Heat Transfer Enhancement for Gas Turbine Blade Leading Edge Cooling Using Curved Double-Wall/Vortex Cooling With Various Disturbing Objects

2019 ◽  
Author(s):  
Xiaojun Fan ◽  
Liang Li ◽  
Jiefeng Wang ◽  
Fan Wu
Keyword(s):  
Heat Transfer ◽  
Nusselt Number ◽  
Pressure Loss ◽  
Total Pressure ◽  
Leading Edge ◽  
Total Pressure Loss ◽  
Wall Impingement ◽  
Pin Fins ◽  
Cooling Passage ◽  
Double Wall

Abstract A new double-wall cooling configuration combined with the vortex cooling is established to study the cooling behavior for the gas turbine blade leading edge. This configuration consists of multiple nozzles, a curved inner cooling passage, a row of bridge holes and a curved outer cooling passage with 4 kinds of disturbing objects (namely smooth wall, pin-fins, dimples and protrusions). Numerical simulations are performed based on the 3D viscous steady Reynolds Averaged Navier-Stokes (RANS) equations and the k-ω turbulence model. The cooling behavior of the Double-wall/vortex cooling configuration is compared with the Double-wall/impingement cooling configuration at the same conditions. Generally, the Double-wall/vortex cooling configuration has a better cooling performance. It is found the Nusselt number of the inner surface for the Double-wall/vortex cooling configuration is 46.7% higher. However, the Double-wall/impingement cooling configuration has a smaller friction coefficient and a total pressure loss. Different disturbing objects have significant influences on the heat transfer performance of the outer surface. The Nusselt number of disturbing objects (pin-fins, dimples and protrusions) is much higher than the smooth wall, and the value is 1.27–2.22 times larger. Configuration with protrusions has the highest globally-averaged Nusselt number. For the heat transfer performance of the inner surface and the total pressure loss coefficient, disturbing objects have no obvious influence. As bridge holes row increases, the overall cooling performance is improved. The globally-averaged Nusselt number of the outer target is enhanced while the total pressure loss is reduced.

Conjugate Heat Transfer Characteristics of Double Wall Cooling With Gradient Diameter of Film and Impingement Holes

2021 ◽  
Author(s):  
Juan He ◽  
Qinghua Deng ◽  
Zhenping Feng
Keyword(s):  
Heat Transfer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Flow Direction ◽  
Turbine Blades ◽  
Total Pressure Loss ◽  
Advanced Technique ◽  
Cooling Effectiveness ◽  
Uniform Pattern ◽  
Double Wall

Abstract Double wall cooling, consisting of internal impingement cooling and external film cooling, is believed to be the most advanced technique in modern turbine blades cooling. In this paper, to improve the uniformity of temperature distribution, a flat plate double wall cooling model with gradient diameter of film and impingement holes was proposed, and the heat transfer and flow characteristics were investigated by solving steady three-dimensional Reynolds-Averaged Navier-Stokes (RANS) equations with SST k-ω turbulence model. The influence of gradient diameter on overall cooling effectiveness and total pressure loss was studied by comparing with the uniform pattern at the blowing ratios ranging from 0.5 to 2. For gradient diameter of film hole patterns, results show that −10% film pattern always has the lowest film flow non-uniformity coefficient. The laterally averaged overall cooling effectiveness of uniform pattern lies between that of +10% and −10% film patterns, but the intersection of three patterns moves upstream from the middle of flow direction with the increase of blowing ratio. Therefore, the −10% film pattern exerts the highest area averaged cooling effectiveness, which is improved by up to 1.6% and 1% at BR = 0.5 and 1 respectively compared with a uniform pattern. However, at higher blowing ratios, the +10% film pattern maintains higher cooling effectiveness and lower total pressure loss. For gradient diameter of impingement hole patterns, the intersection of laterally averaged overall cooling effectiveness in three patterns is located near the middle of flow direction under all blowing ratios. The uniform pattern has the highest area averaged cooling effectiveness and the smallest non-uniform coefficient, but the −10% jet pattern has advantages of reducing pressure loss, especially in the laminated loss.

Elliptical Pin Fins as an Alternative to Circular Pin Fins for Gas Turbine Blade Cooling Applications: Part 1 — Endwall Heat Transfer and Total Pressure Loss Characteristics

2001 ◽  
Author(s):  
Oğuz Uzol ◽  
Cengiz Camci
Keyword(s):  
Heat Transfer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Major Axis ◽  
Total Pressure Loss ◽  
Turbine Blade Cooling ◽  
Pin Fin ◽  
Blade Cooling ◽  
Pin Fins ◽  
Pin Fin Arrays

Detailed experimental investigation of the wall heat transfer enhancement and total pressure loss characteristics for two alternative elliptical pin fin arrays is conducted and the results are compared to the conventional circular pin fin arrays. Two different elliptical pin fin geometries with different major axis lengths are tested, both having a minor axis length equal to the circular fin diameter and positioned at zero degrees angle of attack to the free stream flow. The major axis lengths for the two elliptical fins are 1.67 and 2.5 times the circular fin diameter, respectively. The pin fin arrays with H/D = 1.5 are positioned in a staggered 2 row configuration with 3 fins in the first row and 2 fins in the second row with S/D = X/D = 2. Endwall heat transfer and total pressure loss measurements are performed two diameter downstream of the pin fin arrays (X/D = 2) in a rectangular cross-section tunnel with an aspect ratio of 4.8 and for varying Reynolds numbers between 10000 and 47000 based on the inlet velocity and the fin diameter. Liquid Crystal Thermography is used for the measurement of convective heat transfer coefficient distributions on the endwall inside the wake. The results show that the wall heat transfer enhancement capability of the circular pin fin array is about 25–30% higher than the elliptical pin fin arrays in average. However in terms of total pressure loss, the circular pin fin arrays generate 100–200% more pressure loss than the elliptical pin fin arrays. This makes the elliptical fin arrays very promising cooling devices as an alternative to conventional circular pin fin arrays used in gas turbine blade cooling applications.

Measurement of Blade Tip Heat Transfer and Leakage Flow in a Turbine Cascade With a Multi-Cavity Squealer Tip

2013 ◽  
Author(s):  
Jung Shin Park ◽  
Sang Hoon Lee ◽  
Jae Su Kwak ◽  
Won Suk Lee ◽  
Jin Taek Chung
Keyword(s):  
Heat Transfer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Leading Edge ◽  
Total Pressure Loss ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Measurement Results ◽  
Blade Tip ◽  
Cavity Floor

Tip leakage flow induces high heat transfer to the blade tip and causes significant aerodynamic losses. In this paper, we propose a multi-cavity squealer tip with an additional rib in the squealer cavity. Our study investigated the effects of the rib location and shape on the blade tip heat transfer and the total pressure loss. Experiments were performed in a five-bladed linear cascade using a low speed wind tunnel. The blade chord, pitch, and span length were 126mm, 102.7mm, and 160mm, respectively. The Reynolds number, based on the blade chord and cascade exit velocity, was 2.44×105, and a tip clearance of 1.25% of the blade span was considered. The additional rib was installed in the squealer tip cavity near the leading edge, the mid-chord, and the training edge, respectively. The shape of the rib was also varied from rectangular to triangular in order to minimize the rib surface area exposed to the hot gas. The secondary flow and total pressure loss were measured using a seven-hole probe at one-chord downstream of the blade trailing edge, and the heat transfer coefficient distributions were measured by utilizing the hue-detection based transient liquid crystal technique. Flow measurement results indicated that the proposed multi-cavity tip reduced the total pressure loss. The blade tip heat transfer measurement results showed that the proposed multi-cavity tip was able to reduce the maximum heat transfer region near the cavity floor near the leading edge, but the heat transfer on the second cavity floor increased due to the leakage flow reattachment.

Heat Transfer and Pressure Loss Characteristics of Pin-Fins With Different Shapes in a Wide Channel

2017 ◽  
Author(s):  
Jin Xu ◽  
Jiaxu Yao ◽  
Pengfei Su ◽  
Jiang Lei ◽  
Junmei Wu ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Nusselt Number ◽  
Thermal Performance ◽  
Pressure Loss ◽  
Bottom Surface ◽  
Number Range ◽  
Pin Fin ◽  
Nominal Diameter ◽  
Pin Fins

Convective heat transfer enhancement and pressure loss characteristics in a wide rectangular channel (AR = 4) with staggered pin fin arrays are investigated experimentally. Six sets of pin fins with the same nominal diameter (Dn = 8mm) are tested, including: Circular, Elliptic, Oblong, Dropform, NACA and Lancet. The relative spanwise pitch (S/Dn = 2) and streamwise pitch (X/Dn = 4.5) are kept the same for all six sets. Same nominal diameter and arrangement guarantee the same blockage area in the channel for each set. Reynolds number based on channel hydraulic diameter is from 10000 to 70000 with an increment of 10000. Using thermochromic liquid crystal (R40C20W), heat transfer coefficients on bottom surface of the channel are achieved. The obtained friction factor, Nusselt number and overall thermal performance are compared with the previously published data from other groups. The averaged Nusselt number of Circular pin fins is the largest in these six pin fins under different Re. Though Elliptic has a moderate level of Nusselt number, its pressure loss is next to the lowest. Elliptic pin fins have pretty good overall thermal performance in the tested Reynolds number range. When Re>40000, Lancet has a same level of performance as Circular, but its pressure loss is much lower than Circular. These two types are both promising alternative configuration to Circular pin fin used in gas turbine blade.

Second Vane Total Pressure Loss Due to Endwall Iceform Contouring

2008 ◽  
Author(s):  
Ronald S. LaFleur
Keyword(s):  
Heat Transfer ◽  
Secondary Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Drag ◽  
Pressure Coefficient ◽  
Secondary Flows ◽  
Total Pressure Loss ◽  
Surface Shape ◽  
Exit Plane

The iceformation design method generates an endwall contour, altering the secondary flows that produce elevated endwall heat transfer load and total pressure losses. Iceformation is an analog to regions of metal melting where a hot fluid alters the isothermal surface shape of a part as it is maintained by a cooling fluid. The passage flow, heat transfer and geometry evolve together under the constraints of flow and thermal boundary conditions. The iceformation concept is not media dependent and can be used in analogous flows and materials to evolve novel boundary shapes. In the past, this method has been shown to reduce aerodynamic drag and total pressure loss in flows such as diffusers and cylinder/endwall junctures. A prior paper [1] showed that the Reynolds number matched iceform geometry had a 24% lower average endwall heat transfer than the rotationally symmetric endwall geometry of the Energy Efficiency Engine (E3). Comparisons were made between three endwall geometries: the ‘iceform’, the ‘E3’ and the ‘flat’ as a limiting case of the endwall design space. This paper adds to the iceformation design record by reporting the endwall aerodynamic performances. Second vane exit flow velocities and pressures were measured using an automated 2-D traverse of a 1.2 mm diameter five-hole probe. Exit plane maps for the three endwall geometries are presented showing the details of the total pressure coefficient contours and the velocity vectors. The formation of secondary flow vortices is shown in the exit plane and this results in an impact on exit plane total pressure loss distribution, off-design over- and under-turning of the exit flow. The exit plane contours are integrated to form overall measures of the total pressure loss. Relative to the E3 endwall, the iceform endwall has a slightly higher total pressure loss attributed to higher dissipation of the secondary flow within the passage. The iceform endwall has a closer-to-design exit flow pattern than the E3 endwall.

The Influence of Aero-Derivative Combustor Turbulence and Reynolds Number on Vane Aerodynamics Losses, Secondary Flows, and Wake Growth

2006 ◽  
Author(s):  
F. E. Ames ◽  
J. D. Johnson ◽  
N. J. Fiala
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Pressure Loss ◽  
Total Pressure ◽  
Reynolds Numbers ◽  
Secondary Flows ◽  
Total Pressure Loss ◽  
Surface Boundary ◽  
High Turbulence ◽  
Turbulence Condition

Exit surveys detailing total pressure loss, turning angle, and secondary velocities have been acquired for a fully loaded vane profile in a large scale low speed cascade facility. Exit surveys have been taken over a four-to-one range in Reynolds numbers based on exit conditions and for both a low turbulence condition and a high turbulence condition. The high turbulence condition was generated using a mock aero-derivative combustor. Exit loss, angle, and secondary velocity measurements were acquired in the facility using a five-hole cone probe at two stations representing axial chord spacings of 0.25 and 0.50. Substantial differences in the level of losses, distribution of losses, and secondary flow vectors are seen with the different turbulence conditions and at the different Reynolds numbers. The higher turbulence condition produces a significantly broader wake than the low turbulence case and shows a measurable total pressure loss in the region outside the wakes. Generally, total pressure losses are about 0.02 greater for the high turbulence case compared with the low turbulence case primarily due to the state of the suction surface boundary layers. Losses decrease moderately with increasing Reynolds number. Cascade inlet velocity distributions have been previously documented in an endwall heat transfer study of this same geometry. These exit survey measurements support our understanding of the endwall heat transfer distributions, the secondary flows in the passage, and the origin of losses.

Endwall Contouring and Fillet Design for Reducing Losses and Homogenizing the Outflow of a Compressor Cascade

2014 ◽  
Author(s):  
Oliver Reutter ◽  
Stefan Hemmert-Pottmann ◽  
Alexander Hergt ◽  
Eberhard Nicke
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Leading Edge ◽  
Total Pressure Loss ◽  
Wall Region ◽  
Optimized Design ◽  
Compressor Cascade ◽  
Outflow Region ◽  
Endwall Contouring ◽  
Operating Points

The following paper deals with the development of an optimized fillet and an endwall contour for reducing the total pressure loss and for homogenizing the outflow of a highly loaded cascade with a low aspect ratio. The NACA-65 K48 cascade profile without a fillet and without endwall contouring is used as a basis. Optimizations are performed using the DLR in-house tool AutoOpti and the RANS-solver TRACE. Three operating points at an inflow Mach number of 0.67 with different inflow angles are used to secure a wide operating range of the optimized design. At first only a fillet is optimized. The optimized fillet is small at the leading edge and rather high, wide and thick towards the trailing edge. It reduces the total pressure loss and homogenizes the outflow up to a blade height of 20 %. Following this a combined optimization of the endwall and the fillet is performed. The optimized contour leads to the development of a vortex, which changes the secondary flow in such a way, that the corner separation is reduced, which in turn significantly reduces the total pressure loss up to 16 % in the design operating point. The contour in the outflow region leads to a significant homogenization of the outflow in the near wall region.

Influence of a Novel 3D Leading Edge Geometry on the Aerodynamic Performance of Low Pressure Turbine Blade Cascade Vanes

2014 ◽  
Author(s):  
A. Asghar ◽  
W. D. E. Allan ◽  
M. LaViolette ◽  
R. Woodason
Keyword(s):  
Turbine Blade ◽  
Pressure Loss ◽  
Total Pressure ◽  
Separated Flow ◽  
Leading Edge ◽  
Low Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Low Pressure Turbine ◽  
Edge Geometry

This paper addresses the issue of aerodynamic performance of a novel 3D leading edge modification to a reference low pressure turbine blade. An analysis of tubercles found in nature and used in some engineering applications was employed to synthesize new leading edge geometry. A sinusoidal wave-like geometry characterized by wavelength and amplitude was used to modify the leading edge along the span of a 2D profile, rendering a 3D blade shape. The rationale behind using the sinusoidal leading edge was that they induce streamwise vortices at the leading edge which influence the separation behaviour downstream. Surface pressure and total pressure measurements were made in experiments on a cascade rig. These were complemented with computational fluid dynamics studies where flow visualization was also made from numerical results. The tests were carried out at low Reynolds number of 5.5 × 104 on a well-researched profile representative of conventional low pressure turbine profiles. The performance of the new 3D leading edge geometries was compared against the reference blade revealing a downstream shift in separated flow for the LE tubercle blades; however, total pressure loss reduction was not conclusively substantiated for the blade with leading edge tubercles when compared with the performance of the baseline blade. Factors contributing to the total pressure loss are discussed.

Sweeping Jet Film Cooling on a Turbine Vane

2018 ◽  
Author(s):  
Mohammad A. Hossain ◽  
Lucas Agricola ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey P. Bons
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Convective Heat Transfer Coefficient ◽  
Total Pressure Loss ◽  
Pressure Probe ◽  
Blowing Ratio ◽  
Turbine Vane ◽  
Shaped Hole

The cooling performance of sweeping jet film cooling was studied on a turbine vane suction surface in a low-speed linear cascade wind tunnel. The sweeping jet holes consist of fluidic oscillators with an aspect ratio (AR) of unity and a hole spacing of Pd/D = 6. Infrared (IR) thermography was used to estimate the adiabatic film effectiveness at several blowing ratios and two different freestream turbulence levels (Tu = 0.3% and 6.1%). Convective heat transfer coefficient was measured by a transient IR technique, and the net heat flux benefit was calculated. The total pressure loss due to sweeping jet film cooling was characterized by traversing a total pressure probe at the exit plane of the cascade. Tests were performed with a baseline shaped hole (777-shaped hole) for comparison. The sweeping jet hole showed higher adiabatic film effectiveness than the 777-shaped hole in the near hole region. Although the unsteady sweeping action of the jet augments heat transfer, the net positive cooling benefit is higher for sweeping jet holes compared to 777 hole at particular flow conditions. The total pressure loss measurement showed a 12% increase in total pressure loss at a blowing ratio of M = 1.5 for sweeping jet hole while 777-shaped hole showed a 8% total pressure loss increase at the corresponding blowing ratio.

Investigation on Performance of Compressor Cascade with Tubercle Leading Edge Blade

2020 ◽  
Vol 37 (3) ◽  
pp. 295-303 ◽  
Author(s):  
Tu Baofeng ◽  
Zhang Kai ◽  
Hu Jun
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Large Scale ◽  
Design Method ◽  
Leading Edge ◽  
Experimental Tests ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Linear Compressor Cascade

AbstractIn order to improve compressor performance using a new design method, which originates from the fins on a humpback whale, experimental tests and numerical simulations were undertaken to investigate the influence of the tubercle leading edge on the aerodynamic performance of a linear compressor cascade with a NACA 65–010 airfoil. The results demonstrate that the tubercle leading edge can improve the aerodynamic performance of the cascade in the post-stall region by reducing total pressure loss, with a slight increase in total pressure loss in the pre-stall region. The tubercles on the leading edge of the blades cause the flow to migrate from the peak to the valley on the blade surface around the tubercle leading edge by the butterfly flow. The tubercle leading edge generates the vortices similar to those created by vortex generators, splitting the large-scale separation region into multiple smaller regions.

Sign in / Sign up

Export Citation Format

Share Document