Analytical Study on Cooling and Pressure Loss Characteristics of a Return-Flow Steam-Cooled Gas Turbine Rotating Blade

1994 ◽  
Author(s):  
M. Obata ◽  
Hiroshi Taniguchi ◽  
Kazuhiko Kudo
Keyword(s):  
Gas Turbine ◽  
Pressure Loss ◽  
Analytical Study ◽  
Return Flow ◽  
Rotating Blade

Heat Transfer in a Rotating, Blade-Shaped, Two-Pass Cooling Channel With a Variable Aspect Ratio

2021 ◽  
Author(s):  
I-Lun Chen ◽  
Izzet Sahin ◽  
Lesley M. Wright ◽  
Je-Chin Han ◽  
Robert Krewinkel
Keyword(s):  
Heat Transfer ◽  
Aspect Ratio ◽  
Gas Turbine ◽  
Cross Section ◽  
Pressure Loss ◽  
Planar Geometry ◽  
Internal Cooling ◽  
Cooling Channel ◽  
First Passage ◽  
Rotating Blade

Abstract This study features a rotating, blade-shaped, two-pass cooling channel with a variable aspect ratio. Internal cooling passages of modern gas turbine blades closely follow the shape and contour of the airfoils. Therefore, the cross-section and the orientation with respect to rotation varies for each cooling channel. The effect of passage orientation on the heat transfer and pressure loss is investigated by comparing to a planar channel design with a similar geometry. Following the blade cross-section, the first pass of the serpentine channel is angled at 50° from the direction of rotation while the second pass has an orientation angle of 105°. The coolant flows radially outward in the first passage with an aspect ratio (AR) = 4:1. After a 180-degree tip turn, the coolant travels radially inward into the second passage with AR = 2:1. The copper plate method is applied to obtain the regionally-averaged heat transfer coefficients on all the interior walls of the cooling channel. In addition to the smooth surface case, 45° angled ribs with a profiled cross section are also placed on the leading and trailing surfaces in both the passages. The ribs are placed such that P/e = 10 and e/H = 0.16. The Reynolds number varies from 10,000 to 45,000 in the first passage and 16,000 to 73,000 in the second passage. The rotational speed ranges from 0 to 400 rpm, which corresponds to maximum rotation numbers of 0.38 and 0.15 in the first and second passes, respectively. The blade-shaped feature affects the heat transfer and pressure loss in the cooling channels. In the second passage, the heat transfer on the outer wall and trailing surface is higher than the inner wall and leading surface due to flow impingement and the swirling motion induced by the blade-shaped tip turn. The rotational effect on the heat transfer and pressure loss is lower in the blade-shaped design than the planar design due to the feature of angled rotation. The tip wall heat transfer is significantly enhanced by rotation in this study. The overall heat transfer and pressure loss in this study is higher than the planar geometry due to the blade-shaped feature. The heat transfer and pressure loss characteristics from this study provide important information for the gas turbine blade internal cooling designs.

Experimental Study of Combustion on a Micro Gas Turbine Combustor

2008 ◽  
Author(s):  
Feng-Shan Wang ◽  
Wen-Jun Kong ◽  
Bao-Rui Wang
Keyword(s):  
Gas Turbine ◽  
Pressure Loss ◽  
Loss Factor ◽  
Total Pressure ◽  
Combustion Efficiency ◽  
Total Pressure Loss ◽  
Test Facility ◽  
Inlet Temperature ◽  
Micro Gas Turbine ◽  
Oil Fuel

A research program is in development in China as a demonstrator of combined cooling, heating and power system (CCHP). In this program, a micro gas turbine with net electrical output around 100kW is designed and developed. The combustor is designed for natural gas operation and oil fuel operation, respectively. In this paper, a prototype can combustor for the oil fuel was studied by the experiments. In this paper, the combustor was tested using the ambient pressure combustor test facility. The sensors were equipped to measure the combustion performance; the exhaust gas was sampled and analyzed by a gas analyzer device. From the tests and experiments, combustion efficiency, pattern factor at the exit, the surface temperature profile of the outer liner wall, the total pressure loss factor of the combustion chamber with and without burning, and the pollutants emission fraction at the combustor exit were obtained. It is also found that with increasing of the inlet temperature, the combustion efficiency and the total pressure loss factor increased, while the exit pattern factor coefficient reduced. The emissions of CO and unburned hydrogen carbon (UHC) significantly reduced, but the emission of NOx significantly increased.

Film Cooling and Aerodynamic Performance on Multi-Cavity Squealer Tip of a Turbine Blade

2021 ◽  
Author(s):  
Feng Li ◽  
Zhao Liu ◽  
Zhenping Feng
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Cooling Performance ◽  
Pressure Loss Coefficient ◽  
Blade Tip ◽  
Total Pressure Loss Coefficient

Abstract The blade tip region of the shroud-less high-pressure gas turbine is exposed to an extremely operating condition with combined high temperature and high heat transfer coefficient. It is critical to design new tip structures and apply effective cooling method to protect the blade tip. Multi-cavity squealer tip has the potential to reduce the huge thermal loads and improve the aerodynamic performance of the blade tip region. In this paper, numerical simulations were performed to predict the aerothermal performance of the multi-cavity squealer tip in a heavy-duty gas turbine cascade. Different turbulence models were validated by comparing to the experimental data. It was found that results predicted by the shear-stress transport with the γ-Reθ transition model have the best precision. Then, the film cooling performance, the flow field in the tip gap and the leakage losses were presented with several different multi-cavity squealer tip structures, under various coolant to mainstream mass flow ratios (MFR) from 0.05% to 0.15%. The results show that the ribs in the multi-cavity squealer tip could change the flow structure in the tip gap for that they would block the coolant and the leakage flow. In this study, the case with one-cavity (1C) achieves the best film cooling performance under a lower MFR. However, the cases with multi-cavity (2C, 3C, 4C) show higher film cooling effectiveness under a higher MFR of 0.15%, which are 32.6%%, 34.2%% and 41.0% higher than that of the 1C case. For the aerodynamic performance, the case with single-cavity has the largest total pressure loss coefficient in all MFR studied, whereas the case with two-cavity obtains the smallest total pressure loss coefficient, which is 7.6% lower than that of the 1C case.

Build Direction Effects on Additively Manufactured Channels

2015 ◽  
Author(s):  
Jacob C. Snyder ◽  
Curtis K. Stimpson ◽  
Karen A. Thole ◽  
Dominic Mongillo
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Pressure Loss ◽  
Small Scale ◽  
Turbine Cooling ◽  
Gas Turbine Cooling ◽  
Gas Turbine Heat Transfer ◽  
Manufacturing Techniques ◽  
Traditional Manufacturing

With the advances of Direct Metal Laser Sintering (DMLS), also generically referred to as additive manufacturing, novel geometric features of internal channels for gas turbine cooling can be achieved beyond those features using traditional manufacturing techniques. There are many variables, however, in the DMLS process that affect the final quality of the part. Of most interest to gas turbine heat transfer designers are the roughness levels and tolerance levels that can be held for the internal channels. This study investigates the effect of DMLS build direction and channel shape on the pressure loss and heat transfer measurements of small scale channels. Results indicate that differences in pressure loss occur between the test cases with differing channel shapes and build directions, while little change is measured in heat transfer performance.

Probabilistic and Numerical Modelling of a Lobed Mixer at Windmilling Conditions

2013 ◽  
Author(s):  
Maxime Lecoq ◽  
Nicholas Grech ◽  
Pavlos K. Zachos ◽  
Vassilios Pachidis
Keyword(s):  
Flow Field ◽  
Response Surface ◽  
Gas Turbine ◽  
Pressure Loss ◽  
Total Pressure ◽  
Engine Performance ◽  
Operating Conditions ◽  
Total Pressure Loss ◽  
Operating Point ◽  
Mixing Losses

Aero-gas turbine engines with a mixed exhaust configuration offer significant benefits to the cycle efficiency relative to separate exhaust systems, such as increase in gross thrust and a reduction in fan pressure ratio required. A number of military and civil engines have a single mixed exhaust system designed to mix out the bypass and core streams. To reduce mixing losses, the two streams are designed to have similar total pressures. In design point whole engine performance solvers, a mixed exhaust is modelled using simple assumptions; momentum balance and a percentage total pressure loss. However at far off-design conditions such as windmilling and altitude relights, the bypass and core streams have very dissimilar total pressures and momentum, with the flow preferring to pass through the bypass duct, increasing drastically the bypass ratio. Mixing of highly dissimilar coaxial streams leads to complex turbulent flow fields for which the simple assumptions and models used in current performance solvers cease to be valid. The effect on simulation results is significant since the nozzle pressure affects critical aspects such as the fan operating point, and therefore the windmilling shaft speeds and air mass flow rates. This paper presents a numerical study on the performance of a lobed mixer under windmilling conditions. An analysis of the flow field is carried out at various total mixer pressure ratios, identifying the onset and nature of recirculation, the flow field characteristics, and the total pressure loss along the mixer as a function of the operating conditions. The data generated from the numerical simulations is used together with a probabilistic approach to generate a response surface in terms of the mass averaged percentage total pressure loss across the mixer, as a function of the engine operating point. This study offers an improved understanding on the complex flows that arise from mixing of highly dissimilar coaxial flows within an aero-gas turbine mixer environment. The total pressure response surface generated using this approach can be used as look-up data for the engine performance solver to include the effects of such turbulent mixing losses.

Compressor Characteristics in Gas Turbine Performance Modelling

2001 ◽  
Author(s):  
Geoff Jones ◽  
Pericles Pilidis ◽  
Barry Curnock
Keyword(s):  
Gas Turbine ◽  
Pressure Loss ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Performance Model ◽  
Flow Function ◽  
Performance Modelling ◽  
Turbine Engine ◽  
Loss Parameter ◽  
Modelling Methods

The choice of how to represent the performance of the fans and compressors of a gas turbine engine in a whole-engine performance model can be critical to the number of iterations required by the solver or indeed whether the system can be solved. This paper therefore investigates a number of compressor modelling methods and compares their relative merits. Particular attention is given to investigating the ability of the various representations to model the performance far from design point. It is noted that, for low rotational speeds and flows, matching on pressure ratio will produce problems, and that efficiency is a discontinuous function at these conditions. Thus, such traditional representations of compressors are not suitable for investigations of starting or windmilling performance. Matching on pressure ratio, Beta, the Crainic exit flow function and the true exit flow function is investigated. The independent parameters of isentropic efficiency, pressure loss, a modified pressure loss parameter, specific torque, and ideal and actual enthalpy rises are compared. The requirements of the characteristic choice are investigated, with regard to choosing matching variables and ensuring that relationships are smooth and continuous throughout the operating range of the engine.

scholarly journals Design and analysis of flow rectifier of gas turbine flowmeter

2013 ◽  
Vol 17 (5) ◽  
pp. 1504-1507 ◽  
Author(s):  
Zhi-Fei Li ◽  
Zheng Du ◽  
Kai Zhang ◽  
Dong-Sheng Li ◽  
Zhong-Di Su ◽  
...  
Keyword(s):  
Experimental Data ◽  
Flow Field ◽  
Computational Model ◽  
Pressure Distribution ◽  
Gas Turbine ◽  
Turbulence Model ◽  
Pressure Loss ◽  
Three Dimensional ◽  
Turbine Flowmeter ◽  
Good Agreement

Three-dimensional computational model for a gas turbine flowmeter is proposed, and the finite volume based SIMPLEC method and k-? turbulence model are used to obtain the detailed information of flow field in turbine flowmeter, such as velocity and pressure distribution. Comparison between numerical results and experimental data reveals a good agreement. A rectifier with little pressure loss is optimally designed and validated numerically and experimentally.

scholarly journals Aerodynamic Loss Increase due to Individual Film Cooling Injections From Gas Turbine Nozzle Surface

1998 ◽  
Author(s):  
Ryo Kubo ◽  
Fumio Otomo ◽  
Yoshitaka Fukuyama ◽  
Yuhji Nakata
Keyword(s):  
Gas Turbine ◽  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Pressure Loss ◽  
Rate Ratio ◽  
Flow Rate Ratio ◽  
Design Stage ◽  
Pressure Side ◽  
Mass Flow Rate Ratio

A CFD investigation was conducted on the total pressure loss variation for a linear nozzle guide vane cascade of a gas turbine, due to the individual film injections from the leading edge shower head, the suction surface, the pressure surface and the trailing edge slot. The results were compared with those of low speed wind tunnel experiments. A 2-D Navier-Stokes procedure for a 2-D slot injection, which approximated a row of discrete film holes, was performed to clarify the applicable limitation in the pressure loss prediction during an aerodynamic design stage, instead of a costly 3-D procedure for the row of discrete holes. In mass flow rate ratios of injection to main flow from 0% to 1%, the losses computed by the 2-D procedure agreed well with the experimental losses except for the pressure side injection cases. However, as the mass flow rate ratio was increased to 2.5%, the agreement became insufficient. The same tendency was observed in additional 3-D computations more closely modeling the injection hole shapes. The summations of both experimental and computed loss increases due to individual row injections were compared with both experimental and computed loss increases due to all-row injection with the mass flow rate ratio ranging from 0% to 7%. Each summation agreed well with each all-row injection result. Agreement between experimental and calculated results was acceptable. Therefore, the loss due to all-row injections in the design stage can be obtained by the correlations of 2-D calculated losses from individual row injections. To improve more precisely the summation prediction for the losses due to the present all-row injections, extensive research on the prediction for the losses due to the pressure side injection should be carried out.

Fast Response Wall Pressure Measurement as a Means of Gas Turbine Blade Fault Identification

1991 ◽  
Vol 113 (2) ◽  
pp. 269-275 ◽  
Author(s):  
K. Mathioudakis ◽  
A. Papathanasiou ◽  
E. Loukis ◽  
K. Papailiou
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Pressure Field ◽  
Fast Response ◽  
Gas Turbine Blade ◽  
Unsteady Pressure ◽  
Rotating Blade ◽  
Processing Techniques ◽  
Industrial Gas Turbine ◽  
Unsteady Pressure Measurements

The distortions of the pressure field around rotating blades of turbomachinery components due to alterations of their shape can be utilized for the identification of faults related to the blading. Measurement of the unsteady pressure field near the wall provides information on such flow and pressure distortions and can thus be used for diagnostic purposes. An experimental investigation of the compressor rotating blade pressure field of an industrial gas turbine has been undertaken, in order to demonstrate the feasibility of the abovementioned principle. Various realistic gas turbine blade faults have been examined. Application of the appropriate processing techniques demonstrates that unsteady pressure measurements can be used to identify the occurrence of minor blade faults (not traceable by standard techniques) as well as the kind of fault. The proposed methodology has the potential for being incorporated in a computerized engine health monitoring system.

scholarly journals Experimental Prediction of Lean Blowout Limits for 3kW Micro Gas Turbine Combustor fuelled with LPG

2021 ◽  
Vol 13 (1) ◽  
pp. 89-95
Author(s):  
V. KIRUBAKARAN ◽  
David BHATT
Keyword(s):  
Pressure Drop ◽  
Gas Turbine ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Gas Turbine Combustor ◽  
Inlet Velocity ◽  
Lean Blowout ◽  
Micro Gas Turbine ◽  
Experimental Prediction

The Lean Blowout Limit of the combustor is one of the important performance parameters for a gas turbine combustor design. This study aims to predict the total pressure loss and Lean Blowout (LBO) limits of an in-house designed swirl stabilized 3kW can-type micro gas turbine combustor. The experimental prediction of total pressure loss and LBO limits was performed on a designed combustor fuelled with Liquefied Petroleum Gas (LPG) for the combustor inlet velocity ranging from 1.70 m/s to 11 m/s. The results show that the predicted total pressure drop increases with increasing combustor inlet velocity, whereas the LBO equivalence ratio decreases gradually with an increase in combustor inlet velocity. The combustor total pressure drop was found to be negligible; being in the range of 0.002 % to 0.065 % for the measured inlet velocity conditions. These LBO limits predictions will be used to fix the operating boundary conditions of the gas turbine combustor.

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