An Experimental and Numerical Study of Heat Transfer and Pressure Loss in a Rectangular Channel With V-Shaped Ribs

An experimental and numerical study was conducted to determine the thermal performance of V-shaped ribs in a rectangular channel with an aspect ratio of 2:1. Local heat transfer coefficients were measured using the steady state thermochromic liquid crystal technique. Periodic pressure losses were obtained with pressure taps along the smooth channel sidewall. Reynolds numbers from 95,000 to 500,000 were investigated with V-shaped ribs located on one side or on both sides of the test channel. The rib height-to-hydraulic diameter ratios (e/Dh) were 0.0625 and 0.02, and the rib pitch-to-height ratio (P/e) was 10. In addition, all test cases were investigated numerically. The commercial software FLUENT™ was used with a two-layer k-ε turbulence model. Numerically and experimentally obtained data were compared. It was determined that the heat transfer enhancement based on the heat transfer of a smooth wall levels off for Reynolds numbers over 200,000. The introduction of a second ribbed sidewall slightly increased the heat transfer enhancement whereas the pressure penalty was approximately doubled. Diminishing the rib height at high Reynolds numbers had the disadvantage of a slightly decreased heat transfer enhancement, but benefits in a significantly reduced pressure loss. At high Reynolds numbers small-scale ribs in a one-sided ribbed channel were shown to have the best thermal performance.

3D Flow Prediction and Improvement of Holes Arrangement of a Film-Cooled Turbine Blade Using a Feature-Based Jet Model

A feature-based jet model has been proposed for use in 3D CFD prediction of turbine blade film cooling. The goal of the model is to be able to perform computationally efficient flow prediction and optimization of film-cooled turbine blades. The model reproduces in the near hole region the macro flow features of a coolant jet within a Reynolds-Averaged Navier Stokes (RANS) framework. Numerical predictions of the 3D flow through a linear transonic film-cooled turbine cascade are carried out with the model, with a low computational overhead. Different cooling holes arrangement are computed and the prediction accuracy is evaluated versus experimental data. It shown that the present model provides a reasonably good prediction of the adiabatic film-cooling effectiveness and Nusselt number around the blade. A numerical analysis of the interaction of coolant jets issuing from different rows of holes on the blade pressure side is carried out. It is shown that the upward radial migration of the flow due to the passage secondary flow structure has an impact on the spreading of the coolant and the film cooling effectiveness on the blade surface. Based on this result, a new arrangement of the cooling holes for the present case is proposed that leads to a better spanwise covering of the coolant on the blade pressure side surface.

Heat Transfer in an Airfoil Trailing Edge Configuration With Shaped Pedestals Mounted Internal Cooling Channel and Pressure Side Cutback

Described in this paper is an experimental study of heat transfer over a trailing edge configuration preceded with an internal cooling channel of pedestal array. The pedestal array consists of both circular pedestals and oblong shaped blocks. Downstream to the pedestal array, the trailing edge features pressure side cutback partitioned by the oblong shaped blocks. The local heat transfer coefficient over the entire wetted surface in the internal cooling chamber has been determined by using a “hybrid” measurement technique based on transient liquid crystal imaging. The hybrid technique employs the transient conduction model in a semi-infinite solid for resolving the heat transfer coefficient on the endwall surface uncovered by the pedestals. The heat transfer coefficient over a pedestal can be resolved by the lumped capacitance method with an assumption of low Biot number. The overall heat transfer for both the pedestals and endwalls combined shows a significant enhancement compared to the case with thermally developed smooth channel. Near the downstream most section of the suction side, the land, due to pressure side cutback, is exposed to the stream mixed with hot gas and discharged coolant. Both the adiabatic effectiveness and heat transfer coefficient on the land section are characterized by using the transient liquid crystal technique.

Automatic Optimisation of Pre-Swirl Nozzle Design

The objective of the research described here is to develop and demonstrate use of automatic design methods for pre-swirl nozzles. Performance of pre-swirled cooling air delivery systems depends critically on the design of these nozzles which is subject to manufacturing and stress constraints. The best solution may be a compromise between cost and performance. Here it is shown that automatic optimisation using computational fluid dynamics (CFD) to evaluate nozzle performance can be useful in design. A parametric geometric model of a nozzle with appropriate constraints is first defined and the CFD meshing and solution are then automated. The mesh generation is found to be the most delicate task in the whole process. Direct hill climbing (DHC) and response surface model (RSM) optimisation methods have been evaluated. For the test case considered, significant nozzle performance improvements were obtained using both methods, but the RSM model was preferred.

Blade Tip Heat Transfer and Aerodynamics in a Large Scale Turbine Cascade With Moving Endwall

High resolution Nusselt number (Nu) distributions were measured on the blade tip surface of a large, 1.0 meter-chord, low-speed cascade representative of a high-pressure turbine. Data was obtained at a Reynolds number of 4.0 × 105 based on exit velocity and blade axial chord. Tip clearance levels ranged from 0.56% to 1.68% design span or equally from 1% to 3% of blade chord. An infrared camera, looking through the hollow blade, made detailed temperature measurements on a constant heat flux tip surface. The relative motion between the endwall and the blade tip was simulated by a moving belt. The moving belt endwall significantly to shifts the region of high Nusselt number distribution and reduces the overall averaged Nusselt number on the tip surface by up to 13.3%. The addition of a suction side squealer tip significantly reduced local tip heat transfer and resulted in a 32% reduction in averaged Nusselt number. Analysis of pressure measurements on the blade airfoil surface and tip surface along with PIV velocity flow fields in the gap give an understanding of the heat transfer mechanism.

Time-Resolved Heat Transfer Measurements on the Tip Wall of a Ribbed Channel Using a Novel Heat Flux Sensor: Part II — Heat Transfer Results

Measurements using a novel heat flux sensor were performed in an internal ribbed channel representing the internal cooling passages of a gas turbine blade. These measurements allowed for the characterization of heat transfer turbulence levels and unsteadiness not previously available for internal cooling channels. In the study of heat transfer, often the fluctuations can be equally as important as the mean values for understanding the heat loads in a system. In this study comparisons are made between the time-averaged values obtained using this sensor and detailed surface measurements using the transient thermal liquid crystal technique. The time-averaged heat flux sensor and transient TLC results showed very good agreement, validating both methods. Time-resolved measurements were also corroborated with hot film measurements at the wall at the location of the sensor to better clarify the influence of unsteadiness in the velocity field at the wall on fluctuations in the heat flux. These measurements resulted in turbulence intensities of the velocity and heat flux of about 20%. The velocity and heat flux integral length scales were about 60% and 35% of the channel width respectively, resulting in a turbulent Prandtl number of about 1.7 at the wall.

Enhancement of Film Cooling Performance at the Leading Edge of Turbine Blade

To enhance the film cooling performance in the vicinity of the turbine blade leading edge, the flow characteristics of the film-cooled turbine blade have been investigated using a cylindrical body model. The inclination of the cooling holes is along the radius of the cylindrical wall and 20 deg relative to the spanwise direction. Mainstream Reynolds number based on the cylinder diameter was 1.01×105 and 0.69×105, and the mainstream turbulence intensities were about 0.2% in both Reynolds numbers. CO2 was used as coolant to simulate the effect of density ratio of coolant-to-mainstream. Furthermore, the effect of coolant flow rates was studied for various blowing ratios of 0.4, 0.7, 1.1, and 1.4, respectively. In experiment, spatially-resolved temperature distributions along the cylindrical body surface were visualized using infrared thermography (IRT) in conjunction with thermocouples, digital image processing, and in situ calibration procedures. This comparison shows the results generated to be reasonable and physically meaningful. The film cooling effectiveness of current measurement (0.29 mm × 0.33 min per pixel) presents high spatial and temperature resolutions compared to other studies. Results show that the blowing ratio has a strong effect on film cooling effectiveness and the coolant trajectory is sensitive to the blowing ratio. The local spanwise-averaged effectiveness can be improved by locating the first-row holes near the second-row holes.

The Effects of Varying the Combustor-Turbine Gap

To protect hot turbine components, cooler air is bled from the high pressure section of the compressor and routed around the combustor where it is then injected through the turbine surfaces. Some of this high pressure air also leaks through the mating gaps formed between assembled turbine components where these components experience expansions and contractions as the turbine goes through operational cycles. This study presents endwall adiabatic effectiveness levels measured using a scaled up, two-passage turbine vane cascade. The focus of this study is evaluating the effects of thermal expansion and contraction for the combustor-turbine interface. Increasing the mass flow rate for the slot leakage between the combustor and turbine showed increased local adiabatic effectiveness levels while increasing the momentum flux ratio for the slot leakage dictated the coverage area for the cooling. With the mass flow held constant, decreasing the combustor-turbine interface width caused an increase in uniformity of coolant exiting the slot, particularly across the pressure side endwall surface. Increasing the width of the interface had the opposite effect thereby reducing coolant coverage on the endwall surface.

Passive Shroud Cooling Concepts for HP Turbines: Experimental Investigations

The cooling of rotor shrouds in the first stage of a high-pressure turbine requires special attention as flatter turbine inlet temperature profiles and more highly loaded blades result in increased thermal and mechanical stresses. The use of film cooling and/or internal convective cooling makes the rotor shroud heavier and oversized, restricting the maximum rotational speed. Alternative methods are therefore sought to achieve improved cooling of the shroud. This paper discusses the low speed experimental investigation of two ‘passive’ cooling concepts known as ‘rail cooling’ and ‘platform cooling’. It has been shown experimentally that the modified cooling method, namely the platform cooling, substantially improves the rotor shroud coolant distribution in the critical areas whilst employing significantly lower amounts of coolant.

Methods for Reducing a Thermal Finite Element Model Including the Advection Network Contribution

The life monitoring concept needs on-line calculation to evaluate stresses and temperatures on aircraft engine components, in order to asses fatigue damage accumulation and residual life. Due to the amount of computational time required it is not possible for a full finite element model to operate in real time using the on-board CPU. Stresses and temperatures are then evaluated by using simplified algorithms. In the present work Guyan reduction and component mode synthesis have been applied to a thermal finite element model, including the cooling stream flow — the so called advection network — in order to reduce the size of the solving equation system. The appropriate mathematical formulation for the advection network reduction has been developed. Two reduction methods have been performed, discussed and subsequently applied to a thermal finite element model of a real low pressure turbine disk. The reduced system includes both the disk and the correlated fluid network model, simulating turbine secondary air system. The finite element model is axi-symmetric, with constant convective coefficients. Results of time integration for the reduced and the complete models have been compared. Results show that the proposed techniques gives models with a reduced number of degrees of freedom and at the same time good accuracy in temperature calculation. The reduced models are then suitable for real time computation.