Principal Research Results

Development of the Temperature Prediction Methodfor Hot Gas Path Parts of 1300 ℃ Class Gas Turbines
Background
Electric power companies are working to extend the remaining life of gas turbine (GT) hot gas path parts to reduce maintenance costs in Japan. But there are few effective methods to evaluate the remaining life of GT hot gas path parts quantitatively, so it isimportant to develop the temperature prediction method of GT hot gas path parts to realize the evaluation of them correctly. CRIEPI isdeveloping the NDE (non-destructive evaluation) method to evaluate the thermal barrier performance and the degradation property of
the TBC (Thermal Barrier Coating) which is coated over GT hot gas path parts, and the numerical method based on CFD(Computational Fluid Dynamics), which analyzes flow and thermal field in a gas turbine. Development of NDE method for the performance of the TBC, temperature analysis and analytical life evaluation of a first stage rotor blade of a 1100 ℃ class gas turbine havebeen conducted.

Objectives
The prediction methods of temperature distributions of the combustion liner and the first stage rotor blade for the operating
1300 ℃ class gas turbine. The developed methods shall be verified by comparison between the predicted temperature and actual conditions of ex-service parts.

Principal Results
1. Development of NDE prediction method for temperature distribution of the TBC coated over thecombustion liner
CRIEPI developed the prediction method for the average temperature of TBC (TBC temperature) and the temperature of thesubstrate over the operation term by means of NDE method ＊ 1 and degradation prediction method ＊ 2 for the TBC thermal barrier performance (Fig.1). The predicted temperature distribution on the whole inner surface of the TBC which was used about 15,000h in thermal power plant is shown in Fig.2 by our developed method and NDE apparatus. It is possible to confirm from the result of the predicted TBC temperature distribution that the TBC temperature becomes higher from the center of the combustion liner to the downstream, and some parts show high temperature locally due to the effect of non-uniform cooling of the film cooling air. Using thismethod of predicting TBC temperature makes it possible to predict substrate temperature and to evaluate the remaining life of GT hotgas path parts more correctly.
2. Numerical prediction procedure of temperature profile for a first stage rotor blade
The three-dimensional steady-state numerical analysis was conducted by thermal conjugation of inside and outside fields ofthe blade, which consists of convection heat transfer around the blade, thermal conduction in the blade material and internal blade
cooling. In the present analytical method, cooling structure of the blade was conscientiously modeled and correlations of heat transferand friction in the rib-roughened internal cooling passage derived from Large Eddy Simulation were applied. The predicted temperature profile on the blade wall is in agreement with estimation based on the material of ex-service blades ＊ 6(Fig.3), and the hot spot
predicted by the numerical analysis coincides with the distinctive damages due to oxidation on ex-blades (Fig.4). Accordingly, capability of the present analysis to utilize for the life evaluation of the blade has been verified.

Future Development
On the NDE prediction method for temperature distribution of the TBC coated over the combustion liner, we will hope tomeasure the many actual combustion liners in order to raise its reliability. And with regard to the numerical analysis, transient variations of the blade temperature in start-up and shut-down will be predicted in order to contribute to the life evaluation.
Main Researchers:
Masahiko Morinaga, Ph.D, Research Scientist, Thermal Engineering Sector, Energy Engineering Research Laboratory
Toshihiko Takahashi, Ph.D, Research Scientist, Advanced Power Engineering Sector, Energy Engineering Research Laboratory
References
M.Morinaga, T.Takahashi and T.Fujii, “Development of Actual TBC Exposure Temperature Prediction Method” GT2004-53267,
Turbo Expo 2004, Vienna Austria, 06/14-17/2004
T.Takahashi, K.Watanabe, T.Sakai, and T.Takahashi, 2004, “Numerical Analysis of Heat Transfer for Gas Turbine Hot-gas-path
Parts, - 5th Report, Three Dimensional Conjugate Analysis for 1300 Celsius class First-stage Rotor Blade Temperature -,” CRIEPI
Report., W03011(in Japanese)※ Estimation by means of thickness of the aluminum diffusion layer at the boundary between the coating and the substrate＊1：Morinaga et al., “Development of Nondestructive Inspection Method for The Degradation of Thermal Barrier Coating - ServiceabilityEvaluation of The Developed NDI Method -”, AJTEC2003, pp355, Hawaii, U.S.A., 2003
＊ 2：Fujii et al., “Development of Operating Temperature Prediction Method Using Thermophysical Properties Change of TBC”, ASME Turbo
EXPO 2002, GT-2002-30274, Amsterdam, The Netherlands, 2002
＊ 3：Takahashi et al., “Numerical Analysis of Heat Transfer for Gas Turbine Hot-gas-path Parts, The 3rd report,” CRIEPI Rep.W98006(1999) (inJapanese)＊ 4：Sakai et al., “Damage Assessment for First Stage Blade of a 1100 Celsius class Gas Turbine based on Three dimensional Finite ElementAnalysis, CRIEPI Rep. T01046(2002) (in Japanese)＊ 5：Takahashi et al., “Numerical Analysis of Heat Transfer of Inner Cooling Passage for the First Stage Rotor Blade of 1300 Celsius class Gas
Turbine, The 2nd report,” CRIEPI Rep., W02005(2003) (in Japanese)＊ 6：Okada et al., “Examination of Temperature Estimation and Aluminum-content Prediction for a Gas Turbine Coating,” CRIEPI
Rep.W01022(2002) (in Japanese)6A. Cost reduction and ensuring reliability

Measured by TBC Thermal Barrier
Performance Inspection Apparatus(δ/λλaged)

Circumferential
Position n

Measured TBC
Thickness (δ)

Measured Thermal Conductivityof Aged TBC (λaged)

Circumferential position [°]Operation Time (t )LMP = ( T + 273.15 )･( log10 t + 14 )TBC Operating Temperature [℃]
T = (λaged／λ／λas sprayed + 1.40 ) / 1.25 / (log10t + 14 ) ×104–273.15
Thermocouple Installed Point
Substrate Surface
Temperatureof AirCooling Side

Measured Substrate Thickness

Color of TBC Temperature [℃]360

λaged／λ／λas sprayed = 1.25×10-4LMP –1.40

270180900100

Heat Flux onThermocouple
Installed Point

200300400500

Axial Length from the Combustor Inlet [mm]Substrate Surface Temperatureof TBC Sprayed SideFig.2

Calculating Flow of the TBC Temperature

NDE method and degradation prediction method for thermal barrier
performance of TBC enable to predict the TBC and the Substratetemperature under operation

Predicted result of TBC Temperature

TBC temperature becomes higher from the center of thecombustion liner to the downstream, and some parts
show high temperature locally due to the effect of
non-uniform film cooling.

250
Present analysis
Estimation from ex-service blades
PresentOkada et al.6)

200

Temperature [℃]Temperature [℃ ] (The origin : blade-wall-averaage temperature)

Axial Length from theCombustor Inlet

Film Cooling Air port

Ratio of Aged Thermal Conductivity to VirgineThermal Conductivity

Fig.1

Combustor

100 Span80%
0-100
100

Suction side
Pressure side- 290(a)Predictedtemperature distribution on the blade wall

Span50%(The origin:blade-wall-average temperature)

Thermal Conductivityof Virgin TBC(λas sprayed)

0-100
100

Span20%0Oxidation-100TE

Suction side

LEPressure sideTE

Suction side

LE, TE: Leading edge and Trailing edge

Pressure side(b) Exterior of a ex-service blade (used in the US))
Fig.4 Predicted temperature distribution

Fig.3 Predicted temperature profile on the blade
wall, compared with estimation based on
ex-service bladeson the blade wall and Exterior of an
ex-service blade

Predicted temperature profile is in agreement withestimation based on ex-service blade

Hot spots predicted by numerical analysis coincides
with the distinctive damages on an ex-service blade

7