Philip Poinsatte

Experimental Heat Transfer and Bulk Air Temperature Measurements for a Multipass Internal Cooling Model with Ribs and Bleed

An experimental study was made to obtain heat transfer and air temperature data for a simple three-leg serpentine test section that simulates a turbine blade internal cooling passage with trip strips and bleed holes. The objectives were to investigate the interaction of ribs and various bleed conditions on internal cooling and to gain a better understanding of bulk air temperature in an internal passage. Steady-state heat transfer measurements were obtained using a transient technique with thermochromic liquid crystals. Trip strips were attached to one wall of the test section and were located either between or near the bleed holes. The bleed holes, used for film cooling, were metered to simulate the effect of external pressure on the turbine blade. Heat transfer enhancement was found to be greater for ribs near bleed holes compared to ribs between holes, and both configurations were affected slightly by bleed rates upstream. Air temperature measurements were taken at discrete locations along one leg of the model. Average bulk air temperatures were found to remain fairly constant along one leg of the model.

Convective heat transfer measurements from a NACA 0012 airfoil in flight and in the NASA Lewis Icing Research Tunnel

Local heat transfer coefficients from a smooth and roughened NACA 0012 airfoil were measured using a steady state heat flux method. Heat transfer measurements on the specially constructed 0.533 meter chord airfoil were made both in flight on the NASA Lewis Twin Otter Research Aircraft and in the NASA Lewis Icing Research Tunnel (IRT). Roughness was obtained by the attachment of small, 2 mm diameter, hemispheres of uniform size to the airfoil surface in four distinct patterns. The flight data was taken for the smooth and roughened airfoil at various Reynolds numbers based on chord in the range of 1.24x10(exp 6) to 2.50x10(exp 6) and at various angles of attack up to 4 degrees. During these flight tests the free stream velocity turbulence intensity was found to be very low (less than 0.1 percent). The wind tunnel data was taken in the Reynolds number range of 1.20x10(exp 6) to 4.52x10(exp 6) and at angles of attack from -4 degrees to +8 degrees. The turbulence intensity in the IRT was 0.5 to 0.7 percent with the cloud making spray off. Results for both the flight and tunnel tests are presented as Frossling number based on chord versus position on the airfoil surface for various roughnesses and angle of attack. A table of power law curve fits of Nusselt number as a function of Reynolds number is also provided. The higher level of turbulence in the IRT versus flight had little effect on heat transfer for the lower Reynolds numbers but caused a moderate increase in heat transfer at the higher Reynolds numbers. Turning on the cloud making spray air in the IRT did not alter the heat transfer. Roughness generally increased the heat transfer by locally disturbing the boundary layer flow. Finally, the present data was not only compared with previous airfoil data where applicable, but also with leading edge cylinder and flat plate heat transfer values which are often used to estimate airfoil heat transfer in computer codes.

Turbulence and Heat Transfer Measurements in an Inclined Large Scale Film Cooling Array: Part II—Temperature and Heat Transfer Measurements

The second of a two-part paper, this study focuses on the temperature field and surface heat transfer measurements on a large-scale models of an inclined row of film cooling holes. Detailed surface and flow field measurements were taken and presented in Part I. The model consists of three holes of 1.9-cm diameter that are spaced 3 hole diameters apart and inclined 30° from the surface. Additionally, another model with an anti-vortex adaptation to the film cooling holes is also tested. The coolant stream is metered and cooled to 20°C below the mainstream temperature. A thermocouple is used to obtain the flow temperatures along the jet centerline and at various streamwise locations. Steady state liquid crystal thermography is used to obtain surface heat transfer coefficients. Results are obtained for blowing ratios of up to 2 in order to capture off-design conditions in which the jet is lifted. Film cooling effectiveness values of 0.4 and 0.15 were found along the centerline for blowing ratios of 1 and 2 respectively. In addition, an anti-vortex design was tested and found to have improved film effectiveness. This paper presents the detailed temperature contours showing the extent of mixing between the coolant and freestream and the local heat transfer results.Copyright

Turbulence and Heat Transfer Measurements in an Inclined Large Scale Film Cooling Array: Part I—Velocity and Turbulence Measurements

A large-scale model of an inclined row of film cooling holes is used to obtain detailed surface and flow field measurements that will enable future computational fluid dynamics code development and validation. The model consists of three holes of 1.9-cm diameter that are spaced 3 hole diameters apart and inclined 30° from the surface. The length to diameter ratio of the coolant holes is about 18. Measurements include film effectiveness using IR thermography and near wall thermocouples, heat transfer using liquid crystal thermography, flow field temperatures using a thermocouple, and velocity and turbulence quantities using hotwire anemometry. Results are obtained for blowing ratios of up to 2 in order to capture severe conditions in which the jet is lifted. This first part of the two-part paper presents the detailed velocity component and turbulence stresses along the centerline of the film-cooling hole and at various streamwise locations.Copyright

Heat Transfer Measurements for a Film Cooled Turbine Vane Cascade

Experimental heat transfer and pressure measurements were obtained on a large scale film cooled turbine vane cascade. The objective was to investigate heat transfer on a commercial high pressure first stage turbine vane at near engine Mach and Reynolds number conditions. Additionally blowing ratios and coolant density were also matched. Numerical computations were made with the Glenn-HT code of the same geometry and compared with the experimental results. A transient thermochromic liquid crystal technique was used to obtain steady state heat transfer data on the midspan geometry of an instrumented vane with 12 rows of circular and shaped film cooling holes. A mixture of SF6 and Argon gases was used for film coolant to match the coolant-to-gas density ratio of a real engine. The exit Mach number and Reynolds number were 0.725 and 2.7 million respectively. Trends from the experimental heat transfer data matched well with the computational prediction, particularly for the film cooled case.

Experimental Heat Transfer and Bulk Air Temperature Measurements for a Multipass Internal Cooling Model With Ribs and Bleed

An experimental study was made to obtain heat transfer and air temperature data for a simple 3-leg serpentine test section that simulates a turbine blade internal cooling passage with trip strips and bleed holes. The objectives were to investigate the interaction of ribs and various bleed conditions on internal cooling and to gain a better understanding of bulk air temperature in an internal passage. Steady state heat transfer measurements were obtained using a transient technique with thermochromic liquid crystals. Trip strips were attached to one wall of the test section and were located either between or near the bleed holes. The bleed holes, used for film cooling, were metered to simulate the effect of external pressure on the turbine blade. Heat transfer enhancement was found to be greater for ribs near bleed holes compared to ribs between holes, and both configurations were affected slightly by bleed rates upstream. Air temperature measurements were taken at discreet locations along one leg of the model. Average bulk air temperatures were found to remain fairly constant along one leg of the model.Copyright

Application of Pinniped Vibrissae to Aeropropulsion

Vibrissae (whiskers) of Phoca Vitulina (Harbor Seal) and Mirounga Angustirostris (Elephant Seal) possess undulations along their length. Harbor Seal Vibrissae have been shown to reduce vortex induced vibrations and reduce drag compared to appropriately scaled cylinders and ellipses. Samples of Harbor Seal vibrissae, Elephant Seal vibrissae and California Sea Lion vibrissae were collected from the Marine Mammal Center in California. CT scanning, microscopy and 3D scanning techniques were utilized to characterize the whiskers. Leading edge parameters from the whiskers were used to create a 3D profile based on a modern power turbine blade. The NASA SW-2 cascade wind tunnel facility was used to perform hotwire surveys and pitot surveys in the wake of the ‘Seal Blades’ to provide validation of Computational Fluid Dynamics simulations. Computational Fluid Dynamics simulations were used to study the effect of incidence angles from −37 to +10 degrees on the aerodynamic performance of the Seal blade. The tests and simulations were conducted at a Reynolds number of 100,000 based on inlet conditions and blade axial chord. The Seal blades showed consistent performance improvements over the baseline configuration. It was determined that a fuel burn reduction of approximately 5% could be achieved for a fixed wing aircraft.Copyright

High Reynolds Number and Turbulence Effects on Turbine Heat Transfer

Experimental data on pressure distribution and heat transfer on a turbine airfoil were obtained over a range of Reynolds numbers from 0.75 to 7.0 x 10 6 and a range of turbulence intensities from 1.8 to about 15%. The purpose of this study was to obtain fundamental heat transfer and pressure distribution data over a wide range of high Reynolds numbers and to extend the heat transfer data base to include the range of Reynolds numbers encountered in the Space Shuttle main engine turbopump turbines. The results of this study indicated that Reynolds number and turbulence intensity have a large effect on both the transition from laminar to turbulent flow and the resulting heat transfer. For a given turbulence intensity, heat transfer for all Reynolds numbers at the leading edge can generally be correlated with the Frossling number developed for lower Reynolds numbers. For a given turbulence intensity, heat transfer for the airfoil surfaces downstream of the leading edge can be approximately correlated with a dimensionless parameter. Comparisons of the experimental results were also made with a numerical solution from a two-dimensional Navier-Stokes code. A B C d E MO) h

Numerical and Experimental Examination of Turbulent Mixing of a Heated Jet in Crossflow [STUB]

The injection of fully-developed turbulent heated air from a tube into a cooler turbulent duct flow is examined, as an analogy to film cooled turbine blades. A LES numerical model is developed and applied in which tube and duct turbulence inflow effects are emulated using a divergence-free synthetic eddy method (SEM). For direct comparison, a hot-wire experiment is conducted within the ERB test cell SW-6 at NASA Glenn Research Center. Results related to velocity, temperature, and heat flux are obtained numerically and experimentally for a blowing ratio of 1.2, involving a 36 K temperature difference. The relative effect on the solutions of tube and duct inflow turbulence is systematically evaluated. The impact of inherent low-pass filtering of temperature measurements and probe wire offset on the experimental results are addressed.

Investigation of Spiral and Sweeping Holes

Surface infrared thermography, hotwire anemometry, and thermocouple surveys were performed on two new film cooling hole geometries: spiral/rifled holes and fluidic sweeping holes. The spiral holes attempt to induce large-scale vorticity to the film cooling jet as it exits the hole to prevent the formation of the kidney shaped vortices commonly associated with film cooling jets. The fluidic sweeping hole uses a passive in-hole geometry to induce jet sweeping at frequencies that scale with blowing ratios. The spiral hole performance is compared to that of round holes with and without compound angles. The fluidic hole is of the diffusion class of holes and is therefore compared to a 777 hole and Square holes. A patent-pending spiral hole design showed the highest potential of the non-diffusion type hole configurations. Velocity contours and flow temperature were acquired at discreet cross-sections of the downstream flow field. The passive fluidic sweeping hole shows the most uniform cooling distribution but suffers from low span-averaged effectiveness levels due to enhanced mixing. The data was taken at a Reynolds number of 11,000 based on hole diameter and freestream velocity. Infrared thermography was taken for blowing ratios of 1.0, 1.5, 2.0, and 2.5 at a density ratio of 1.05. The flow inside the fluidic sweeping hole was studied using 3D unsteady RANS.Copyright