Determination of Non-Uniform Heat Transfer Coefficients Around a Solid Aluminum Alloy Complex Shape During Quenching by Using an Inverse Method

2002 ◽  
Author(s):  
Pen-Chung Chen ◽  
Deborah A. Kaminski ◽  
Robert W. Messler
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Inverse Method ◽  
Flux Measurement ◽  
Transient Temperature ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Uniform Heat ◽  
The Cost ◽  
Transfer Problems

Gas turbine systems include complex heat transfer problems. Especially, the cooling efficiency is critical to the operation of gas turbine. In order to achieve the desired cooling condition, one needs to know the distribution of heat transfer on the components; however, the cost to implement a full-scale gas turbine test is tremendous. Therefore, many researchers used simplified models to acquire the test data; certain experiments can provide heat flux measurement, whereas other techniques can measure heat transfer coefficients. The direct measurement of heat transfer coefficients on the surface of components is extremely difficult. In such situations, the inverse method using transient temperature measurements taken within the part can be used to determine heat transfer coefficients. By combining experiments and numerical modeling, this presentation attempts to provide an effective and robust method to determine heat transfer coefficients on the part’s surface during cooling. Though the setting of the present paper is the quenching of a part, the technique presented is proposed for in-service heat load. To characterize the present situation, i.e., non-uniform heat transfer coefficients occurring during quenching, a unique methodology for employing inverse heat conduction was developed to obtain heat transfer coefficients from temperature responses. In conventional inverse approaches, the heat transfer coefficient is assumed to be uniform around the periphery, but this approach sometimes is unrealistic, especially for complex shaped parts. In this study, experimental data were used to find parameters in a heat transfer correlation, rather than to determine the coefficients directly. The resulting analysis provided an improved fit to measurements compared to conventional inverse approaches. The method developed was robust and is extendable to parts of arbitrary shape.

Use or Transient Temperature Measurements to Determine Combustor Liner Heat Transfer Coefficients

1979 ◽  
Author(s):  
Barry Schlein
Keyword(s):  
Heat Transfer ◽  
Convective Heat Transfer ◽  
Gas Turbine ◽  
Convective Heat ◽  
Numerical Solutions ◽  
Temperature Measurements ◽  
Transient Temperature ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Convective Heat Transfer Coefficients

The use of transient temperature measurements to determine convective heat transfer coefficients is extended to the more complex environment of the gas turbine combustion chamber. Numerical solutions to a transient liner wall heat balance demonstrate that the rate of metal temperature with time is exponential and that the rate is only a function of the convective heat transfer coefficients. Data taken during a snap acceleration of a gas turbine confirms the exponential temperature response. The combination of numerical analysis and transient temperature measurements provides an approximate method of determining linear heat transfer coefficients.

Heat Transfer Coefficients and Film Cooling Effectiveness on the Squealer Tip of a Gas Turbine Blade

2003 ◽  
Vol 125 (4) ◽  
pp. 648-657 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Rotor Blade ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio

Experimental investigations were performed to measure the detailed heat transfer coefficients and film cooling effectiveness on the squealer tip of a gas turbine blade in a five-bladed linear cascade. The blade was a two-dimensional model of a first stage gas turbine rotor blade with a profile of the GE-E3 aircraft gas turbine engine rotor blade. The test blade had a squealer (recessed) tip with a 4.22% recess. The blade model was equipped with a single row of film cooling holes on the pressure side near the tip region and the tip surface along the camber line. Hue detection based transient liquid crystals technique was used to measure heat transfer coefficients and film cooling effectiveness. All measurements were done for the three tip gap clearances of 1.0%, 1.5%, and 2.5% of blade span at the two blowing ratios of 1.0 and 2.0. The Reynolds number based on cascade exit velocity and axial chord length was 1.1×106 and the total turning angle of the blade was 97.9 deg. The overall pressure ratio was 1.2 and the inlet and exit Mach numbers were 0.25 and 0.59, respectively. The turbulence intensity level at the cascade inlet was 9.7%. Results showed that the overall heat transfer coefficients increased with increasing tip gap clearance, but decreased with increasing blowing ratio. However, the overall film cooling effectiveness increased with increasing blowing ratio. Results also showed that the overall film cooling effectiveness increased but heat transfer coefficients decreased for the squealer tip when compared to the plane tip at the same tip gap clearance and blowing ratio conditions.

scholarly journals Solving the inverse heat conduction problem using NVLink capable Power architecture

2017 ◽  
Vol 3 ◽  
pp. e138 ◽  
Author(s):  
Sándor Szénási
Keyword(s):  
Heat Transfer ◽  
Data Transfer ◽  
Heat Conduction Problem ◽  
Inverse Heat Conduction Problem ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Inverse Heat Transfer ◽  
Long Time ◽  
History Of ◽  
The Cost

The accurate knowledge of Heat Transfer Coefficients is essential for the design of precise heat transfer operations. The determination of these values requires Inverse Heat Transfer Calculations, which are usually based on heuristic optimisation techniques, like Genetic Algorithms or Particle Swarm Optimisation. The main bottleneck of these heuristics is the high computational demand of the cost function calculation, which is usually based on heat transfer simulations producing the thermal history of the workpiece at given locations. This Direct Heat Transfer Calculation is a well parallelisable process, making it feasible to implement an efficient GPU kernel for this purpose. This paper presents a novel step forward: based on the special requirements of the heuristics solving the inverse problem (executing hundreds of simulations in a parallel fashion at the end of each iteration), it is possible to gain a higher level of parallelism using multiple graphics accelerators. The results show that this implementation (running on 4 GPUs) is about 120 times faster than a traditional CPU implementation using 20 cores. The latest developments of the GPU-based High Power Computations area were also analysed, like the new NVLink connection between the host and the devices, which tries to solve the long time existing data transfer handicap of GPU programming.

scholarly journals Effect of Discrete Ribs on Heat Transfer and Friction Inside Narrow Rectangular Cross Section Cooling Passage of Gas Turbine Blade

2021 ◽  
Vol 10 (6) ◽  
pp. 192-209
Author(s):  
Karthik Krishnaswamy ◽  
◽  
Srikanth Salyan ◽  
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Turbine Blades ◽  
Local Heat Transfer ◽  
Local Heat ◽  
Hydraulic Diameter ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Cooling Passage

The performance of a gas turbine during the service life can be enhanced by cooling the turbine blades efficiently. The objective of this study is to achieve high thermohydraulic performance (THP) inside a cooling passage of a turbine blade having aspect ratio (AR) 1:5 by using discrete W and V-shaped ribs. Hydraulic diameter (Dh) of the cooling passage is 50 mm. Ribs are positioned facing downstream with angle-of-attack (α) of 30° and 45° for discrete W-ribs and discerte V-ribs respectively. The rib profiles with rib height to hydraulic diameter ratio (e/Dh) or blockage ratio 0.06 and pitch (P) 36 mm are tested for Reynolds number (Re) range 30000-75000. Analysis reveals that, area averaged Nusselt numbers of the rib profiles are comparable, with maximum difference of 6% at Re 30000, which is within the limits of uncertainty. Variation of local heat transfer coefficients along the stream exhibited a saw tooth profile, with discrete W-ribs exhibiting higher variations. Along spanwise direction, discrete V-ribs showed larger variations. Maximum variation in local heat transfer coefficients is estimated to be 25%. For experimented Re range, friction loss for discrete W-ribs is higher than discrete-V ribs. Rib profiles exhibited superior heat transfer capabilities. The best Nu/Nuo achieved for discrete Vribs is 3.4 and discrete W-ribs is 3.6. In view of superior heat transfer capabilities, ribs can be deployed in cooling passages near the leading edge, where the temperatures are very high. The best THPo achieved is 3.2 for discrete V-ribs and 3 for discrete W-ribs at Re 30000. The ribs can also enhance the power-toweight ratio as they can produce high thermohydraulic performances for low blockage ratios.

Laminar and Transitional Boundary Layer Structures in Accelerating Flow With Heat Transfer

1986 ◽  
Vol 108 (1) ◽  
pp. 116-123 ◽  
Author(s):  
K. Rued ◽  
S. Wittig
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Gas Turbine ◽  
Turbine Blades ◽  
Pressure Gradients ◽  
Transfer Coefficients ◽  
Gas Turbine Blades ◽  
Heat Transfer Coefficients ◽  
Layer Structures ◽  
Flow Heat

The accurate prediction of heat transfer coefficients on cooled gas turbine blades requires consideration of various influence parameters. The present study continues previous work with special efforts to determine the separate effects of each of several parameters important in turbine flow. Heat transfer and boundary layer measurements were performed along a cooled flat plate with various freestream turbulence levels (Tu = 1.6−11 percent), pressure gradients (k = 0−6 × 10−6), and cooling intensities (Tw/T∞ = 1.0−0.53). Whereas the majority of previously available results were obtained from adiabatic or only slightly heated surfaces, the present study is directed mainly toward application on highly cooled surfaces as found in gas turbine engines.

Using Gradient Heat Flux Measurement to Experimentally Determine Local Heat Transfer Coefficient on Combustion Chamber Surface in a Diesel Engine

2019 ◽  
pp. 87-96
Author(s):  
V.B. Sapozhnikov ◽  
V.Yu. Mityakov ◽  
A.V. Mityakov ◽  
A.V. Vintsarevich ◽  
D.V. Gerasimov
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Combustion Chamber ◽  
Local Heat Transfer ◽  
Flux Measurement ◽  
Local Heat ◽  
Transfer Coefficients ◽  
Engine Operation ◽  
Heat Transfer Coefficients

We used gradient thermometry to determine local heat transfer coefficients on the fire deck surface. We studied two modes of engine operation, that is, motored and fired. We show that the heat transfer coefficient distribution over the fire deck surface is inhomogeneous. Our investigation results may be used to validate existing models of heat transfer in a combustion chamber.

Heat Transfer in a Vane Trailing Edge Passage With Conical Pins and Pin-Turbulator Integrated Configurations

2014 ◽  
Author(s):  
J. Kruekels ◽  
S. Naik ◽  
A. Lerch ◽  
A. Sedlov
Keyword(s):  
Heat Transfer ◽  
Pressure Drop ◽  
Gas Turbine ◽  
Internal Heat ◽  
Transfer Coefficients ◽  
Trailing Edge ◽  
Heat Transfer Coefficients ◽  
Trailing Edges ◽  
Converging Channel ◽  
Internal Heat Transfer

The trailing edge sections of gas turbine vanes and blades are generally subjected to extremely high heat loads due to the combined effects of high external accelerating Mach numbers and gas temperatures. In order to maintain the metal temperatures of these trailing edges to a level, which fulfills the mechanical integrity of the parts, highly efficient cooling of the trailing edges is required without increasing the coolant consumption, as the latter has a detrimental effect on the overall gas turbine performance. In this paper the characteristics of the heat transfer and pressure drop of two novel integrated pin bank configurations were investigated. These include a pin bank with conical pins and a pin bank consisting of cylindrical pins and intersecting broken turbulators. As baseline case, a pin bank with cylindrical pins was studied as well. All investigations were done in a converging channel in order to be consistent with the real part. The heat transfer and pressure drop of all the pin banks were investigated initially with the use of numerical predictions and subsequently in a scaled experimental wind tunnel. The experimental study was conducted for a range of operational Reynolds numbers. The TLC (thermochromic liquid crystal) method was used to measure the detailed heat transfer coefficients in scaled Perspex models representing the various pin bank configurations. Pressure taps were located at several positions within the test sections. Both local and average heat transfer coefficients and pressure loss coefficients were determined. The measured and predicted results showed that the local internal heat transfer coefficient increases in the flow direction. This was due to the flow acceleration in the converging channel. Furthermore, both the broken ribs and the conical pin banks resulted in higher heat transfer coefficients compared with the baseline cylindrical pins. The conical pins produced the highest average internal heat transfer coefficients in contrast to the pins with the broken ribs, though this was also associated with a higher pressure drop.

scholarly journals Verification of Thermal Models of Internally Cooled Gas Turbine Blades

2018 ◽  
Vol 2018 ◽  
pp. 1-10 ◽  
Author(s):  
Igor Shevchenko ◽  
Nikolay Rogalev ◽  
Andrey Rogalev ◽  
Andrey Vegera ◽  
Nikolay Bychkov
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blades ◽  
Heat Removal ◽  
Leading Edge ◽  
Jet Impingement ◽  
Local Heat Transfer ◽  
Transfer Coefficients ◽  
Single Experiment ◽  
Heat Transfer Coefficients

Numerical simulation of temperature field of cooled turbine blades is a required element of gas turbine engine design process. The verification is usually performed on the basis of results of test of full-size blade prototype on a gas-dynamic test bench. A method of calorimetric measurement in a molten metal thermostat for verification of a thermal model of cooled blade is proposed in this paper. The method allows obtaining local values of heat flux in each point of blade surface within a single experiment. The error of determination of local heat transfer coefficients using this method does not exceed 8% for blades with radial channels. An important feature of the method is that the heat load remains unchanged during the experiment and the blade outer surface temperature equals zinc melting point. The verification of thermal-hydraulic model of high-pressure turbine blade with cooling allowing asymmetrical heat removal from pressure and suction sides was carried out using the developed method. An analysis of heat transfer coefficients confirmed the high level of heat transfer in the leading edge, whose value is comparable with jet impingement heat transfer. The maximum of the heat transfer coefficients is shifted from the critical point of the leading edge to the pressure side.

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