Ronald Scott Bunker

A Review of Shaped Hole Turbine Film-Cooling Technology

Film cooling represents one of the few game-changing technologies that has allowed the achievement of todays high firing temperature, high-efficiency gas turbine engines. Over the last 30 years, only one major advancement has been realized in this technology, that being the incorporation of exit shaping to the film holes to result in lower momentum coolant injection jets with greater surface coverage. This review examines the origins of shaped film cooling and summarizes the extant literature knowledge concerning the performance of such film holes. A catalog of the current literature data is presented, showing the basic shaping geometries, parameter ranges, and types of data obtained. Specific discussions are provided for the flow field and aerodynamic losses of shaped film hole coolant injection. The major fundamental effects due to coolant-to-gas blowing ratio, compound angle injection, cooling hole entry flow character, and mainstream turbulence intensity are each reviewed with respect to the resulting adiabatic film effectiveness and heat transfer coefficients for shaped holes. A specific example of shaped film effectiveness is provided for a production turbine inlet vane with comparison to other data. Several recent unconventional forms of film hole shaping are also presented as a look to future potential improvements

A Review of Turbine Blade Tip Heat Transfer

Abstract: This paper presents a review of the publicly available knowledge base concerning turbine blade tip heat transfer, from the early fundamental research which laid the foundations of our knowledge, to current experimental and numerical studies utilizing engine‐scaled blade cascades and turbine rigs. Focus is placed on high‐pressure, high‐temperature axial‐turbine blade tips, which are prevalent in the majority of todays aircraft engines and power generating turbines. The state of our current understanding of turbine blade tip heat transfer is in the transitional phase between fundamentals supported by engine‐based experience, and the ability to a priori correctly predict and efficiently design blade tips for engine service.

Effect of Trench Width and Depth on Film Cooling From Cylindrical Holes Embedded in Trenches

The present study is an experimental investigation of film cooling from cylindrical holes embedded in transverse trenches. Different trench depths are considered with two trench widths. Trench holes can occur when blades are coated with thermal barrier coating (TBC) layers. The film-hole performance and behavior will be different for the trench holes compared to standard cylindrical holes that are flush with the surface. The trench width and depth depend on the mask region and the thickness of the TBC layer. Detailed heat transfer coefficient and film effectiveness measurements are obtained simultaneously using a single test transient IR thermography technique. The study is performed at a single mainstream Reynolds number based on freestream velocity and film-hole diameter of 11,000 at four different coolant-to-mainstream blowing ratios of 0.5, 1.0, 1.5, and 2.0. The results show that film effectiveness is greatly enhanced by the trenching due to the improved two-dimensional nature of the film and lateral spreading. The detailed heat transfer coefficient and film effectiveness contours provide a clear understanding of the jet-mainstream interactions for different hole orientations. Computational fluid dynamics simulation using FLUENT was also performed to determine the jet-mainstream interactions to better understand the surface heat transfer coefficient and film effectiveness distributions.

Heat Transfer Coefficients on the Squealer Tip and Near-Tip Regions of a Gas Turbine Blade With Single or Double Squealer

Detailed heat transfer coefficient distributions on a gas turbine squealer tip blade were measured using a hue detection based transient liquid-crystals technique. The heat transfer coefficients on the shroud and near tip regions of the pressure and suction sides of a blade were also measured. Squealer rims were located along (a) the camber line, (b) the pressure side, (c) the suction side, (d) the pressure and suction sides, (e) the camber line and the pressure side, and (f) the camber line and the suction side, respectively. Tests were performed on a five-bladed linear cascade with a blow down facility. The Reynolds number based on the cascade exit velocity and the axial chord length of a blade was 1.1 ×10 6 and the overall pressure ratio was 1.2. Heat transfer measurements were taken at the three tip gap clearances of 1.0%, 1.5%, and 2.5% of blade span. Results show that the heat transfer coefficients on the blade tip and the shroud were significantly reduced by using a squealer tip blade. Results also showed that a different squealer geometry arrangement changed the leakage flow path and resulted in different heat transfer coefficient distributions. The suction side squealer tip provided the lowest heat transfer coefficient on the blade tip and near tip regions compared to the other squealer geometry arrangements.

Gas Turbine Heat Transfer: Ten Remaining Hot Gas Path Challenges

The advancement of turbine cooling has allowed engine design to exceed normal material temperature limits, but it has introduced complexities that have accentuated the thermal issues greatly. Cooled component design has consistently trended in the direction of higher heat loads, higher through-wall thermal gradients, and higher in-plane thermal gradients. The present discussion seeks to identify ten major thermal issues, or opportunities, that remain for the turbine hot gas path (HGP) today. These thermal challenges are commonly known in their broadest forms, but some tend to be little discussed in a direct manner relevant to gas turbines. These include uniformity of internal cooling, ultimate film cooling, microcooling, reduced incident heat flux, secondary flows as prime cooling, contoured gas paths, thermal stress reduction, controlled cooling, low emission combustor-turbine systems, and regenerative cooling. Evolutionary or revolutionary advancements concerning these issues will ultimately be required in realizable engineering forms for gas turbines to breakthrough to new levels of performance. Herein lies the challenge to researchers and designers. It is the intention of this summary to provide a concise review of these issues, and some of the recent solution directions, as an initial guide and stimulation to further research.

Effect of Squealer Geometry Arrangement on a Gas Turbine Blade Tip Heat Transfer

This study investigates the effect of a squealer tip geometry arrangement on heat transfer coefficient and static pressure distributions on a gas turbine blade tip in a five-bladed stationary linear cascade. A transient liquid crystal technique is used to obtain detailed heat transfer coefficient distribution. The test blade is a linear model of a tip section of the GE E 3 high-pressure turbine first stage rotor blade. Six tip geometry cases are studied: (1) squealer on pressure side, (2) squealer on mid camber line, (3) squealer on suction side, (4) squealer on pressure and suction sides, (5) squealer on pressure side plus mid camber line, and (6) squealer on suction side plus mid camber line. The flow condition during the blowdown tests corresponds to an overall pressure ratio of 1.32 and exit Reynolds number based on axial chord of 1.1×10 6

Effect of Tip Gap and Squealer Geometry on Detailed Heat Transfer Measurements Over a High Pressure Turbine Rotor Blade Tip

The present study explores the effects of gap height and tip geometry on heat transfer distribution over the tip surface of a HPT first-stage rotor blade. The pressure ratio (inlet total pressure to exit static pressure for the cascade) used was 1.2, and the experiments were run in a blow-down test rig with a four-blade linear cascade. A transient liquid crystal technique was used to obtain the tip heat transfer distributions. Pressure measurements were made on the blade surface and on the shroud for different tip geometries and tip gaps to characterize the leakage flow and understand the heat transfer distributions. Two different tip gap-to-blade span ratios of 1% and 2.6% are investigated for a plane tip, and a deep squealer with depth-to-blade span ratio of 0.0416. For a shallow squealer with depth-to-blade span ratio of 0.0104, only 1% gap-to-span ratio is considered. The presence of the squealer alters the tip gap flow field significantly and produces lower overall heat transfer coefficients. The effects of different partial squealer arrangements are also investigated for the shallow squealer depth. These simulate partial burning off of the squealer in real turbine blades. Results show that some partial burning of squealers may be beneficial in terms of overall reduction in heat transfer coefficients over the tip surface.

Film Cooling Effectiveness Due to Discrete Holes Within a Transverse Surface Slot

The goal of many turbine airfoil film cooling schemes is the achievement of a tangentially injected 2D layer of protective film over the surface. In common nomenclature, this is referred to as 2D slot film cooling, which can achieve adiabatic effectiveness levels approaching unity at the injection location. Since continuous and uninterrupted slots are not structurally feasible in the high pressure turbine components, other approximate film cooling geometries have been sought. The present study examines two film cooling geometries which are formed by the combination of internal discrete film holes feeding continuous 2D surface slots. Experiments have been performed within a flat plate wind tunnel test section, which includes an accelerating freestream condition to model the surface of a turbine airfoil. As suggested by the experiments of Wang et al. [1], a normal 2D surface slot is located transverse to the mainstream flow direction. The slot is fed by a row of discrete coolant supply holes oriented in the spanwise direction with inclination angle of 30-degrees, pitch-to-diameter ratio of 3.57, and length-to-diameter ratio of 5.7. The slot depth-to-hole diameter ratio is S/D of 3. Two such slots were tested, one with axial width-to-hole diameter ratio of 1.13, and the other with ratio of 1.5. Tests were conducted for supply hole blowing ratios of 0.75 to 4, density ratios of 1.8, and a freestream approach turbulence intensity of 4.5%. The holes-within-slot film effectiveness data are compared with both axial and radial film data, ie. S/D equal to zero, obtained in the same test section. The holes-in-slot geometries demonstrate two important characteristics, (1) a relative insensitivity of the adiabatic film effectiveness to blowing rate, and (2) no observed film blow-off at high blowing rates. In addition, a novel film cooling arrangement is demonstrated in which the surface slot is very shallow, forming a narrow trench with S/D of only 0.43. It is shown that this novel surface geometry yields the best film effectiveness of all cases tested.Copyright

Jet mixing in a slot

Abstract Jet mixing inside a slot has the potential to be applied for cooling in various components in gas turbines. An experimental program was conducted to focus on investigating the flow mixing behavior inside the slots. Various parameters including orientation angle, inclination angle, slot width, effect of primary flow and slot depth were systematically examined. An array of seven jets with a jet-to-jet spacing of five diameters and velocity ratio of unity was used in the experiment. The results indicated that the flow distribution at the slot exit becomes more uniform as the orientation angle was increased from 0° to 60°. Wider slot width brought about high non-uniformity and was clearly undesirable. The optimum slot depth was found to range from 2 to 2.8 times the jet diameter. The flow field was the most uniform where the shear layers of the two adjacent jets met. Regions of low velocity and high pressure caused non-uniform velocity distribution downstream where the jets met, so deeper slots do not necessarily render better uniformity. When the orientation angle increased from 0 to 30° and 60°, the jets bent toward the wall in the inclined direction caused by the Coanda effect. The presence of the primary flow forced the jet to spread faster in the lateral direction and caused some non-uniform flow pockets at the slot exit. The compound angle configuration (60° jet orientation and 30° slot inclination angles) was discovered as the best choice.

Heat Transfer and Friction Factors for Flows Inside Circular Tubes With Concavity Surfaces

Heat transfer and friction coefficients measurements have been obtained for fully developed, turbulent internal flows in circular tubes with six different concavity (dimple) surface array geometries. Two different concavity depths and three different concavity array densities were tested using tube bulk flow Reynolds numbers from 20,000 to 90,000. Liquid-crystal thermography was used to measure the temperature distributions on the outside of the concavity tubes. Using the average heat transfer coefficient for the fully developed region, the overall heat transfer enhancements are compared to baseline smooth tube results. Friction coefficients are also compared to values for a smooth tube. Dimple depths of 0.2-0.4 relative to the dimple surface diameter were used, with surface area densities ranging from 0.3 to 0.7. Dimple arrays were all in-line geometries. The results showed that heat transfer enhancements for dimpled internal surfaces of circular passages can reach factors of 2 or more when the relative dimple depth is greater than 0.3 and the dimple array density is about 0.5 or higher. The associated friction factor multipliers for such configurations are in the range of 4-6. The present study provides a first insight into the heat transfer and friction effects of various concavity arrays for turbulent flows.