

4) Y. Kuo and K. Kakehi: Proc. International Gas Turbine Congress 2015 (Gas Turbine Society of Japan), pp. 1031-1034.

5) L. N. Carter, C. Martin, P. J. Withers and M. M. Attallah: Journal of Alloys and Compounds. 615 (2014) 338-347.

6) Y. Idell, L. E. Levine, A. J. Allen, F. Zhang, C. E. Campbell, G. B. Olson, J. Gong, D. R. Snyder and H. Z. Deutchman: JOM. 68 (2016) 950-951

7) K. Kakehi: Master Thesis of Nagoya University, (1985).

8) M. Sugiyama, T. Umeda and J. Matsuyama: Tetsu-to-hagané. 60 (1974) 1094-1112.

9) Q. Jia and D. Gu: Optics & Laser Technology. 62 (2014) 161-171.

10) K. Kakehi: Mechanical Engineering, Nikkan Kogyo Shimbun. 64 (2016) 76-82.

11) W. Shifeng, L. Shuai, W. Qingsong, C. Yan, Z. Sheng and S. Yusheng: J. Mat. Proc. Technol. 214 (2014) 2660-2667.

12) M. Seifi, A. Salem, J. Beuth, O. Harrysson and J. J. Lewandowski: JOM. 68 (2016) 747-764.


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161 / 298

B-17

Effect of Re on creep behavior of single crystal Ni base superalloy Takeshi OHASHI Naoki NOMURA Koji KAKEHI (TMU)

ABSTRACT

Due to its good creep and TMF properties, the single crystal (SC) superalloy is the primary material to

manufacture the blades of the gas turbine and the aero-engine. The third or further generation Ni based SC

superalloys have been successively developed by adding more and more Re as strengthening elements. But

applying massive Re into SC alloys significantly causes many deleterious problems, such as i) the increase of the

density of alloys as higher than 9, reducing the engine efficiency, ii) the excessively high cost of alloys and iii)

the resource strategy. Thus the new alloying design for reducing Re is urgently desired. The effect of Re on creep

behavior of single-crystal superalloys at 750 /750MPa was investigated to consider alternative elements.

Key words: Single Crystal, Superalloy, Creep, Rhenium

1

P W(

) PWA1480

Re

Re 2-3mass% 2

Re 4-6 mass% 3

Ni

Re

Re

Step

Solution

)

Re 850

Re

2)

{111}<112> Re

{111}<112> TMF

TMF4)

Re

LSC(Low-cost Single Crystal)

1 2 {111}<112>

750 Re

2

Re

LSC-15 LSC-11m 1 2

PWA1480 CMSX-4 (Table 1)

Table 1

Table 2

[001] 2 °

2.8×2.8×19.6 mm

Alloy Co Cr Mo W Al Ti Ta Re Hf Ni

LSC-15 6 7 1.5 10 6 - 5.5 - 0.1 bal.

LSC-11m 8 7 1.8 8.2 5.4 1.2 6.2 0.8 0.1 bal.

PWA1480 5 10 - 4 5 1.5 12 - - bal.

CMSX-4 9 6.5 0.6 6 5.6 1 6.5 3 0.1 bal.

Table 1. Chemical composition of alloys (mass %).


162 / 298

{111}<112>

750 /750 MPa

TEM (Transmission

Electron Microscopy) TEM

{001} {111}

3

3.1 TEM

TEM Fig.1

Fig.2 LSC PWA1480

CMSX-4

CMSX-4 1

Fig.3

CMSX-4

5)

a<112>6) 1

1

3 TEM

1

2h TEM Fig.3

LSC-15 LSC-11m PWA1480

CMSX-4

<112>

1

Fig.4 CMSX4

(Fig.4(c))

LSC-15

a<112> 6)

1

LSC-11m

(Fig.4(b))

Fig.5(a) PWA1480 1

0.8% (111)

TEM TEM

a/2<110>

Fig.5(b) CMSX-4

1.4% (111)

TEM CMSX-4

a/3<112>

3.2

TEM EDS(Energy Dispersive x-ray

Spectroscopy)

3

Ki

(1)

i

i

(1)

Ki Fig.6

PWA1480 7)

Fig.6 Al Ti Ta Cr Co

Mo W Re

LSC11m Re Re

Alloy Solution heat treatmenta Aging treatment

LSC-15 Full solution treatment

1100 °C/4 h/GFC+

870 °C/20 h/GFC LSC-11m

PWA1480 1282°C/1 h + 1287°C/2 h + 1294°C/1 h/GFC 1080°C/4 h/AC+ 870°C/32

h/AC

CMSX-4

1277 °C/2 h + 1288 °C/2 h + 1296 °C/3 h + 1304 °C/3

hr+ 1313 °C/2 h + 1316 °C/2 h + 1318 °C/2 h +1321 °C/2

h /GFC

1140 °C/4 h/GFC +

870 °C/20 h/GFC

Table 2. Heat treatment schedule of alloys.


163 / 298

Fig. 1 TEM microstructures of (a) LSC-15, (b) LSC-11m, (c) PWA1480 and (d) CMSX4 after heat treatment

Fig. 2 (a) creep curves and (b) relationship creep strain and creep rate.

Fig. 3 TEM micrographs of the [001] samples interrupted at 2 h creep at 750 °C and 750 MPa: (a) LSC-15

.15%); (b) LSC-11m ; (c) PWA1480 ; (d) CMSX-4 ( . The foil normal

was cut along the (001) plane to observe the interfacial dislocations.

Fig.4 TEM image of (a)LSC- - - 750 /750MPa


164 / 298

Fig. 6 Partitioning coefficients of the alloying elements

in the standard heat treated condition.

3.3

) )

(Table 3)

(2)

10)

(3)

(4)

i

Table 3 PWA1480

Fig.7

Re Mo W

Re Mo W LSC-15 LSC-11m

CMSX-4

PWA1480

Ta

4

4.1

Fig.3 CMSX-45)

1 CMSX-4

PWA1480 CMSX-4 LSC-15 LSC-11m

6.805 6.206 6.123 6.443

5.984 5.011 4.718 5.068

-0.1283 -0.2131 -0.2592 -0.2389

Table 3. calculated by equations of (3), (4) and (5)

Fig. 5 TEM micrograph of the

[001] samples deformed to the

start of secondary creep at 750 °C

and 750 MPa: (a) PWA1480

; (b) CMSX4 .

The specimens were sectioned on

the (111) plane.


165 / 298

a/3<112>

(Fig.5(b)) a<112>

Link and Feller Kniepmeier11)

{111}<112> -

(i)

a/6<112>

(ii)

2

{111}<112>12)

W

Mo Re Fig.7

CMSX-4 LSC-15 LSC11m

CMSX-4 Re W

Mo Re Fig.6

13)

Re

Re

Fig.3 2 12)

Re

{111}<112>

Re

Ni

W 1/5 14,15) Re

Re

LSC-15 LSC-11m PWA1480 2 h

a<112>

1 CMSX-4

1 Re

a<112>

4.2

LSC-11m

(Fig.4(b))

Ru

16) Ru

LSC-11m

[001]

17)

(5)

b

(SFE)

16)

hcp Re Co18,19 ,20)

Fig.8 LSC-11m Re Co (mol%)

CMSX-4

LSC-11m LSC-15

(Table 3)

LSC-11m (5)

<112>

16,17) LSC-11m

<112>

Fig.220)

Fig.7 The refractory alloy elements in the matrix at the

standard heat treated condition.


166 / 298

Fig.8 The amount of Re and Co in the matrix at the

standard heat treated condition.

4.3 PWA1480 2 Ta

Ta Ni3Al 1 mol% Ta

1mol% Al

180 mJm 2 240mJm 2 7,21

22) Ta

22,23)

(SSF: superlattice stacking fault)19)

PWA1480

(Fig.7)

<112>

(Fig.3(c))

5

4 750 /750MPa

(1) Re Mo W

(2) Re

a<112>

1

(3) CMSX4 Re

a<112>

1

(4) LSC-11m

a<112>

(5) Ta <112>

1)

, , 2004, 141-149. 2) Madeleine Durand-Charre, The Microstructure of Superalloys, Gordon and Breach Science Publishers, 1998 3) Z. Zhua, H. Basoaltob, N. Warnkena, R.C. Reed, Acta Mater. 60 (2012) 4888-4900. 4) , Ni

Vol.75 No.7, pp.21-28, (2005). 5) , Vol.62, No.8pp.611-614, (2004),

C.M.F. Rae, N. Matan, R.C. Reed, Materials Science and Engineering A300 (2001) 125 134

) D.Blavette P.Caron T.Khan, Superalloys 1988:proceedings of the Sixth International Symposium on Superalloys,1988, (1988) 305-314.8) F.Diologent, P.Caron, Materials Science and Engineering A385, 2004, 245-257 9) A.Royer, P.Bastie, M.Veron, Acta Materialia, 46, 1998, 5357-5368 10) R. C. Reed, The SUPERALLOYS Fundamentals and Applications (2006) 11) T. Link and M. Feller-Kniepmeier: Metall. Trans. A., 1992, vol. 23A, pp. 99-105. 12) K. Kakehi, Metall. and Mater. Trans. A, 30A, pp.1249-1259 (1999) 13) P.J. Warren, A. Cerezo, G.D.W. Smith, Materials Science and Engineering, A250(1998), 88-92. 14) M. Hattori, N. Goto, Y. Murata1, T. Koyama and M. Morinaga, Materials Transactions, 46(2005), 163-166 15) M. S. A. Karunaratne, P. Carter and R. C. Reed: Mater. Sci. Eng. A281 (2000) 229 233.

N. Tsuno, K. Kakehi, C.M.F. Rae and R. Hashizume, Metallurgical and Materials Transactions, 40A (2009), pp. 269-27217) R.A. Hobbs, G.J. Brewster, C.M.F. Rae and S. Tin, Superalloys2008, 171-180 18) S. Ma, L. Carroll, T.M. Pollock, Acta Materialia, 55(2007), 5802-5812. 19) C.M. Rae, K. Kakehi, R.C. Reed, Proc. 7th Liege Conference on Materials for Advanced Power Engineering 2002, (2002), pp.207-216. 20) H. Murakami, T. Yamagata, H. Harada, M. Yamazaki, Mater. Sci. Eng. A 223 (1997) 54-58. 21 P. Veyssi`ere, G. Saada, Dislocations in Solids, F. Nabarro, M. Duesbery, eds., Vol. 10, Elsevier Sciences, 1996, 314-317. 22) Y. Mishima, S. Ochiai and T. Suzuki, Acta Metallurgica, 33(1985), 1161-1169.

B.H. Kear and D.P. Pope, Superalloys, Supercomposites and Superceramics, J.K. Tien, ed., (San Diego, Academic Press), 1989, 545.


167 / 298

44 2016.10

B-18

Ni

Present Research Status of the Direct and Complete Recycling Method for

Ni-base Single Crystal Superalloy Turbine Airfoils

*Satoshi UTADA, Shinsuke SUZUKI (Waseda Univ.), Hiroshi HARADA, Makoto OSAWA, Kyoko KAWAGISHI,

Toshiharu KOBAYASHI and Tadaharu YOKOKAWA (NIMS)

ABSTRACT

In collaborative work of Waseda University and National Institute for Materials Science (NIMS), we proposed

direct and complete recycle method that enables drastic cost reduction and stable supply of Ni-base single

crystal superalloy turbine components. The method includes collecting the turbine components, remelting and

single crystals. Recycled material showed high-temperature properties equal to or better than genuine ingot

material. The project is scheduled to scale up to commercial ingot size for practical use. This paper introduces

research results and status of our recycling method.

Key words: Ni-base single crystal superalloys, Recycle, Creep, Cyclic oxidation, Turbine airfoil

Ni1) Ni

(

)

(NIMS) Ni

Trent1000

1150°C2)

W Ta Re Ru

NIMS

4 6 Ni 3 5)

JST/ALCA ( :

Advanced Low Carbon Technology Research and

Development Program)

2013-19

4

Ni

2.1

1980 Curwick 6) deBarbadillo7) Papp8)

Papp 1986

5500

70% 20%


168 / 298

8)

Woulds9)

50%

Lamberigts 10) Rupp 11) Prasad 12) Song13) Innola 14) Ni

2.2

Srivastava

Re15)

Stoller 16)

2.3 17)

Ni Re

Lu 18) Ni Co

Fig. 1 Schematic diagram of conventional


169 / 298

3.1

Ni PWA1484

21)

TMS-1700(MGA1700)

PWA1484

10ppm

22 24)

(1)

(2)

(3) (

)

(4)

CaO 25) 100ppm

Ni CaO

8ppm24) CaO

(

)

3.2

2 Ni

PWA148426)

PWA1484

(1) (Genuine)

Al2O3

(2) (Simple recycle)

Al2O3

(3) CA Al2O3

(4) CA+CaO CaO

(

X

)

CA+CaO CA

PWA1484

( 1315°C 4h 1080°C 4h

704°C 24h) 26)

1100°C /137MPa

1100°C 1h 1

3.3

Figure 220)

Al2O3 2

CaO CA+CaO

CaO

Genuine Simple recycle CA CA+CaO0

5

10

Impu

rity

Con

tent

(pp

m)

Materials

S NO S NO

S

N

O

S

NO

Fig. 2 Content of major impurities such as sulfur, oxygen, and

nitrogen for each material. Standard deviations of

each impurity content are shown as error bars.20)


170 / 298

Ni

Fig.319) CA+CaO

TCP

CA+CaO

NIMS-ADP (Alloy

Designing Program, )27)

Fig.4

Fig.520)

( )

Zr Si

Y Ni

CA+CaO

3.4

CaO

Ni

(1)

(2) CaO

( )

0 50 100 150

0

5

10

15

20

103%

CA

Stra

in (

%)

Time /h

Simple recycleCA+CaO

Genuine

70%

Fig. 3 The results of creep test at 1100 °C / 137 MPa. 19)

150 175 200 225

100

125

150

103%

100%

79%

70%

Genuine

CA+CaO

CA

Act

ual C

reep

Rup

utre

Tim

e /h

Predicted Creep Rupture Time /h

Simple recycle

Fig.4 The results of creep test at 1100 °C / 137 MPa and the

relationship between predicted and actual creep rupture

time are shown. 20)

0 20 40 60-0.015

-0.010

-0.005

0.000

0.005 CA+CaO

CA

Simplerecycle

Wei

ght C

hang

e / m

gm

m-2

Number of 1h Oxidation Cycles

Genuine

Fig. 5 Results of cyclic oxidation test up to 50 cycles at

1100°C are shown as weight changes of specimens. 20)


171 / 298

(3)

(4)

Ni

NIMS

NIMS

JST ALCA

2013-19

1)

: :

vol. 43, No. 5 (2015), pp. 349-356. 2) Yokokawa, T., Harada, H., Mori, Y., Kawagishi, K.,

Koizumi, Y., and Kobayashi, T.: Design of Next Generation Ni-base Single Crystal Superalloys Containing Ir: Towards 1150 Temperature Capability, Proceedings of Superalloy 2016, (2016), to be published.

3) Koizumi, Y., Kobayashi, T., Yokokawa, T., Zhang, J., Osawa, M., Harada, H., Aoki, Y., and Arai, M.: Development of next-generation Ni-base single crystal superalloys, Proceedings of Superalloys 2004, (2004), pp. 35 43.

4) Sato, A., Harada, H., Yeh, A.-C., Kawagishi, K., Kobayashi, T., Koizumi, Y., Yokokawa, T., and Zhang, J. X.: A 5th generation SC superalloy with balanced high temperature properties and processability, Proceedings of Superalloys 2008, (2008), pp. 131 138.

5) Kawagishi, K., Yeh, A., Yokokawa, T., Kobayashi, T., Koizumi, Y., and Harada, H.: Development of an Oxidation-resistant High-Strength Sixth-Generation Single-Crystal Superalloy TMS-238, Proceedings of Superalloys 2012, (2012), pp. 9 13.

6) Curwick, L. R., Petersen, W. A., and DeBarbadillo, J. J.: Superalloy Scrap-Generation and Recycling, Proceedings of Superalloys 1980, (1980), pp. 21 30.

7) deBarbadillo, J. J.: Nickel-Base Superalloys; Physical Metallurgy of Recycling, Metallurgical Transactions A, Physical Metallurgy and Materials Science, Vol. 14 A, No. 3 (1983). pp. 329 341.

8) Papp, J. F.: Superalloy Recycling 1976-1986, Proceedings of Superalloys 1988, (1988), pp. 367 376.

9) Woulds, M. J.: Recycling of Engine Serviced Superalloys, Proceedings of Superalloys 1980, (1980), pp. 31 41.

10) Lamberigts, M., Lecomte-Beckers, J., and Drapier, J.-M.: Reverting foundry nickel-base superalloys, Proceedings of International Conference on High Temperature Alloys for Gas Turbine and Other Applications, (1986), pp. 777 786.

11) Rupp, S., Bienvenu, Y., Massol, J., Ablitzer, D., Gardaix, E., Mancha, H., Allibert, M., and Martin-Garin, R.: Laboratory Study on the Recycling of Nickel Base Superalloy Foundry Scrap, Proceedings of International Conference on High Temperature Alloys for Gas Turbine and Other Applications, (1986), pp. 787 797.

12) Prasad, V. V. S., Rao, A. S., Prakash, U., Rao, V. R., Rao, P. K., and Gupt, K. M.: Recycling of superalloy scrap through electro slag remelting, ISIJ International, Vol. 36, No. 12 (1996), pp. 1459 1464.

13) Song, J., Xiao, Q., Zang, C., Xiao, C., Li, S., and Han, Y.: Microstructure and Mechanical Properties of Recycled Ni3Al Base Alloy IC6, Chinese Journal of Aeronautics, Vol. 15, No. 1 (2002), pp. 55 60.

14) Innola, M., Victor, S., and Heorhiy, M.: Refining and recycling of the Nickel based heat-resistant alloys used in Aviation, Proceedings of 3rd International Conference on Life Cycle Management, 2007.

15) Srivastava, R. R., Kim, M., Lee, J., Jha, M. K., and Kim, B.-S.: Resource recycling of superalloys and hydrometallurgical challenges, Journal of Material Science, Vol. 49, No. 14 (2014), pp. 4671 4686.

16) Stoller, V., Olbrich, A., Meese-Marktscheffel, J., Mathy, W., Erb, M., Nietfeld, G., and Gille, G.: Process for electrochemical decomposition of superalloys, Patent, US20080110767 A1, 15-May-2008.

17) :

Vol. 2, No. 2 (2015). 18) Lu, X., Matsubae, K., Nakajima, K., Nakamura, S., and

Nagasaka, T.: Thermodynamic Considerations of Contamination by Alloying Elements of Remelted End-of-Life Nickel- and Cobalt-Based Superalloys, Metallurgical and Materials Transactions B, Vol. 47, No. 3 (2016), pp. 1785 1795.

19) Utada, S., Joh, Y., Osawa, M., Yokokawa, T., Kobayashi, T., Kawagishi, K., Suzuki, S., and Harada, H.: Direct Recycle of Used Single Crystal Superalloy Turbine Blades, Proceedings of the International Gas Turbine Congress 2015 Tokyo, (2015), pp. 1039 1043.

20) Utada, S., Joh, Y., Osawa, M., Yokokawa, T., Kobayashi,


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T., Kawagishi, K., Suzuki, S., and Harada, H.: High Temperature Properties of a Single Crystal Superalloy PWA1484 Directly Recycled after Turbine Blade Use, Proceedings of Superalloys 2016, (2016), to be published.

21) : Ni

40(2012), pp. 181-184.

22) Joh, Y., Kobayashi, T., Yokokawa, T., Kawagishi, K., Osawa, M., Suzuki, S., and Harada, H.: Effects of Sulphur on High Temperature Properties of a Ni-base Single Crystal Superalloy, TMS-1700, Proceedings of10th Liège Conference on Materials for Advanced Power Engineering, (2014), pp. 538 544.

23) Joh, Y., Utada, S., Osawa, M., Kobayashi, T., Yokokawa, T., Kawagishi, K., Suzuki, S., and Harada, H.: Effect of Sulfur on Creep Strength of Ni-Base Single-Crystal Superalloy, TMS-1700, Materials Transactions, Vol. 57, No. 8 (2016), pp. 1305 1308.

24) : 2 Ni

PWA1484 CaO

(2015), p. 71. 25) : Al

Al Vol. 14, (1987), pp. 1684-1960.

26) Cetel, A. D. and Duhl, D. N.: Second-Generation Nickel-Base Single Crystal Superalloy, Proceedings of Superalloys 1988, (1988), pp. 235-244.

27) Harada, H., Ohno, K., Yamagata, T., Yokokawa, T., and Yamazaki, M.: Phase Calculation and its Use in Alloy Design Program for Nickel-Base Superalloys, Proceedings of Superalloys 1988, (1988), pp. 733 742.


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44 2016.10

B-22

Ti-6Al-4V

Creep Properties and Microstructure of Ti-6Al-4V parts build by Additive Manufacturing

*Yuya Sakurai(Tokyo Metropolitan Univ.), Koji KAKEHI(Tokyo Metropolitan Univ.)

ABSTRACT

To investigate possibility of application for aircraft parts, Ti-6Al-4V specimens were build by two types of

additive manufacturing technic and compared with convetional mill-annealed materials. The electron backscatter

diffraction (EBSD) analysis revealed SLM and EBM specimens were found to consist of acicul

-annealed material. Results of tensile test

performed at room temperature and 450 °C revealed the anisotropy of ductility in SLM and EBM specimens was

pronounced at high temperature. SLM and EBM specimens showed superior creep properties than the mill-

annealed specimens at 450 °C/400 MPa.

Key words: Ti-6Al-4V, selective laser melting, electron beam melting, microstructure, creep properties

Ti

1)

Boeing

B-787 Ti 15 %2)

Ti-6Al-4V

2 SLM(selective

laser melting) EBM(electron beam melting)

EOS SLM EOSINT M280

EOS Titanium Ti64 35 mm

SLM

AMS 4911 650 °C /3

h( ) SLM

SLM Arcam

EBM Arcam A2X Arcam Titanium

Ti6Al4V 45 mm

EBM

3.1 mm

( 2.8×3.0×19.6 mm)

0° 90°

(top view)

(side view)

954 °C

/2 h( )+538 °C/4 h( ) (STA)

as-built

STA

(RM) 3.17 mm

450 °C

450 °C/400 MPa

(EBSD)

(SEM)


181 / 298

Fig.1 IPF maps of SLM as-built specimen from (a) top

view, (b) side view.

Fig.2 Lower magnification observation of Fig. 1 from

grain boundary.

3.1

Figs. 1, 2 SLM IPF( )

EBSD

SLM

650 °C

Fig. 1

Fig. 2

Fig.

2(a)

Fig. 2(b)

SLM

-

transus(995 °C 3))

Fig. 3 EBM IPF EBM

Burgers

-

Widmanstätten 4)

Fig.3 Wide area observation of as-built EBM specimen from

-Widmanstätten

Fig.4 (a) IPF map (b) phase map of as-received RM. The

SLM SLM

Fig. 4 RM

IPF

EBSD

7.4 %

1.0 %


182 / 298

Table 1 Tensile properties of SLM, EBM and RM

specimens at room temperature.

YS(MPa) UTS(MPa) EL(%)

SLM0° 1090 1150 8.890° 1020 1160 8.5

SLMSTA

0°90°

11501120

12701280

2.42.6

EBM0° 930 1000 8.390° 930 1000 8.8

RM 930 1010 16.3RM STA 1160 1230 6.1

Table 2 Tensile properties of SLM and EBM

specimens at 450°C.

YS(MPa) UTS(MPa) EL(%)

SLM0° 830 970 11.590° 850 950 8.7

EBM0°90°

510560

610670

13.08.8

3.2

3.2.1

Table 1 SLM EBM RM

RM (EL) 16.3 %

SLM 8.7 % EBM 8.6 % 2

EBM RM

SLM (UTS) 0.2 % (YS)

10 %

3.2.2 450 °C

Table 2 450 °C SLM EBM

SLM 1.3 EBM

1.5 0° 90°

SLM EBM 10 %

450 °C

0.2 %

SLM EBM 1.5

3.2.2 450°C /400MPa

Fig. 5 SLM EBM RM 450 °C

/400 MPa

RM RM STA

Fig.5 (a) Creep strain-time curves and (b) strain rate-Creep

strain curves at 450°C/400 MPa.

RM 20 % 10 %

SLM EBM

450 °C

Fig. 5(b)

RM

SLM

4.1

Table 1, 2

(450 °C)

IN718

IN718

5) Ti-6Al-4V

66)

IN718


183 / 298

4.2 SLM EBM

3.1 SLM EBM

EBM

-Widmanstätten EBM

SLM

EBM

SLM EBM

EBM

0.5 0.8 Tm

( )7)

SLM Ar

EBM

EBM SLM

1. Ti-6Al-4V

SLM

2. EBM -Widmanstätten

SLM EBM

3. 450 °C

SLM EBM

SLM EBM

1.5

4. 450 °C /400 MPa

EBM SLM

1) :

, (2015)2) , :

, , Vol.64, No.2 (2014) pp.28-32.

3) G. Lütjering, J. C. Williams: Titanium, second Edition, (Springer, Berlin Heidelberg, 2007) pp. 34

4) R. Pederson: Licentiate Thesis, Luleå University of Technology (2002)

5) K.Kakehi: Mechanical Engineering, Nikkan Kogyo Shimbun, Vol.64 (2016) pp.76-82

6) : Vol.53 No.1 (2014) pp.11-177) : Vol.3 No.3

(2014) pp.152-157


184 / 298

ABSTRACT

Key words


185 / 298

TMS-238

TMS-238MoIr

CMSX-4


186 / 298

TMS238Single Crystal (SC)

Directional Solidified(DS)

Conventional Cat(CC)

Wrought

TMS-238MoIr

CMSX-4


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C-1

Measurement of Heat Release Fluctuationin Gas Turbine Combustor at High Temperature and Pressure

*Shinichi FUKBA, Yuichiro KIMURA, Yusuke TANAKA,Toshihiko SAITOH, Satoshi TAKIGUCHI, Keijiro SAITOH (MHI)

ABSTRACT

Mitsubishi Hitachi Power Systems, Ltd. (MHPS) has been developing highly-efficient gas turbine combined

cycle (GTCC) power generation with the Research & Innovation Center of Mitsubishi Heavy Industries, Ltd.

(MHI). M501J type, the world's first 1600°C-class gas turbine was developed by utilizing the technology

acquired in the national project “The technology development of high efficiency gas turbine which started

from fiscal 2004. For the next generation GTCC, it is essential to understand the details of unsteady phenomena

of the combustion process, which is the key to enhance the capability of low emissions and reliability under the

condition of high combustion temperature.

This paper presents the evaluation of combustion characteristics of a full-scale gas turbine combustor under a

fully reproduced condition of an actual engine, by means of both the experiment and numerical simulation. The

flame chemiluminescence data from high pressure combustion test are analyzed to understand the flame

structure. In addition, Large Eddy Simulation (LES) are conducted and simulation results are validated with

experimental results.

Key words: Combustor, Flame Chemiluminescence, LES


189 / 298

Table 1

10mm

Table

Fig.1 Fig.2

1700

Fig.3

8

2

7

( ~ )

Fig.3

5 kHz

Fig.1 GT

31ata 550°C

Figure. 1

Figure. 2


190 / 298

Figure. 3

Table 2 ,

Fig.4

7mm

300-340nm

OH* (307nm)3)

50kHz

Figure. 4

Table

Fig.5

(20sec)

LES

FLUENT ver15.0

LES SGS Smagorinsky Lilly

FGM (Flamelet Generated

Manifold)

~ ~

NOx

NOx

FGM


191 / 298

Figure.5

(1)

Fig.6 a) b)

c)

d) e)

a)

b)

c)

Fig.6-e)

Fig.6-b) c)

Fig.6-d)

~

~

45deg

90deg

(2) 4)

Fig.6-d)

ave

)(22

v

Lxff (2)

f x

L

avev 0x


192 / 298

(2)

~

5)

LES

FGM

1700

1700 C )

1), , Vol.52, No.2 (2015)

pp.2-9.2)

, , Vol.52, No.2 (2015) pp.15-22.

3) F.Güthe, D.Guyot, et al.: Chemiluminescence as diagnostic tool in the development of gas turbines, Applied Physics B, Vol.107, Issue.3 (2012) pp.619-636.

4) CFD

53 (2015)pp.106-107.

5) A.Innocenti, A.Andreini, et al.: Numerical analysis of the dynamic flame response and thermo-acoustic stability of a full-annular lean partially-premixed combustor,Proceedings of ASME Turbo Expo 2016,GT2016-57182.


193 / 298

C-2

Reduction of PM emission from an RQL combustor with aviation bio fuel

*Hitoshi FUJIWARA, Mitsumasa MAKIDA, Kazuo SHIMODAIRA, Hideshi YAMADA, Keiichi OKAI, Takuya MIZUNO (Japan Aerospace Exploration Agency),

andMasaharu NAKAMURA(Mitsubishi Heavy Industries)

ABSTRACT

Key words: Hydro-treated ester and fatty acid (HEFA), non-volatile PM, aromatic content

Kerosene( ) HEFA

Table 1 HEFA

Kerosene

Kerosene 2

HEFA


194 / 298


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Makida, M., Yamada, H. and Shimodaira, K., “Detailed Research on Rich-Lean Type Single Sector and full annular combustor for small aircraft engine,” ICAS2014-0628, 2014.


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44 2016 10

C-5 NSP

YSEC

Numerical Analysis of Annular Combustor for NSP Micro Jet Engine

*Hiroto SUZUKI (Niigata Univ.), Koji MATSUBARA and Yusaku MATSUDAIRA(Niigata Univ.)

Hiroyuki KOSHIKIZAWA and Kazuyuki ABE(YSEC), Kakuya IWATA (AIST)

ABSTRACT

NSP (Niigata Sky Project) has been developing less than one hundred horsepower micro jet engine, which had

never been commercialized. This micro jet engine needs to improve fuel evaporation tubes, adopt fuel injectors

and change figures of combustor in order to achieve high efficiency combustion. Therefore, I have been

forwarding this research by numerical analysis. This paper reports numerical analysis model of annular

combustor, adequacies of numerical analysis results and examples of numerical analysis for improving fuel

evaporation tubes. Besides, this report also considers fuel evaporating phenomenon and combustion

phenomenon of evaporating fuel, which are difficult to observe in the experiment. After this, we are planning to

apply this analysis to the study of new combustor, which vibratory and incomplete combustion are less likely to

occur.

Key words: Micro Jet Engine, Combustor, Numerical Analysis, Annular, Fuel Evaporation Tube

100PS

NIIGATA SKY PROJECT 2011

(1) 2015 43

NSP

NSP

ANSYS Fluent ver.16.0

Realizable k-

k-

PDF

6

Fig.1

1/6

Fig.2


203 / 298

Fig.1 1/6 Combustor Model

Fig.2 Air and Fuel Stream

Fig.3

Fig.3 10mm

Fig.3 Close-Up of Fuel Evaporation Tube

( Fuel Flow, Nozzle Exit)

Fig.4 Fig.5

408

8.4mm 131

Fig.4 Computational Grid

Fig.5 Computational Grid in Fuel Evaporation Tube

6 Fig.4

[ ] 1/1

1/6

1/6

(C10H22) 320 [K]


204 / 298

6 [g/s]

10 [μm]

18 [m/s]

79 [%] 21 [%] 300 [K]

480 [g/s]

80

PDF

C10H22

320 [K] 47 [ ]

NSP

4 [rpm]

300 [K] 27 [ ]

Fig.6

Fig.6 Mean Temperature against Mixture Fraction

Fig.7 (A)

(a)

(b)

Fig.7(B) (A)

Fig.8

Fig.7 Cross-Sections for Visualization

Fig.8 Velocity Contour

(a)

(b)[m/s]

0

350

175

88

263

(B)

(A)


205 / 298

Fig.9

556 [m/s]

Fig. 9 Velocity Vectors on cross-section (b)

Fig.10 Fig.11

Fig.12

80

2491 [K]

Fig.10 Mass Fraction of C10H22

Fig.11 Mass Fraction of O2

Fig.12 Temperature contour


206 / 298

Table 1(2)

80

Table 1 Comparison of Simulated Temperatures

with Theory (K) Present Theory Maximum 2491.8 2464.0 Exit 809.8 801.8

Table 1 27.8[K]

8[K]

1/6

NIIGATA SKY PROJECT

1)

50 2A11 (2012)

2) JSME pp

124 (2002)

3)

43

C-3 (2015)


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C-10

Development of the Micro Gas Turbine for 250kW class SOFC-MGT hybrid systems

*Masato Kawano, Keisuke Arai and Yoshinori Yasui (Toyota Turbine and Systems Inc.)

ABSTRACT

This report describes development and a proof evaluation of the micro gas turbine to use for the hybrid system

which coupled micro gas turbine with SOFC(Solid Oxide Fuel Cells). Gas turbine supplies compressed air to

SOFC. Gas turbine generates electricity with the fuel which finished a reaction in SOFC and with the air which

finished a reaction in SOFC. We remodeled the main body of gas turbine to fit SOFC system, and developed the

combustor which could burn with low-calorie fuel from SOFC. We developed the control to link SOFC, and to

control the gas turbine for SOFC.

Key words: Micro Gas Turbine, Solid Oxide Fuel Cells (SOFCs), Hybrid System, Compressor, Combustor,

Control,


233 / 298

Fig.1 Flow diagram of SOFC-MGT hybrid system.


234 / 298

Table 1 Specifications of Micro Gas Turbine

Fig.4 Micro Gas Turbine cross section view.

Table 2 Gas condition from SOFC.

Fig.5 Power decreasing TG051R for SOFC

Fig.3 50kW Class Micro Gas Turbine.


235 / 298

Fig.6 Combustor cross section view

Fig.7 Combustion test Result

Fig.8 Operation mode change


236 / 298

Fig.9 Appearance of model 10 SOFC-MGT

Installed at the Tokyo Gas Co.,Ltd’.s Senju Techno

Station It was provided by MHPS

Fig.10 Long term behavior of model 10 SOFC-MGT

hybrid system installed at Tokyo Gas Co.,Ltd’.s Senju

Techno Station It was provided by MHPS

Fig.11 Experiment to reduce main fuel

Fig.12 Appearance of model 15 SOFC-MGT hybrid

system installed at the Next-Generation Fuel Cell

Research Center(NEXT-FC) in Kyushu University

It was provided by MHPS

Fig.13 MGT installed in SOFC.BOP.


237 / 298


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C-12

Ammonia utilization for gas turbine

*Norihiko IKI(AIST), Osamu KURATA(AIST), Takayuki MATSUNUMA(AIST), Takahiro INOUE(AIST), Taku TSUJIMURA(AIST), Hirohide FURUTANI(AIST), Hideaki

KOBAYASHI(Tohoku Univ.) and Akihiro HAYAKAWA(Tohoku Univ.)

ABSTRACT

Ammonia is one of the candidates of hydrogen carrier. AIST carried out demonstration tests with the aim to

show the potential of ammonia-fired power plant. 50kW class turbine system firing kerosene is selected as a

base model. A standard combustor is replaced by a prototype combustor which enables a bi fuel supply of

kerosene and fuel gas. The gas turbine started firing kerosene and increased its electric power output. After

achievement of stable power output, ammonia gas was started to be supplied and its flow rate increased

gradually. Over 40kW power output was achieved by firing ammonia gas only and over 40kW power output was

also achieved by co-firing methane and ammonia. Ammonia gas supply increases NOx in the exhaust gas

dramatically. However NOx removal equipment via Selective Catalytic Reduction can reduce NOx successfully.

Key words: Micro gas turbine, Ammonia, Combustion, Hydrogen carrier, Methane

Table 1 Properties and combustion characteristics of ammonia

33.4 1.177MPa 30

0.85MPa 20 0.76 g/l 0 0.1013MPa

( ) 0.59 1 25

651

840 930 15 28%

18.6MJ/kg

7cm/s 1/5


244 / 298

Table 2

Gas pressure reducing valve

50kW classgas turbine

NOxremoval

equipmentFuel gas compressor

50kg cylinder (liquefied ammonia)

gas<0.2MPa Gas

<0.5MPa

Bypass

Heating by water

gas>0.35MPa

Fig.1 Temporary test facility of a micro gas firing ammonia gas with kerosene

Fuel gas compressor

50kW classgas turbine

NOxremoval

equipment

1 ton cylinder (liquefied ammonia)

Vaporizer

Gas<0.5MPa

Liquid Pressure reducing valve

gasliquid<0.2MPa<216kg/day

Electric Heating

Fig.2 Test facility of a micro gas turbine firingammonia gas

Fig.4 Replacement of a gas turbine combustor

12

Fig.3 Prototype gas turbine combustor


245 / 298

Pow

er [k

W],

Ker

osen

e[L/

h], N

itrog

en[N

m3 /h

]

Rot

aing

spe

ed

[rpm

]

Power

Kerosene

Rotating speed

Nitrogen

Fig.5 Power output and fuel consumption with nitrogen supply

Fig.6 21kW power generation operation co-firing kerosene and ammonia at 75,000rpm of rotating speed


246 / 298

Fig.7 41.8kW power generation operation firing ammonia at 80,000rpm of rotating speed

Fig.8 41.8kW power generation operation co-firing methane and ammonia at 80,000rpm of rotating speed


247 / 298

Fig.9 NOx emission (41.8kW power generation, 80,000rpm of rotating speed)

Fig.10 Flame observation (39.1kW power generation, 75,000rpm of rotating speed)


248 / 298

1) http://www.takachiho.biz/pdf/NH3.pdf

2) http://www.ube-ind.co.jp/japanese/eco/sds/sds7664-41-7_20160608.pdf

3) Vol.58, No.183 (2016) pp.41-48.

4) Hayakawa, A., Goto, T., Mimoto, R., Kudo, T.,

Kobayashi, H.: Fuel, 159 (2015) pp.98-106

5)

Vol.58, No.183 (2016) pp.41-48.

6) th annual NH3 Fuel

Conference (2014)

7) F th

annual NH3 Fuel Conference (2015)

8) e Firing Kerosene and

-43689,

(2015) pp.1-5

9) Kurata, O. et al,,

icope 2015,

ICOPE-15-1139 (2015) pp.1-7

10)

ASME Turbo Expo 2015, GT2016-56954 (2016) pp.1-6


249 / 298

C-16

Technologies to enhance the efficiency for Shin-Sendai Thermal Power Station Unit No.3 series

*Satoshi TAKANO (Tohoku Electric Power Co., Inc.) ABSTRACT

Tohoku Electric Power Co., Inc. is working diligently to improve Gas Turbine Combined Cycle (GTCC)

efficiency in order to meet society’s requirement reducing environmental impact and supplying inexpensive

stably electric power, since installed Japan’s first GTCC in Higashi-Nigata Thermal Power Station. Moreover,

when decide to install its gas turbine, the company provides the manufacture with feedback about knowledge

that based on deep understanding of the operation and maintenance technologies, to improve design and

reliability of new frame and technology.

Shin-Sendai Thermal Power Station unit No.3 is first installed the M701F5 gas turbine which based on its

predecessor the M701F4 structurally and which incorporates advanced component technologies applied to other

gas turbine model. This unit are planned to achieve the highest level of efficiency in the world (more than

60%LHV), which are adopted the latest gas turbine firing 1500 class turbine inlet temperature.

This paper introduces Shin-Sendai Power Station unit No.3 which achieve higher efficiency and higher

operability.

Key words: Gas Turbine, Combined Cycle, High Efficiency, Turbine Inlet Temperature, Electric Power,

M701F5, Shin-Sendai Thermal Power Station

1970

2

GTCC 3

1090MW 1984

3 TIT

1100 M701D

48%

LHV GTCC

3

1989 6

M701G TIT1450

4-1 805MW

55% LHV

1999

1959

40 1 3

GTCC 4 446MW

2010 7 4

TIT 1440

HN3 : Higashi-Niigata Unit No.3

HN4-1, 2 : Higashi-Niigata Unit No.4-1, No.4-2

S4 : Sendai Unit No.4

SS3 : Shin-Sendai Unit No.3

Fig.1 Historical trend of thermal efficiency


250 / 298

4

TIT 58%(1)

GTCC

GTCC

2016 7

3 490MW×

M701F5

GTCC

MHPS

60.9% LHV 3

Fig.

1

3

1 2 GTCC

24 1 27

12 3-1

28 7 3-2

Table 1 Fig. 2

Table 1 Comparison of Shin-Sendai units

3

LNG 160,000kl×2

Fig. 2 Aerial view of No.3 Shin -Sendai power station

5m

2

2.2 M701F53 4

F4

MHPS TIT 1500 F5

F5

F5

3 F5

4

F4

Fig. 3 Relationship among frames

Item New Unit Original Unit

No.3-1 No.3-2 No.1 No.2

Type of

Generation

Combined-Cycle

(Gas and Steam

Turbine)

Conventional

(Steam Turbine)

Output 490MW 490MW 350MW 600MW

Fuel LNG Heavy

Oil

Heavy Oil

Natural

Gas

Commercial

Operation DEC,2015 JUL,2016 JAN,1971 JUN,1973


251 / 298

Fig. 4 Features of M701F5 gas turbine

(2) Fig. 3, 4

F5 F4

NACA

CDA CDA

F

F G H J

(Fig. 5)

Fig. 5 Comparison between NACA and CDA

1500 GAC (G Air Cooled

4 G1

GAC

1600 J NOx

(Fig. 6)

Fig. 6 Combustor parts among GAC, J and F5

TIT

Fig. 7 F5

TIT 1600 J

19 9

MHPS

Fig. 8

Fig. 7 Turbine blade technologies

Fig. 8

F5 4

3

TBC Thermal Barrier

Coating TBC

TBC

TBC Fig.

9

Fig. 9

Note: as of September 2015

CO2 laser

Cooling Air

Video Camera

Specimen


252 / 298

F5 MHPS

MHPS

2.4

M701F5

3

MHPS

TIT 30

TIT

F5

1600 J TIT

TIT

TIT

Fig. 10

NOx

LCF

F5 1600ºC J

J TIT

F5

NOx

TIT

TIT

Fig. 10

TIT

TIT

Fig. 10 Challenges toward TIT increase

3-1

1500

Fig. 11

Fig. 11 Temporary measurements

IGV

Fig. 12

1

Fig. 13


253 / 298

Fig. 12 Compressor inlet flow

Fig. 13 Pressure dynamics in compressor

F5

GAC J

MHPS

NOx

Fig. 14, 15 and 16

Fig. 14 Combustor pressure dynamics result

Fig. 15 Combustor metal temperature result

Fig. 16 NOx emission result

1600 J

1500 F5

Fig. 17 Turbine row 1 vane metal temperature

1 50

Fig. 17

F5

F4

F4

F4

Fig. 18

4 F4


254 / 298

Fig. 19

Fig. 18 Rotor vibration

Fig. 19 Vibratory stress at turbine exhaust

100

TBC

Fig. 20

Fig. 20 Inspection result after 100% load operation

3

3

F5

1 1

3

60.9 LHV

3


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C-18

Fundamental research and technology development of JAXA AeroenginesAnalysis, instrumentation, and testing technologies for jet engine R&D

*Hisao FUTAMURA, Takeshi YAMAMOTO (JAXA)

ABSTRACT

JAXA Propulsion research unit holds various test facilities relating aircraft propulsion systems, those are kept

good condition for our own research use and for others’ demands. In order to lead developing engine

technologies up to date, we have to introduce advanced technologies for analysis, instrumentation, and testing to

support precise and accurate data production. The current fundamental research activities and facilities

innovation are presented in this paper with research projects and background information.

Key words: Engine testing, Exhaust emission, Engine noise, Numerical analysis, Thermodynamics

5%

JAXA

2014 (2014 6 )1)

( 26 8

)2)

( 1)3) 27 4


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4) U.S. Gulf Coast Kerosene-Type Jet Fuel Spot Price FOB, US$ per gallon, US Energy Information Administration

5) World energy investment 2016, OECD/IEA, 20166) 6.

1944 2 Pp.4827) GAS TABLES, J. H. Keenan, J. Kaye, John Wiley &

sons, 19458) GAS TABLES international version, J. Keenan, J. Chao,

J. Kaye, John wiley & sons, 19839) NIST-JANAF Thermochemical Tables 4th Edition, M.

Chase, JPCRD monograph No.9, NIST, 199810) Conceptual design analysis on the Porpulsion system for

High-speed transport aircrafts, 2005/9/10, ISABE 2005-1294, H. Futamura, R. Yanagi

11), 2011/2/28,

, ,JAXA-RM-10-016,

12)

( 2), 2014/3/1,, ,

JAXA-RM-13-019,13) NASA Glenn Coefficients for Calculating

Thermodynamic Properties of Individual Species, B. McBride, M. Zehe, S. Gordon, NASA/TP-2002-211556, Sept. 2002

14) The GERG-2004 Wide-Range Equation of State for Natural Gases and Other Mixtures, O. Kunz, R. Klimeck, W. Wagner, M. Jaeschke, GERG TM15, 2007

15) The GERG-2008 Wide-Range Equation of State for Natural Gases and Other Mixtures: An Expasion of GERG-2004, O. Kunz, W. Wagner, J. Chemical & engineering data 2012, vol. 57, pp.3032-3091


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=δ

θδ

θ

−−−×

( )

−−

=

−κ

κκ

κκ


270 / 298


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( )

−=−− κκ

−+=− η


272 / 298

Total mass [kg] 1.4Size [mm] 104×400

Maximum thrust [N] 83.4Max rotating speed (Design point) [rpm] 130,000

Max exhaust gas temperature [ ] 800Max fuel flow rate [m /min] Approx. 270

C-20

Research on the chevron nozzles using a small turbojet engine *Gakuyo TAMADA and Toshiaki TSUCHIYA (Kanazawa Institute of Technology)

ABSTRACT

This paper describes a research on a jet noise reduction device, a chevron nozzle. Although chevron nozzles

reduce the jet noise by an effect of promoting the mixing of high velocity jet and external flow, they tend to

provide a thrust loss at the same time. In this research, experiments have been conducted using a small turbo-jet

engine to evaluate the effect of chevron nozzles on the engine performance from both aspects of noise reduction

and thrust loss. In addition, this research has conducted CFD analysis by Large-Eddy Simulation (LES) in order

to investigate effect of the mixing of the jet shear layer by Chevron at the nozzle exit. Based on the results of

CFD analysis, it has been found that the chevron nozzle with straight section leads to a decrease in the

magnitude of vorticity.

Key words: Chevron nozzle, Jet engine, Jet noise, Thrust, Noise reduction device

(Sophia Precision J850

: )

( )

1

1

ECU

2

Fig. 1 Experimental apparatus

Table 1 Engine specification

Fig. 2 Control system of the turbojet engine J850


273 / 298

(Nozzle )

(Nozzle )

1)

3

4 5 2

6

Fig. 3 The tested chevron nozzles

Fig. 4 Chevron shape

Fig. 5 Chevron penetration

Table 2 Geometric details of the chevron nozzles

Fig. 6 Photographs of the five chevron (plus one baseline)

nozzles used in the experiment abcdefghijk

( )

90,000 120,000[rpm]

10,000[rpm]

10 7

Fig. 7 Sound level meter locations

RION NL-22

5

Chevroncount

Chevronangle [deg]

Chevronlength[mm]

Exitdiameter

[mm]

Chevronpenetration

[mm]TYPE1 0 0 0 35 -TYPE2 5 60 6.71 35 0.52TYPE3 5 45 11.62 35 0.91TYPE4 5 30 20.12 35 1.57TYPE5 10 45 5.81 35 0.45TYPE6 20 45 2.91 35 0.23TYPE1 0 0 0 35 -TYPE2 5 60 6.71 35 0TYPE3 5 45 11.62 35 0TYPE4 5 30 20.12 35 0TYPE5 10 45 5.81 35 0TYPE6 20 45 2.91 35 0

Nozzle ID

Nozzle

Nozzle


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8 (Overall sound

pressure level : OASPL)

Near-Field

Far-Field 2

120,000[rpm] 9

120,000[rpm]

(a) A Group (Left: Nozzle Right: Nozzle )

(b) B Group (Left: Nozzle Right: Nozzle )

Fig. 8 Comparison of OASPL (Position )

2.0 2.5[dB]

A

B TYPE3 TYPE5,6

( )

St

(St=0.2

)

(St=1.2 )

(a) FFT analysis results (A Group , Nozzle )

(b) FFT analysis results (A Group , Nozzle )

(c) FFT analysis results (B Group , Nozzle )

(d) FFT analysis results (B Group , Nozzle )

Fig. 9 Comparison of measurement position and

(120,000[rpm])Abcdefghijklmnopqrst

10

11


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11(a),(b)

11(a),(b)

8

TYPE1

12(a),(b)

TYPE1

12



Fig. 10 Comparison of specific thrusts



Fig. 11 Comparison of equilibrium running lines

(a) OASPL (b) Corrected thrust

Fig. 12 Effective outlet area ratio

(a) OASPL (b) Corrected thrust

Fig. 13 Chevron penetration rate

Cor

rect

ed s

peci

fic

thru

st

Cor

rect

ed s

peci

fic

thru

st

Cor

rect

ed s

peci

fic

thru

st

Cor

rect

ed s

peci

fic

thru

st


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13(a),(b) (Chevron

penetration rate)

( )

13

14

Fig.14 Chevron penetration rate vs. Effective outlet area

ratio0123456789012345678901234567

0

LES

Software Cradle SCRYU/Tetra V11

SGS (Subgrid Scale model) DSM(Dynamic

Smagorinsky Model)

2 (Q )

15

(a) Z=0

(b) X=0 (Nozzle exit)

Fig.15 The cut section view of the mesh

(Nozzle inlet)

(Inlet)

( ) (Outlet)

(Nozzle)

(Side)

16


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Fig.16 Cut pattern along jet axis

Fig.17 Contours of the average magnitude of vorticity

(Left:TYPE1 Right:TYPE4)1234567

Fig.18 Iso-surfaces of Q parameter colored by the

magnitude of vorticity (Left:TYPE1 Right:TYPE4)

(a) TYPE1

(b) TYPE4

Fig.19 Magnitude of vorticity contours of the jet shear

layer at X/Dout=0.0~5.01234567890

17,19 TYPE4 TYPE1

18

20

20 0.0 5.0

TYPE1 TYPE4

Fig.20 Magnitude of vorticity (With and without chevrons)

CFD

(1)

(2)

(3)

(4)

(5)

1)

2)

,

53

No.167-1,(2016)

TYPE1 TYPE4

TYPE1 TYPE4

Vorticity


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