A combined experimental and computational study has been performed to investigate the detailed distribution of convective heat transfer coefficients on the first stage blade tip surface for a geometry typical of large power generation turbines (>100MW). This paper is concerned with the design and execution of the experimental portion of the study, which represents the first reported investigation to obtain nearly full surface information on heat transfer coefficients within an environment which develops an appropriate pressure distribution about an airfoil blade tip and shroud model. A stationary blade cascade experiment has been run consisting of three airfoils, the center airfoil having a variable tip gap clearance. The airfoil models the aerodynamic tip section of a high pressure turbine blade with inlet Mach number of 0.30, exit Mach number of 0.75, pressure ratio of 1.45, exit Reynolds number based on axial chord of 2.57.10 6, and total turning of about 110 degrees. A hue detection based liquid crystal method is used to obtain the detailed heat transfer coefficient distribution on the blade tip surface for flat, smooth tip surfaces with both sharp and rounded edges. The cascade inlet turbulence intensity level took on values of either 5% or 9%. The cascade also models the casing recess in the shroud surface ahead of the blade. Experimental results are shown for the pressure distribution measurements on the airfoil near the tip gap, on the blade tip surface, and on the opposite shroud surface. Tip surface heat transfer coefficient distributions are shown for sharp-edge and rounded-edge tip geometries at each of the inlet turbulence intensity levels.

Full-surface heat transfer coefficient distribution measurements have been made using a liquid crystal thermography technique for several cases of normally impinging jet arrays onto a flat, smooth surface within a region bounded on three sides. While the impingement target plate remains of a fixed size, the impingement jet array has been changed to cover a wide range of conditions, extending beyond the currently available literature data. Axial and lateral jet spacing values of x/D and y/D of 3, 6, and 9 have been used, all with square orientation and in-line jets. The jet plate-to-target surface distance z/D has been varied from 1.25 to 5.5. Jet Reynolds numbers ranged from 14,000 to 65,000. In the sparse array limiting case, the number of jet rows is four in the axial direction and three in the lateral direction. For the dense array limiting case, the number of jet rows is 26 in the axial direction and 20 in the lateral direction. Using both heat transfer and pressure distribution measurements, results are compared to the existing correlation of Florschuetz et al. [1], showing excellent agreement in regions of common parameters. In regions not previously reported in the literature, the present study extends the streamwsie row-averaged heat transfer coefficient correlation of [1] with a modified correlation for design use.

An introduction is given to a new rotating wheelspace test vehicle known as the GE Hot Gas Ingestion Rig (HGIR). This scaled 1.5 stage turbine rig is configured similar to a current generation heavy duty gas turbine. It has a broad spectrum of measurement capability, including radial and circumferential ports for CO2 measurements that are used to measure the sealing effectiveness from candidate rim seal geometries. Engine-matched conditions are presented in a non-dimensional form that demonstrate the value of this fully capable test facility, including static pressure signatures at stage 1 nozzle exit, exit Reynolds number, exit Mach number and rotational Reynolds number. This paper also provides details of the operating conditions and assessment of a thermal steady-state condition achieved consistently throughout each test. Part I of this two-part paper focuses on the geometric details of this new state-of-the-art wheelspace rig, the measurement capabilities currently available and planned, and the results from the baseline geometry. The test data from this test vehicle are used to validate reduced order models, including unsteady CFD models. Details of the CFD modeling and validation are presented in the Part II paper Ding et al. [1]. Measurement uncertainties for all key parameters as well as the repeatability of the test rig to reproduce test conditions are presented to demonstrate the rigor taken in the design and operation of this testing facility.

A combined experimental and computational study has been performed to investigate the detailed distribution of convective heat transfer coefficients on the first-stage blade tip surface for a geometry typical of large power generation turbines (>100 MW). This paper is concerned with the design and execution of the experimental portion of the study, which represents the first reported investigation to obtain nearly full surface information on heat transfer coefficients within an environment that develops an appropriate pressure distribution about an airfoil blade tip and shroud model. A stationary blade cascade experiment has been run consisting of three airfoils, the center airfoil having a variable tip gap clearance. The airfoil models the aerodynamic tip section of a high-pressure turbine blade with inlet Mach number of 0.30, exit Mach number of 0.75, pressure ratio of 1.45, exit Reynolds number based on axial chord of 2.57×106, and total turning of about 110 deg. A hue detection based liquid crystal method is used to obtain the detailed heat transfer coefficient distribution on the blade tip surface for flat, smooth tip surfaces with both sharp and rounded edges. The cascade inlet turbulence intensity level took on values of either 5 or 9 percent. The cascade also models the casing recess in the shroud surface ahead of the blade. Experimental results are shown for the pressure distribution measurements on the airfoil near the tip gap, on the blade tip surface, and on the opposite shroud surface. Tip surface heat transfer coefficient distributions are shown for sharp edge and rounded edge tip geometries at each of the inlet turbulence intensity levels. [S0889-504X(00)01902-4]

Experiments and numerical simulations were conducted to understand the heat transfer characteristics of a stationary gas turbine combustor liner cooled by impingement jets and cross flow between the liner and sleeve. Heat transfer was also aided by trip-strip turbulators on the outside of the liner and in the flowsleeve downstream of the jets. The study was aimed at enhancing heat transfer and prolonging the life of the combustor liner components. The combustor liner and flow sleeve were simulated using a flat plate rig. The geometry has been scaled from actual combustion geometry except for the curvature. The jet Reynolds number and the mass-velocity ratios between the jet and cross flow in the rig were matched with the corresponding combustor conditions. A steady state liquid crystal technique was used to measure spatially resolved heat transfer coefficients for the geometric and flow conditions mentioned above. The heat transfer was measured both in the impingement region as well as over the turbulators. A numerical model of the combustor test rig was created that included the impingement holes and the turbulators. Using CFD, the flow distribution within the flow sleeve and the heat transfer coefficients on the liner were both predicted. Calculations were made by varying the turbulence models, numerical schemes, and the geometrical mesh. The results obtained were compared to the experimental data and recommendations have been made with regard to the best modeling approach for such liner-flow sleeve configurations.