scholarly journals A Short-Duration Blowdown Tunnel for Aerodynamic Studies on Gas Turbine Blading

1982 ◽  
Author(s):  
N. C. Baines ◽  
M. L. G. Oldfield ◽  
T. V. Jones ◽  
D. L. Schultz ◽  
P. I. King ◽  
...  
Keyword(s):  
Mach Number ◽  
Gas Turbine ◽  
Large Scale ◽  
Pressure Fluctuations ◽  
Secondary Flows ◽  
Atmospheric Conditions ◽  
Pressure Distributions ◽  
Line Pressure ◽  
Exit Mach Number ◽  
Turbine Blading

A blowdown tunnel has been constructed for testing large-scale models of gas turbine blading to measure pressure distributions and losses, and for fundamental studies of boundary layers and secondary flows. It can accommodate cascades of approximately 100-mm chord, 300-mm span, and sufficient passages to achieve periodicity, or larger blades up to about 300-mm chord with fewer passages. The exit Reynolds number per metre range is 3 × 105 to 6 × 107 at unity exit Mach number, and the maximum exit Mach number is typically 1.45. These limits are well in excess of likely developments in gas turbine technology in the foreseeable future. The tunnel has a running time of 3–5 sec, which means that a small and inexpensive compressor plant is sufficient to supply the system. A pressure regulator produces a constant supply presssure regardless of line pressure fluctuations. The cascade exit pressure is controlled by tandem ejector pumps which produce sub-atmospheric conditions. The cascade Reynolds and Mach numbers may thus be independently and continuously varied.

A Cold Heat Transfer Tunnel for Gas Turbine Research on an Annular Cascade

1993 ◽  
Author(s):  
R. F. Martinez-Botas ◽  
A. J. Main ◽  
G. D. Lock ◽  
T. V. Jones
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Region Of Interest ◽  
Secondary Flows ◽  
Test Duration ◽  
Pressure Distributions ◽  
Exit Mach Number ◽  
Nozzle Guide ◽  
Turbine Design ◽  
Annular Cascade

The Oxford University Blowdown Tunnel has been substantially modified to test a large annular cascade of high pressure nozzle guide vanes (mean blade diameter of 1.11 m and axial chord of 0.0673 m). The new transonic facility has been constructed to obtain complete contours of heat transfer coefficient for both the end walls and blade surfaces using the transient liquid crystal technique, to measure pressure distributions and losses, and to study fundamental aspects of boundary layers and secondary flows. The facility allows an independent variation of Reynolds and Mach numbers, providing aerodynamic and heat transfer measurements in the region of interest for gas turbine design. The mass flow rate through the cascade at NGV design conditions (exit Mach number 0.96 and Reynolds number 2.0 × 106) is 38 kg/s and the pressure-regulated test duration exceeds 7 seconds.

Numerical Modeling of Heat Transfer and Pressure Losses for Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2007 ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. CFD predictions of blade tip heat transfer are compared to test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; they are flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25, 2.0, and 2.75% of blade span. The tip heat transfer results of the numerical models agree fairly well with the data and are comparable to other CFD predictions in the open literature.

Experiments of Compressible Homogeneous and Isotropic Turbulence Interacting With Expansion Waves

2003 ◽  
Author(s):  
Savvas S. Xanthos ◽  
Yiannis Andreopoulos
Keyword(s):  
Mach Number ◽  
Shock Tube ◽  
Total Pressure ◽  
Large Scale ◽  
Isotropic Turbulence ◽  
Pressure Fluctuations ◽  
Incident Shock ◽  
Expansion Waves ◽  
Curvature Effects ◽  
Homogeneous And Isotropic Turbulence

The interaction of traveling expansion waves with grid-generated turbulence was investigated in a large-scale shock tube research facility. The incident shock and the induced flow behind it passed through a rectangular grid, which generated a nearly homogeneous and nearly isotropic turbulent flow. As the shock wave exited the open end of the shock tube, a system of expansion waves was generated which traveled upstream and interacted with the grid-generated turbulence; a type of interaction free from streamline curvature effects, which cause additional effects on turbulence. In this experiment, wall pressure, total pressure and velocity were measured indicating a clear reduction in fluctuations. The incoming flow at Mach number 0.46 was expanded to a flow with Mach number 0.77 by an applied mean shear of 100 s−1. Although the strength of the generated expansion waves was mild, the effect on damping fluctuations on turbulence was clear. A reduction of in the level of total pressure fluctuations by 20 per cent was detected in the present experiments.

Measurements of Secondary Flows Downstream of a Turbine Cascade at Off-Design Incidence

2004 ◽  
Author(s):  
M. W. Benner ◽  
S. A. Sjolander ◽  
S. H. Moustapha
Keyword(s):  
Large Scale ◽  
Prediction Method ◽  
Leading Edge ◽  
Field Measurements ◽  
Inviscid Flow ◽  
Secondary Flows ◽  
Subsequent Paper ◽  
Empirical Prediction ◽  
Pressure Distributions ◽  
Passage Vortex

This paper presents experimental results of the secondary flows from two large-scale, low-speed, linear turbine cascades for which the incidence was varied. The aerofoils for the two cascades were designed for the same inlet and outlet conditions and differed mainly in their leading-edge geometries. Detailed flow field measurements were made upstream and downstream of the cascades and static pressure distributions were measured on the blade surfaces for three different values of incidence: 0, +10 and +20 degrees. The results from this experiment indicate that the strength of the passage vortex does not continue to increase with incidence, as would be expected from inviscid flow theory. The streamwise acceleration within the aerofoil passage seems to play an important role in influencing the strength of the vortex. The most recent off-design secondary loss correlation (Moustapha et al. [1]) includes leading-edge diameter as an influential correlating parameter. The correlation predicts that the secondary losses for the aerofoil with the larger leading-edge diameter are lower at off-design incidence; however, the opposite is observed experimentally. The loss results at high positive incidence have also high-lighted some serious shortcomings with the conventional method of loss decomposition. An empirical prediction method for secondary losses has been developed and will be presented in a subsequent paper.

scholarly journals Conceptual Mean-Line Design of Single and Twin-Shaft Oxy-Fuel Gas Turbine in a Semiclosed Oxy-Fuel Combustion Combined Cycle

2013 ◽  
Vol 135 (8) ◽  
Author(s):  
Majed Sammak ◽  
Egill Thorbergsson ◽  
Tomas Grönstedt ◽  
Magnus Genrup
Keyword(s):  
Mach Number ◽  
Gas Turbine ◽  
Mass Flow ◽  
Rotational Speed ◽  
Gas Turbines ◽  
Combined Cycle ◽  
Fuel Gas ◽  
Power Turbine ◽  
Exit Mach Number ◽  
Line Design

The aim of this study was to compare single- and twin-shaft oxy-fuel gas turbines in a semiclosed oxy-fuel combustion combined cycle (SCOC–CC). This paper discussed the turbomachinery preliminary mean-line design of oxy-fuel compressor and turbine. The conceptual turbine design was performed using the axial through-flow code luax-t, developed at Lund University. A tool for conceptual design of axial compressors developed at Chalmers University was used for the design of the compressor. The modeled SCOC–CC gave a net electrical efficiency of 46% and a net power of 106 MW. The production of 95% pure oxygen and the compression of CO2 reduced the gross efficiency of the SCOC–CC by 10 and 2 percentage points, respectively. The designed oxy-fuel gas turbine had a power of 86 MW. The rotational speed of the single-shaft gas turbine was set to 5200 rpm. The designed turbine had four stages, while the compressor had 18 stages. The turbine exit Mach number was calculated to be 0.6 and the calculated value of AN2 was 40 · 106 rpm2m2. The total calculated cooling mass flow was 25% of the compressor mass flow, or 47 kg/s. The relative tip Mach number of the compressor at the first rotor stage was 1.15. The rotational speed of the twin-shaft gas generator was set to 7200 rpm, while that of the power turbine was set to 4800 rpm. A twin-shaft turbine was designed with five turbine stages to maintain the exit Mach number around 0.5. The twin-shaft turbine required a lower exit Mach number to maintain reasonable diffuser performance. The compressor turbine was designed with two stages while the power turbine had three stages. The study showed that a four-stage twin-shaft turbine produced a high exit Mach number. The calculated value of AN2 was 38 · 106 rpm2m2. The total calculated cooling mass flow was 23% of the compressor mass flow, or 44 kg/s. The compressor was designed with 14 stages. The preliminary design parameters of the turbine and compressor were within established industrial ranges. From the results of this study, it was concluded that both single- and twin-shaft oxy-fuel gas turbines have advantages. The choice of a twin-shaft gas turbine can be motivated by the smaller compressor size and the advantage of greater flexibility in operation, mainly in the off-design mode. However, the advantages of a twin-shaft design must be weighed against the inherent simplicity and low cost of the simple single-shaft design.

Numerical Modeling of Heat Transfer and Pressure Losses for an Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2009 ◽  
Vol 1 (2) ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. Computational fluid dynamics (CFD) predictions of blade tip heat transfer are compared with test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; a flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25%, 2.0%, and 2.75% of the blade span. The tip heat transfer results of the numerical models agree well with data. For the case in which side-by-side comparison with test measurements in the open literature is possible, the magnitude of the heat transfer coefficient in the “sweet spot” matches data exactly and shows 20–50% better agreement with experiment than prior CFD predictions of this same case.

Numerical Investigation of Aerothermal Characteristics of the Blade Tip and Near-Tip Regions of a Transonic Turbine Blade

2015 ◽  
Vol 137 (9) ◽  
Author(s):  
A. Arisi ◽  
S. Xue ◽  
W. F. Ng ◽  
H. K. Moon ◽  
L. Zhang
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Rotor Blade ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Tip Leakage Flow ◽  
Blade Tip ◽  
Exit Mach Number

In modern gas turbine engines, the blade tips and near-tip regions are exposed to high thermal loads caused by the tip leakage flow. The rotor blades are therefore carefully designed to achieve optimum work extraction at engine design conditions without failure. However, very often gas turbine engines operate outside these design conditions which might result in sudden rotor blade failure. Therefore, it is critical that the effect of such off-design turbine blade operation be understood to minimize the risk of failure and optimize rotor blade tip performance. In this study, the effect of varying the exit Mach number on the tip and near-tip heat transfer characteristics was numerically studied by solving the steady Reynolds averaged Navier Stokes (RANS) equation. The study was carried out on a highly loaded flat tip rotor blade with 1% tip gap and at exit Mach numbers of Mexit = 0.85 (Reexit = 9.75 × 105) and Mexit = 1.0 (Reexit = 1.15 × 106) with high freestream turbulence (Tu = 12%). The exit Reynolds number was based on the rotor axial chord. The numerical results provided detailed insight into the flow structure and heat transfer distribution on the tip and near-tip surfaces. On the tip surface, the heat transfer was found to generally increase with exit Mach number due to high turbulence generation in the tip gap and flow reattachment. While increase in exit Mach number generally raises he heat transfer over the whole blade surface, the increase is significantly higher on the near-tip surfaces affected by leakage vortex. Increase in exit Mach number was found to also induce strong flow relaminarization on the pressure side near-tip. On the other hand, the size of the suction surface near-tip region affected by leakage vortex was insensitive to changes in exit Mach number but significant increase in local heat transfer was noted in this region.

scholarly journals Comparison of Visualized Turbine Endwall Secondary Flows and Measured Heat Transfer Patterns

1984 ◽  
Vol 106 (1) ◽  
pp. 168-172 ◽  
Author(s):  
R. E. Gaugler ◽  
L. M. Russell
Keyword(s):  
Heat Transfer ◽  
Large Scale ◽  
Horseshoe Vortex ◽  
Secondary Flows ◽  
Turbine Vane ◽  
A New Technique ◽  
Exit Mach Number ◽  
Visualization Techniques ◽  
Neutrally Buoyant ◽  
First Time

Various flow visualization techniques were used to define the secondary flows near the endwall in a large scale turbine vane cascade. The cascade was scaled up from one used to generate endwall heat transfer data under a joint NASA-USAF contract. A comparison of the visualized flow patterns and the measured Stanton number distributions was made for cases where the inlet Reynolds number and exit Mach number were matched. Flows were visualized by using neutrally buoyant helium-filled soap bubbles, by using smoke from oil soaked cigars, and by a new technique using permanent marker pen ink dots and synthetic wintergreen oil. For the first time, details of the horseshoe vortex and secondary flows can be directly compared with heat transfer distributions. Near the cascade entrance, there is an obvious correlation between the two sets of data, but well into the passage the effect of secondary flow is not as obvious.

The impingement of underexpanded axisymmetric jets on wedges

1976 ◽  
Vol 76 (2) ◽  
pp. 307-336 ◽  
Author(s):  
P. J. Lamont ◽  
B. L. Hunt
Keyword(s):  
Mach Number ◽  
Mach Disk ◽  
Front Face ◽  
Convergent Nozzle ◽  
Free Jet ◽  
The Third ◽  
Axisymmetric Jets ◽  
Pressure Distributions ◽  
Major Factors ◽  
Exit Mach Number

This paper reports an experimental investigation into the impingement of underexpanded axisymmetric jets on each of three wedges arranged symmetrically in the jets. Three jets were used; two were produced by a convergent-divergent nozzle of exit Mach number 2.2 which was operated at underexpansion ratios of 1·2 and 2, while the third jet was produced by a convergent nozzle operated at an underexpansion ratio of 4. The base widths of the wedges equalled the exit diameter of the convergent-divergent nozzle, their apex angles were 90°, 60° and 45° and they were situated within or just downstream of the first cell of each jet.Detailed front-face pressure distributions were obtained for 15 different configurations. Shadowgraph photographs were taken of these and other cases. The results show a variety of possible flow patterns. The major factors determining which flow pattern occurs are the combination of the centre-line Mach number and wedge apex angle and the jet shock strength and position. It was found that observed shock intersections can be reconstructed by means of shock polars, but that a four-shock confluence will not always occur when it is possible according to the shock polars. In one case there is evidence of the existence of a β triple-shock confluence. An interesting and unusual flow involving a stagnation bubble was obtained when the wedges were situated downstream of the free-jet Mach disk.

Effect of Airfoil Shape and Turning Angle on Turbine Airfoil Aerodynamic Performance at Transonic Conditions

2011 ◽  
Author(s):  
Santosh Abraham ◽  
Kapil Panchal ◽  
Srinath V. Ekkad ◽  
Wing Ng ◽  
Barry J. Brown ◽  
...  
Keyword(s):  
Mach Number ◽  
Gas Turbine ◽  
Turbine Blades ◽  
Flow Characteristics ◽  
Airfoil Surface ◽  
Turning Angle ◽  
Gas Turbine Blades ◽  
Mach Numbers ◽  
Exit Mach Number ◽  
Turbine Airfoils

Performance data for high turning gas turbine blades under transonic Mach numbers is significantly lacking in literature. Performance of three gas turbine airfoils with varying turning angles at transonic flow conditions was investigated in this study. Midspan total pressure loss, secondary flow field and static pressure measurements on the airfoil surface in a linear cascade setting were measured. Airfoil curvature and true chord were varied to change the loading vs. chord for each airfoil. Airfoils A, D and E are designed to operate at different velocity triangles. Velocity triangle requirements (inlet/exit Mach number and gas angles) come from 1D and 2D models that include calibrated loss systems. One of the goals of this study was to use the experimental data to confirm/refine loss predictions for the effect of various Mach numbers and gas turning angles. The cascade exit Mach numbers were varied within a range from 0.6 to 1.1. The airfoil turning angle ranges from 120° to 138°. A realistic inlet/exit Mach number ratio, that is representative of that seen in a real engine, was obtained by reducing the inlet span with respect to the exit span of the airfoil, thereby creating a quasi 2D cascade. In order to compare the experimental results and study the detailed flow characteristics, 3D viscous compressible CFD analysis was also carried out.

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