Question

consider the turbine blade tip available in the laboratory of University and discuss the needs of cooling technology specially specially to be used for the blood tree using photo and 77 develop a mathematical model and show the effect of wearing the dimensions of the blood type on its performance discuss also the effect of film cooling and unsteady turbulence something pulling film hall and gauge arrangement to be made and enhancement in the cooling techniques which we can implement Jet impingement of multiple jets effect then pin fin cooling been array and partial length in arrangement to be made effect of pension orientation pin fin dimple cooling to be used on the reap tabulated calling and also discuss the f fusion cooling techniques of gas turbine

Answer 1

Introduction to Cooling

Gas turbines play a vital role in the today’s industrialized society, and as the demands for power increase, the power output and thermal efficiency of gas turbines must also increase. One method of increasing both the power output and thermal efficiency of the engine is to increase the inlet temperature of the gas entering the turbine. In the advanced gas turbines of today, the turbine inlet temperature can be as high as 1500°C; however, this temperature exceeds the melting temperature of the metal airfoils. The heat transferred to the turbine blade is substantially increased as the turbine inlet temperature is continuously increased. Thus, it is very important to cool the turbine blades for a long durability and safe operation. Cooling the blade must include cooling of the key regions being exposed to the hot gas. The blade tip region is such a critical area and is indeed difficult to cool. This results from the tip clearance gap where the complex tip leakage flow occurs and thereby local high heat loads prevail. Cooling air around 650°C is extracted from the compressor and passes through the airfoils. With the hot gases and cooling air, the temperature of the blades can be lowered to approximately 1000°C, which is permissible for reliable operation of the engine.

It is widely accepted that the life of a turbine blade can be reduced by half if the temperature prediction of the metal blade is off by only 30°C. In order to avoid premature failure, designers must accurately predict the local heat transfer coefficients and local airfoil metal temperatures. By preventing local hot spots, the life of the turbine blades and vanes will increase. However, due to the complex flow around the airfoils it is difficult for designers to accurately predict the metal temperature. At the leading edge of the vane, the heat transfer coefficients are very high. As the flow splits and travels along the vane, the heat fluxes decreases. Along the suction side of the vane, the flow transitions from laminar to turbulent, and the heat transfer coefficients increase. As the flow accelerates along the pressure surface, the heat transfer coefficients also increase. The trends are similar for the turbine blade. The heat flux at the leading edge is very high and continuously decreases as the flow travels along the blade. On the suction surface, the flow transitions from laminar to turbulent, and the heat flux sharply increases.

Types of Cooling

There are three major types for cooling of gas turbine blades. These are explained as follows,

• Convection Cooling

It works by passing cooling air through passages internal to the blade. Heat is transferred by conduction through the blade, and then by convection into the air flowing inside of the blade. A large internal surface area is desirable for this method, so the cooling paths tend to be serpentine and full of small fins. A variation of convection cooling is impingement cooling. It works by hitting the inner surface of the blade with high velocity air. This allows more heat to be transferred by convection than regular convection cooling does. Impingement cooling is often used on certain areas of a turbine blade, like the leading edge, with standard convection cooling used in the rest of the blade.

• Film Cooling

The second major type of cooling is film cooling. This type of cooling works by pumping cool air out of the blade through small holes in the blade. This air creates a thin layer (the film) of cool air on the surface of the blade, protecting it from the high temperature air. The air holes can be in many different blade locations, but they are most often along the leading edge. One consideration with film cooling is that injecting the cooler bleed into the flow reduces turbine efficiency. That drop in efficiency also increases as the amount of cooling flow increases. The drop in efficiency, however, is usually reduces the increase in overall performance produced by the higher turbine temperature.

• Transpiration cooling

This is third major type of cooling, is similar to film cooling, in this it creates a thin film of cooling air on the blade. But it is different in that air is leaked through a porous shell rather than injected through holes. This type of cooling is effective at high temperatures as it uniformly covers the entire blade with cool air.

Needs of Cooling Technology for Blade Tip

To satisfy the fast developments of advanced gas turbines, the operating temperature must be increased to improve the thermal efficiency and output work of the gas turbine engine. However, the heat transferred to the turbine blade is substantially increased as the turbine inlet temperature is continuously increased. Thus, it is very important to cool the turbine blades for a long durability and safe operation. Due to an unavoidable gap clearance between the blade tip and casing, the hot gas flowing through the gap results in a large thermal load on the blade tip. The potential damage due to the large heat load will lead to blade oxidation. Hence, the blade tip is a key region that needs cooling. The turbine blades are cooled by the use of extracted air from the compressor of the gas turbine. This extraction results in a reduction of the thermodynamic efficiency and power output. Too little coolant flow results in high blade temperature. If a proper cooling system is designed, the gain from high firing temperature is so significant that it can outweigh the losses in the efficiency and power output, and offset the complexity and cost of the cooling technology. The turbine blade tip and near-tip regions are difficult to cool and are subjected to potential damage because of the high heat load caused by tip leakage flow. A common way to cool the tip is to extract the cooling air from the internal coolant passages through some film holes that are located on the blade surface discretely. This cooling is known as film cooling. The relatively cool air passes these holes and forms a thin protective film to protect the tip surface from the highly hot mainstream. A high and uniform cooling effectiveness will ensure overall performance of the blade surface cooling. In general, a higher blowing ratio at a specific temperature ratio gives a higher film cooling performance, and thereby the heat is transferred to the blade surface and hence the protection of surface is improved. It is important to optimize the amount of coolant for film cooling at the engine operating conditions. For a better cooling performance, it is necessary to study the film cooling holes pattern which affect the film cooling performance. As the turbine inlet temperature increases, the heat transferred to the turbine blade also increases. The level and variation in the temperature within the blade material, which cause thermal stresses, must be limited to achieve reasonable durability. The operating temperatures are far above the permissible metal temperatures. Therefore, there is a critical need to cool the blades for safe operation. The blades are cooled with extracted air from the compressor of the engine. Since this extraction incurs a penalty on the thermal efficiency and power output of the engine, it is important to understand and optimize the cooling technology for a given turbine blade geometry under engine operating conditions.

Fig. 2.1 shows the common cooling technology with three major internal cooling zones in a turbine blade with strategic film cooling in the leading edge, pressure and suction surfaces, and blade tip region. The leading edge is cooled by jet impingement with film cooling, the middle portion is cooled by serpentine rib-roughened passages with local film cooling and the trailing edge is cooled by pin fins with trailing edge injection. The blade tip internal cooling mostly focuses on the rotational effects inside the turbine blade with heat transfer passage. Also on the unsteady high free-stream turbulence effects on the turbine blade film cooling performance with standard and shaped film-hole geometry. The details of these techniques are discussed in chapter 3.

Comparison of various tip cooling configurations and their effects on film effectiveness and heat transfer coefficients were presented. Above figure shows the clearance gap and tip film cooling configuration. Four film cooling configurations were tested :- (1) discrete slot injection, (2) round hole injection, (3) pressure side flared hole injection, and (4) grooved-tip cavity injection. These four configurations are as shown in figure 2.3 given below.

It was found that for case 4 the overall film cooling performance varied significantly with injection locations and that among the plane-tip injections the discrete slot injection provided better performance than the others.

As per review by B. Sunden and G. Xie [1] numerically study of various film holes configurations on plane and squealer tips of a turbine blade has been done. From this three configurations were tested: (1) the camber arrangement, (2) the upstream arrangement, and (3) the two rows arrangement. The effects of rotation were also observed. They found that at high blowing ratios the latter two cases provided better film cooling performance on the plane and squealer tips than the former one. Higher blowing ratios resulted in a higher cooling effectiveness on the shroud for all cases. They also found that rotation decreased the plane-tip film cooling effectiveness while it slightly affected the squealer-tip film cooling due to the large cavity depth.

Film Cooling

The turbine blade tip and near-tip regions are difficult to cool and are subjected to potential damage because of the high heat load caused by tip leakage flow. A common way to cool the tip is to extract the cooling air from the internal coolant passages through some film holes that are located on the blade surface discretely. This cooling is known as film cooling. In film cooling, the coolant air is discharged through small holes in the turbine surface to form a protective film between the turbine blade and the hot combustor discharge gas. The relatively cool air passes these holes and forms a thin protective layer/film to protect the tip surface from the highly hot mainstream. Figure 2.4 depicts the film cooling concept.

The performance of the film cooling depends on the coolant-to-hot mainstream pressure ratio, temperature ratio, the holes location and configuration. A high and uniform cooling effectiveness will ensure overall performance of the blade surface cooling. A higher blowing ratio at a specific temperature ratio gives a higher film cooling performance, and thereby the heat is transferred to the blade surface and hence the protection of surface is improved. A very high a blowing ratio leads to jet penetration into the mainstream resulting in a reduced cooling performance. While very small a blowing ratio does not force enough coolant to cover the hot surface. Thus, it is important to optimize the amount of coolant for film cooling at the engine operating conditions. For a better cooling performance, it is necessary to study the film cooling holes pattern, e.g., shape, angle, location, and distribution, which affect the film cooling performance. The effect of wake passing on the showerhead film cooling performance of a turbine blade has been investigated experimentally [2]. The experiments were performed in an annular turbine cascade with an upstream rotating row of cylindrical rods. Nickel thin-film gauges were used to determine local film effectiveness and Nusselt number values for various injectants, blowing ratios, and Strouhal numbers. Results indicated a reduction in film effectiveness with increasing Strouhal number, as well as the expected increase in film effectiveness with blowing ratio. The main effect of the wake passing was a reduction in effectiveness caused by enhanced film mixing, and the shock passing effect was found to produce large fluctuations in the heat transfer rate. Wake passing from upstream blade rows causes periodic fluctuations in the magnitude and relative direction of the flow velocity in downstream blade rows. However, due to the complexity of the unsteady flow field, the design of turbine film cooling schemes has tended to rely on steady databases. Because of its importance in turbine design, there has been much investigation into the behaviour of turbine film cooling flows. James D. Heidmann and Barbara L. Lucci found that an increase in wake passing frequency causes a decrease in film effectiveness over most of the blade surface for all cases considered. They used a rotating wheel wake generator with cylindrical bars upstream of a showerhead cooled blunt body. Heated air was used as the injectant. Increasing wake passing frequency was found to reduce film effectiveness, especially at lower blowing ratios where the influence of the wake on the low momentum film is strongest. The wake effect was reduced as free-stream turbulence increased. It was found that the presence of film injection causes earlier boundary layer transition on the suction surface. Heat transfer coefficients were increased slightly with wake passing, but film effectiveness values were greatly reduced. They performed a similar experiment with the addition of trailing edge coolant ejection from the wake-producing bars. The addition of wake coolant was found to have a relatively small effect on downstream blade heat transfer coefficient, but to appreciably reduce leading edge film effectiveness below the wake case with no coolant ejection. [2]

Effects of Unsteady Turbulence on Film Cooling

• Unsteady Effect on Turbine Blade Film Cooling

Recent studies focus on combustion generated high turbulence effect on turbine stator heat transfer with or without film cooling. Recent studies also focus on providing highly detailed heat transfer coefficient and film cooling effectiveness distributions on turbine blades under unsteady high turbulence flow conditions. One important finding from these studies is that unsteady high turbulent flows do not dramatically change heat transfer coefficient distributions on the film-cooled blade, but significantly reduce the film cooling effectiveness [3]. To optimize the film cooling performance, the effects of film hole size, length, spacing, shape, and orientation on turbine blade surface heat transfer distributions need to be considered. Results show that the shaped film-cooling hole provides better film cooling performance than the standard cylindrical film-cooling hole [2]. They used the rotating spoked wheel-type wake generators to simulate the effect of unsteady wake on downstream stationary blade film cooling performance. They used a transient liquid crystal technique to measure the detailed heat transfer coefficient and film cooling effectiveness distributions on the same film-cooled blade from the same unsteady wake simulation facility [2]. Authors concluded that heat transfer coefficients for a film-cooled blade are much higher compared to a blade without film injection. In particular, film injection causes earlier boundary layer transition on the suction surface of the blade. Unsteady wakes only slightly enhance heat transfer coefficients but significantly reduce film-cooling effectiveness on a film-cooled blade [2]. This is because the heat transfer coefficients on a film cooled blade are already very high due to high turbulence and mixing caused by jet interaction with the mainstream.

• Film Hole Shape Effect on Turbine Blade Film Cooling

To improve cooling effectiveness, one solution is to contour the film-hole geometry. Film cooling holes with a diffuser shaped expansion at the exit portion of the holes are believed to improve the film cooling performance on a gas turbine blade. The increased cross-sectional area at the film hole exit compared to a standard cylindrical hole leads to the reduction of the coolant jet velocity for a given blowing ratio. The momentum flux of the jet exiting the shaped hole and the penetration of the jet into the mainstream will be reduced, which results in an increased film cooling effectiveness. Also lateral expansion of the hole provides an improved lateral spreading of the jet, which leads to a better lateral film cooling coverage of the blade. The film cooling hole pattern is presented in Fig. 2.5,

A few previous studies have shown that expanding the exit of the cooling hole improves film-cooling performance compared to a standard cylindrical hole. For example, Authors compared a forward-expanded hole to a cylindrical hole, both of them having compound angle injection for the flat-plate film cooling. They found that there is a larger lateral spreading of the forward expanded hole even though the spatially averaged effectiveness is the same for both cases [3]. They also studied detailed measurements of the flat-plate film cooling effectiveness and heat transfer coefficients downstream of a single film-cooling hole by using an IR camera method. Film cooling holes with and without exit expansions are tested and compared under steady and unsteady wake flow conditions. Detailed heat transfer coefficient and film cooling effectiveness distributions downstream of the injection were measured by using a transient liquid crystal image technique. They found that both fan-shaped and laidback fan-shaped holes have much lower heat transfer coefficients right after the film injection location when compared with cylindrical holes under the same unsteady wake flow conditions. They have almost the same boundary-layer transition location as the cylindrical film hole case, but their heat transfer coefficients are higher after transition into the turbulent region. In general, fan shaped holes provide better film cooling effectiveness than laidback fan-shaped holes and consequently much better film cooling effectiveness than cylindrical holes. Both fan-shaped holes and laidback fan-shaped holes provide lower span-wise averaged heat flux ratio and thus better thermal protection over the blade surface, especially under unsteady wake flow conditions[6].

Apart from external film cooling the blade tip region, a number of serpentine passages can be used as channels for internal coolant air to cool the blade. These cooling passages wind through the blade but are not limited to a simple straight channel. After the potential hot spots on the airfoil surface are identified, the internal cooling schemes can be developed. Many techniques have been developed to enhance the heat transfer in these passages. The cooling passages located in the middle of the airfoils are often lined with rib tabulators. Near the leading edge of the blade, jet impingement (coupled with film cooling) is commonly used. Jet impingement is also used throughout the cross-section of the stator vanes. Pin-fins and dimples can be used in the trailing edge portion of the vanes and blades. These techniques have also been combined to further increase the heat transfer from the airfoil walls. A typical cooled turbine vane is shown in Fig. 3.1.  

As shown in the fig. 3.1 the vane is hollow, so cooling air can pass through the vane internally. The coolant is extracted from the internal channel for impingement and pin-fin cooling. Jet impingement is a very aggressive cooling technique which very effectively removes heat from the vane wall. However, this technique is not readily applied to the narrow trailing edge. The vane trailing edge is cooled using pin-fins (an array of short cylinders). The pin-fins increase the heat transfer area while effectively mixing the coolant air to lower the wall temperature of the vanes. After impinging on the walls of the airfoil, the coolant exits the vane and provides a protective film on the vane’s external surface. Similarly, the coolant travelling through the pin-fin array is ejected from the trailing edge of the airfoil.

Impingement Cooling

Impingement cooling is commonly used near the leading edge of the airfoils, where the heat loads are the greatest. With the cooling jets impinging the blade wall, the leading edge is well suited for impingement cooling because of the relatively thick blade wall in this area. Impingement can also be used near the mid-chord of the vane. Figure 3.1 shows jet impingement located throughout the cross-section of an inlet guide vane. The effect of jet-hole size and distribution, cooling channel cross-section and target surface shape all have significant effects on the heat transfer coefficient distribution. Jet impingement near the mid-chord of the blade is very similar to impingement on a flat plate [5].

• Jet Impingement from Multiple Jets

As shown in figure 3.1, many jets are used to increase the heat transfer from the vane wall. It has been shown that multiple jets perform very differently from a single jet striking a target surface. They concluded that for multiple jets, the Nusselt number is strongly dependent on the Reynolds number, while there is no significant dependence on the jet-to-target plate spacing. The difference is due to the jet cross-flow from the spent jets. Studies by authors showed that the mass from one jet moves in the cross-jet flow direction, and this flow can alter the performance of neighbouring jets [5].

The cross-flow attempts to deflect a jet away from its impinging location on the target plate. In situations with very strong cross-flow and sufficiently large jet-to-target plate spacing, the cross-flow can completely deflect the jet away from the impingement surface. They reported that cross-flow decreases the overall heat transfer from the impingement surface. They determined that cross-flow enhances the convective heat transfer, but the enhancement from the jets decreases, as the jets are deflected. Controlling the direction of the cross-flow they obtained detailed distributions of the heat transfer coefficients for three target plates. Their results clearly indicate when the cross-flow travels in two opposite directions; the heat transfer enhancement on the target plate is much greater than when the cross-flow is restricted to one direction [5]. The heat transfer enhancement on the target plate decreases near the edges due to the decreased coolant flow for film cooling. A. Schulz and H.-J. Bauer [4] investigated cross-flow through a confined space; they also considered cross-flow travelling in one direction and two directions. This study also concluded that increasing cross-flow results in degraded heat transfer; however, the heat transfer coefficient distribution is much more uniform [4]. A typical test model used by author is shown in Fig. 3.2,

As shown in this fig. the coolant jets impinge on the target surface from the jet plate in an inline array. As the coolant travels along the test surface, the spent air from the upstream jets effects the heat transfer coefficient distributions of the downstream jets, and this effect increases as more spent air accumulates on the target surface.

• Jet Impingement on a Curved Surface

The above studies investigated the heat transfer on flat target plates. The results obtained for flat plates can be applied to impingement near the mid-chord of the blade. However, the effect of target surface curvature must be considered when implementing jet impingement near the leading edge of the airfoil. The curvature of the airfoil creates different cross-flow behaviour, and therefore, the heat transfer coefficients on the curved surface are different than those on the flat surface. Author studied impingement on a curved surface, and this group concluded that the average Nusselt number ratio increases as the curvature of the target plate increases. The effect of target surface shape was also pursued. They concluded that a sharper nose radius yields a more uniform Nusselt number distribution compared to a smooth-nosed chamber. This study was also extended to include the effect of coolant extraction for film cooling. [5].

Pin-Fin Cooling

Due to manufacturing constraints in the very narrow trailing edge of the blade, pin-fin cooling is typically used to enhance the heat transfer from the blade wall in this region. As the coolant flows past the pin, the flow separates and wakes are shed downstream of the pin. In addition to this wake formation, a horseshoe vortex forms just upstream of the base of the pin, and the vortex wraps around the pins. This horseshoe vortex creates additional mixing, and thus enhanced heat transfer. Many factors must be considered when investigating pin-fin cooling. The type of pin-fin array and the spacing of the pins in the array effect the heat transfer distribution in the channel. The pin size and shape also have a profound impact on the heat transfer in the cooling passage.

• Pin Array and Partial Length Pin Arrangement

  There are two array structures commonly used. One is the inline array and the other is the staggered array. Figure 3.3 shows a typical experimental test model with a staggered array of pin-fins.

A closer spaced array shows a higher heat transfer coefficient. Their observations clearly indicate that addition of pin-fins significantly enhances the heat transfer coefficient. However, the addition of pins also increases the pressure drop in the flow channel. They showed that the heat transfer coefficient on the pin surface for both arrays is consistently higher than that of the channel end wall. The pin surface heat transfer is observed to be 10 to 20 percent higher for the presented case. Author studied the effects of partial length pins in a rectangular channel. The surface containing pins is not affected by the pin tip clearance. Whereas the opposite surface, that does not have pins, shows a decrease in heat transfer coefficient with an increase in the pin tip clearance. The friction factor is lower for partial pins compared to full-length pins. In general, the heat transfer coefficient decreases with partial length pins.

• Effect of Pin Shape and Array Orientation

Metzger [5] studied the effects of pin shape and array orientations. They reported the effect of flow incident angle on oblong pins. All incident angles except 90 yield higher Nusselt numbers than circular pins. The increase in the friction factor is associated with the flow turning caused by oblong pins. The pin shapes mostly studied are straight cylinders. However, the casting or other manufacturing processes cannot make perfect cylinders and these manufacturing imperfections may affect the heat transfer performance. Straight cylinders in staggered array formation have the highest heat transfer followed by filleted cylinders in the staggered formation. It has been noticed that the fillet cylinder inline formation has better heat transfer than the straight cylinders in the inline formation. Though a staggered array gives higher heat transfer coefficients, performance of the inline straight cylinders is best among the group and the fillet cylinders in staggered formation are the worst. In a different experimental work the effect of stepped diameters on mass transfer coefficients has been studied. The diameter of the pin is axially varied. The base diameter was greater than the centre diameter and no fillet radius is provided. The array configuration was staggered. Results showed that the mass transfer increases or remains the same compared to a straight cylinder pin array when the radius is varied, but the pressure drop reduces significantly for the stepped diameter cylindrical pins. By the use of cube and diamond shaped pins to enhance the heat transfer coefficient from a surface it is observed that the cube-shaped pins have the highest mass transfer coefficients among the shapes considered and round pins have the lowest mass transfer coefficients [5]. Corresponding pressure loss coefficients are higher for the cube and diamond shaped pins relative to the circular pins.

• Pin-Fin Cooling With Ejection

The trailing edge pin-fin channel normally has ejection holes through which the spent coolant exhausts to the main stream flow. They investigated the effects of the length of coolant ejection holes on the heat transfer coefficient in pin-fins. The length of the ejection hole can significantly alter the discharge rate of coolant. More coolant ejection reduces the Nusselt number significantly from no ejection. This decrease in the heat transfer coefficient can be explained by the fact that coolant mass is extracted from the coolant channel before its cooling capacity is fully utilized. Results indicate that the correlation based on the local Reynolds number can predict the heat transfer coefficient distribution for lower coolant ejection but does not adequately predict the heat transfer coefficients at higher ejection rates. They also found that increasing the ejection degrades the endwall heat transfer near the tall wall opposite of the ejection, and the heat transfer on the channel endwall surface near the ejection holes is increased. They also concluded that square, diamond, and circular pin-fin arrays enhance the heat transfer equally in channels with large ejection flows.

Dimple Cooling

In recent years, dimples have been considered as an alternative to pin-fin cooling. Dimpled cooling is a very desirable alternative due to the relatively low pressure loss penalty and moderate heat transfer enhancement. A typical test section for dimple cooling studies is shown in figure 3.4; this figure also shows the dimple induced secondary flow.

  These concave dimples induce flow separation and reattachment with pairs of vortices. The areas of high heat transfer include the areas of flow reattachment on the flat surface immediately downstream of the dimple. The heat transfer in the dimpled channel is typically 2 to 2.5 times greater than the heat transfer in a smooth channel with a pressure loss penalty of 2 to 4 times that of a smooth channel. These values show little dependence on Reynolds number and channel aspect ratio. However, the dimple size, dimple depth (depth-to-print diameter ratio = 0.1 to 0.3), distribution, and shape (cylindrical, hemispheric, teardrop) each effect the heat transfer distribution in the channel. Recent studies have investigated the influence of these factors on the heat transfer in rectangular channels. Dimples have also been investigated in a circular channel and similar levels of heat transfer enhancement and frictional losses were measured. They compared the heat transfer enhancement due a single dimple on both flat and curved surfaces. From this study it was shown that the surface curvature significantly influences the heat transfer enhancement [3]. The heat transfer is further enhanced on a surface that is concavely shaped (compared to a flat surface); however, a convexly curved surface with a dimple decreases the level of heat transfer enhancement.

Rib Tabulated Cooling

Apart from external film cooling the blade tip region, a number of serpentine passages can be used as channels for internal coolant air to cool the blade. These cooling passages wind through the blade but are not limited to a simple straight channel. A common serpentine passage may consist of a first pass, a sharp 180⁰ turn, and a second pass. A typical serpentine passage is schematically shows in Figure 3.5.

The coolant flows radially outward from the hub and then turns 180⁰ and travels radially inward from the tip to the hub. Also, rib tabulators might be mounted on the leading or trailing walls to enhance the heat transfer between the blade wall and coolant. The flow field in the turn/bend is very complex. The rotation alters the flow and hence the heat transfer coefficient distribution, the rotation effect should be considered. Rib tabulators are the most frequently used method to enhance the heat transfer in the internal serpentine cooling passages. The rib turbulence promoters are typically cast on two opposite walls of the cooling passage. Fundamental studies have been conducted to understand the coolant flow through a stationary ribbed channel. The studies show as the coolant passes over a rib oriented 90° to the mainstream flow, the flow near the channel wall separates. Reattachment follows the separation, and the boundary layer reattaches to the channel wall; this thinner, reattached boundary layer results in increased heat transfer coefficients in the ribbed channel. This rib induced secondary flow is shown in Fig 3.6. C

With this additional mixing, the heat is more effectively dissipated from the wall, and thus additional heat transfer enhancement. Because only the flow near the wall of the cooling channel is disturbed by the ribs, the pressure drop penalty by ribs affordable. Because ribs are the most common heat transfer enhancement technique for the serpentine cooling passages, many studies have been conducted to study the effects of channel cross-section, rib configuration, and coolant flow Reynolds number. As shown in figure 3.6, the aspect ratio of the channels changes from the leading to the trailing edge of the blade. Near the leading edge of the blade, the channel may have an aspect ratio around ¼, but near the trailing edge, much broader channels are present with aspect ratios around 4

Rib Induced Secondary Flow [5].

Je With angled ribs performing superior to orthogonal ribs, many researchers have extended their studies to include a wide variety of rib configurations. Han et al. showed that V-shaped ribs (Fig. 3.7) outperform the angled ribs; for a given pressure drop, the V shaped ribs give more heat transfer enhancement. Numerous other studies have shown the same conclusion that V-shaped ribs perform better than the traditional angled ribs in a variety of channels and flow conditions. In an effort to further increase the heat transfer performance of the rib turbulators, discrete rib configurations were introduced. Fig. 3.7 shows discrete (or broken) ribs are similar to the traditional ribs, but they are broken in one or more locations, so the rib is not continuous. In the majority of cooling channels, discrete ribs were shown to outperform the continuous angled.

The majority of ribs used in experimental studies have a square cross-section; however, studies have investigated the heat transfer enhancement of various profiled ribs. Delta-shaped ribs were also studied and these ribs were also shown to result in higher heat transfer enhancement than the traditional angled ribs [5]. They investigated the performance of ribs leaning into or away from the flow. They concluded the traditional square ribs give greater heat transfer enhancement and less frictional losses than the ribs leaning into or away from the flow. When the blades are cast, the ribs are unlikely to have sharp edges as the previous studies have considered. The ribs are likely to have rounded edges, and this was taken into account.

• Rotational Effect on Rib Tabulated Cooling

Rotation induces Carioles and centrifugal forces that produce cross-stream secondary flow in the rotating coolant passages; therefore, heat transfer coefficients in rotor coolant passages are very much different from those in non rotating frames. Heat transfer in rotating coolant passages is very different from that in stationary coolant passages. Both Carioles and rotating buoyancy forces alter the flow and temperature profiles in the rotor coolant passages and affect their surface heat transfer coefficient distributions. It is very important to determine the local heat transfer distributions in the rotor coolant passages under typical engine cooling flow, coolant-to-blade temperature difference and rotating conditions. Effects of coolant passage cross-section and orientation on rotating heat transfer are also important. Since the direction of the Coriolis force is dependent on the direction of rotation and flow, the Coriolis force acts in different directions in the two-passes. For radial outward flow, the Coriolis force shifts the core flow towards the trailing wall. If both the trailing and leading walls are symmetrically heated, then the faster moving coolant near the trailing wall would be cooler than the slower moving coolant near the leading wall. Rotational buoyancy is caused by a strong centrifugal force that pushes cooler heavier fluid away from the center of rotation. In the first channel rotational buoyancy affects the flow in a similar fashion as the Coriolis force and causes a further increase in flow and heat transfer near the trailing wall of the first channel; whereas, the Coriolis force favors the leading side of the second channel. The rotational buoyancy in the second channel tries to make the flow distribution more uniform in the duct.

• Rotational Effect on Jet Impingement Cooling

Several studies were mentioned above that investigate jet impingement cooling. Although a number of impingement studies have been completed, only a few studies consider the effect of rotation on impingement cooling. Epstein et al. [5] studied the effect of rotation on impingement cooling in the leading edge of a blade. They reported that the rotation decreases the impingement heat transfer, but the effective heat transfer is better than a smooth rotating channel. The zero staggered cooling jets show lower heat transfer coefficients compared to that with a staggered angle. They reported the effect of rotation on the leading edge impingement cooling by using the naphthalene sublimation technique. Their experiment did not include the rotating buoyancy effect. The jet direction has an offset angle with respect to the rotation direction. They found that the rotation decreases the impingement heat transfer for all staggered angles. The effect of rotation is least when jet direction has an angle of 45° to rotation direction

However, a maximum of 40% reduction in heat transfer is noted when jet direction is perpendicular to rotation direction. This may be because the Coriolis force creates a swirl action on the spent flow and also deflects the jet when jet direction is parallel to rotation direction. They studied the effect of rotation on swirling impingement cooling in the leading edge of a blade. They found that screw-shaped swirl cooling can significantly improve the heat-transfer coefficient over a smooth channel and the improvement is not significantly dependent on the temperature ratio and rotational forces. Parsons et al. [5] studied the effect of rotation on impingement cooling in the mid-chord region of the blade. A central chamber serves as the pressure chamber, and jets are released in either direction to impinge on two heated surfaces. The jet impinging directions have different orientations with respect to the direction of rotation. They reported that the rotation decreases the impingement heat transfer on both leading and trailing surfaces with more effect on the trailing side up to 20% heat transfer reduction. Akella and Han [6] studied the effect of rotation on impingement cooling for a two-pass impingement channel configuration with smooth walls. The difference that spent jets from the trailing channel is used as cooling jets for the leading channel. Therefore, the cross-flow in the trailing side is radial outward; for the leading side, it is radial inward. They reported that irrespective of the direction of rotation, the heat transfer coefficient on the first-pass and second-pass impinging wall decreases up to 20% in the presence of rotation.. They also found that the rotation decreases impingement heat transfer on the first-pass and second-pass ribbed wall and extended their earlier work to include heat transfer on a smooth target wall with film coolant extraction from a channel with four heated walls.

• Rotational Effect on Pin-Fin Cooling

Pin-fin cooling has been investigated for many years, but only recently has the effect of rotation been considered in channels with pin-fins. Recently, Willett and Bergles [7] studied the effect of rotation on heat transfer in narrow rectangular channels with smooth and with typical pin-fin array, respectively, including channel orientation effect with respect to the plane of rotation. They found that the heat transfer enhancement in the pin-fin channel due to rotation and buoyancy was less than the enhancement in the smooth channel. They showed that heat transfer enhancement mainly is due to pin-fin flow disturbance; pin-fins significantly reduce the effect of rotation, but they do not eliminate the effect. Wright et al. studied the effect of rotation on heat transfer in narrow rectangular channels with typical pin-fin array used in turbine blade trailing edge design and oriented at 150° with respect to the plane of rotation. Results show that turbulent heat transfer in a stationary pin-fin channel can be enhanced up to 3.8 times that of a smooth channel; rotation enhances the heat transferred from the pin-fin channels up to 1.5 times that of the stationary pin-fin channels. Most importantly, for narrow rectangular pin-fin channels oriented at 150° with respect to the plane of rotation, heat transfer enhancement on both the leading and trailing surfaces increases with rotation. This provides positive information for the cooling designers.

Effusive cooling techniques of gas turbines

Numerous experimental and numerical studies have shown the potential deriving from the adoption of transpiration cooling for advanced cooling problems. The main drawback of pure film cooling is actually that the heat sink potential of the cooling air is not effectively utilized. Consequently attention is focused on wall cooling schemes that make more efficient use of cooling air, allowing the designer more latitude in overall secondary flow management. Fig 3.9 shows the comparison of different cooling techniques with generation of film and flow path.

The transpiration cooling is obtained by means of porous walls. They combine two heat exchange effects which are the convective one at the cooler surface and through the wall and the film one at the warmer surface where hot and cold gases mix. The mixing produces both an increase of the heat exchange coefficient by convection, dependent on the fluids flow features, and a drop of the gas temperature at the surface. The net effect is however a decrease of the heat fluxes from the hot gas to the component’s surface. Although they have been extensively used to demonstrate the suitability of porous materials and evaluate their thermodynamic efficiencies, problems occurring when the component both thermal and mechanical stress have limited the application of this technology to the turbo machinery field has to cope with both thermal and mechanical stress have limited the application of this technology to the turbomachinery field.

In effusion cooling systems, the inner convective heat transfer and the heat shielding process are strongly related in that an intense heat extraction by convection through the holes reduces the ability of the cooling air to lessen the heat transfer between the hot gas and the blade’s wall, due to its higher temperature when it is injected into the main flow. Recently, a new technology called Poroform has been presented. It allows manufacturing components having walls equipped with micro holes, obtained by galvanic electroforming techniques. The fabrication process allows the fabrication of effusive systems with the required distribution of hole diameter and density on the cooled surface, to achieve isothermal conditions for large component areas. According to this technology, the component is replicated onto a matrix serving as a cathode for the galvanic bath. On the matrix, dielectric prefabricated areas prevent the metallic growth during the deposition process, thus ending as holes on the electroformed wall. Fig 3.10 shows different wall manufacturing technologies suitable for effusive air cooling system.

The 2D model developed for designing and analyzing the geometrical parameters of effusive cooling systems. It proves that a very effective cooling system performance can be obtained either by varying both the diameter and distribution of the holes or by fixing the hole spacing with minor effects on cooling effectiveness and cooling air consumption. The latter solution copes quite well with a simplification in the design and fabrication matrix for the electroforming process. The analysis showed that for both cases the blade can withstand temperatures up to 1200 K using the effusive cooling system described herein, and that the coolant mass flow for the manufacturing-optimized case resulted to be about unchanged compared to the thermo-fluid-dynamic optimized one

SUMMARY

To satisfy the increasingly high inlet temperature, turbine blade cooling becomes an important aspect. With modern gas turbines operating at extremely high temperatures, it is necessary to implement various cooling methods, so that turbine blades survive in the path of the hot gases. Simply passing coolant air through the airfoils does not provide adequate cooling; therefore, it is necessary to implement techniques that will further enhance the heat transfer from the airfoil walls. These cooling techniques include blade end-wall cooling, leading-edge cooling, trailing-edge cooling, and tip cooling. The blade tip is one of the critical regions to be cooled due to the high thermal load over the tip surface. Therefore, highly accurate and highly detailed local heat transfer and flow data related to such regions are needed for analysis, and cooling schemes must be designed to prevent the failure due to the local hot spots. For external cooling, a common technique is to add film cooling through the tip and near-tip region. The cooling performance is affected significantly by most conditions, such as film cooling hole configuration, location, distribution, and the representative flow conditions. For internal cooling, serpentine cooling passages are designed inside blades, so that the heat from the pressure side and suction side is picked up by the turning coolant extracted from compressors. The serpentine channel configuration, aspect ratio and orientation, rib configuration and location, and rotation and bend/turn geometry affect significantly on the internal cooling efficiency. The increases in allowable blade metal temperature and film cooling effectiveness are more beneficial than improvements in other efficiency parameters such as outlet temperature and compressor pressure ratio.

More studies are needed for the blade-shaped coolant passages with high performance turbulators and with or without film cooling holes under realistic coolant flow, thermal, and rotating conditions. Also, more studies are needed for rotating impingement cooling with or without film coolant extraction as well as rotating pin-fin cooling with or without trailing edge ejection in order to guide the efficient rotor blade internal cooling designs. Highly accurate and highly detailed local heat transfer data is needed to aid engineers in their design of blades for advanced gas turbines.