This paper describes the aero-thermal development and validation of the GT36 heavy duty gas turbine. The turbine which has evolved from the existing and proven GT26 design, consists of an optimised annulus flow path, higher lift aerofoil profiles, optimised aerodynamic matching between the turbine stages and new and improved cooling systems of the turbine vanes and blades. A major design feature of the turbine has been to control and reduce the aerodynamic losses, associated with the aerofoil profiles, trailing edges, blade tips, endwalls and coolant ejection. The advantages of these design changes to the overall gas turbine efficiency have been verified via extensive experimental testing in high-speed cascade test rigs and via the utilisation of high fidelity multi-row computational fluid dynamics design systems. The thermal design and cooling systems of the turbine vanes, blades have also been improved and optimised. For the first stage vane and blade aerofoils and platforms, multi-row film cooling with new and optimised diffuser cooling holes have been implemented and validated in high speed linear cascades. Additionally, the internal cooling design features of all the blades and vanes were also improved and optimised, which allowed for more homogenous metal temperatures distributions on the aerofoils. The verification and validation of the internal thermal designs of all the turbine components has been confirmed via extensive testing in dedicated Perspex models, where measurements were conducted for local pressure losses, overall flow distributions and local heat transfer coefficients. The turbine is currently being tested and undergoing validation in the GT36 Test Power Plant in Birr, Switzerland. The gas turbine is heavily instrumented with a wide range of validation instrumentation including thermocouples, pressure sensors, strain gauges and five-hole probes. In addition to performance mapping and operational validation, a dedicated thermal paint validation test will also be performed.

This paper investigates the aerodynamic and film cooling effectiveness characteristics of a first stage turbine high lift guide vane and its corresponding downstream blade. The vane and blade geometrical profiles and operating conditions are representative of that normally found in a heavy-duty gas turbine. Both the vane and the blade airfoils consist of multi-row film cooling holes located at various axial positions along the airfoil chord. The film cooling holes are geometrically three-dimensional in shape and depending on the location on the airfoil; they can be either symmetrically fan shaped or non-symmetrically fan shaped. Additionally the film cooling holes can be either compounded or in-line with the external flow direction.Numerical studies and experimental investigations in a linear cascade have been conducted at vane and blade exit isentropic Mach number of 0.8. The influence of the coolant flow ejected from the film cooling holes has been investigated for both the vane and the blade profiles.For the nozzle guide vane, the measured film cooling effectiveness compared well with the predictions, especially on the pressure side. The suction side film cooling effectiveness, which consisted of two pre-throat film rows, proved very effective up-to the suction side trailing edge. For the blade, there was a reasonable comparison between the measured and predicted film cooling effectiveness. Again the blade pre-throat fan shaped cooling holes proved very effective up-to the suction side trailing edge. For the vane, the impact of varying the blowing ratios showed a strong variation in the film cooling effectiveness on the pressure side. However, on the blade, the effect of varying the blowing ratio had a greater impact on the suction side film effectiveness compared to the pressure side.

This paper describes the aero-thermal design and validation of an advanced axial flow turbine. This turbine, which has evolved from the existing and proven GT26/GT24 design consists of an optimised annulus flow path using high lift airfoil profiles and improved aerodynamic matching between the turbine stages. A major design feature of the turbine has been to control and reduce the aerodynamic losses, with particular attention being devoted to minimising the secondary, trailing edge and blade tip losses. The advantages of these design changes to the overall turbine efficiency has been verified by extensive controlled experimentation in high-speed cascade test facilities; by the utilisation of 3D multi-row computational fluid dynamics analysis tools, and via engine tests.In addition to the aerodynamic design modifications of the turbine, the thermal designs of the turbine vanes, blades and heat-shields were also optimised. For the first stage film cooled vane and blade airfoils and platforms, both the film cooling layout and operating characteristics were improved. And for all the internally cooled airfoils, the internal heat transfer design features were additionally optimised, which allowed for more homogenous metal temperature distributions on the airfoil and endwall surfaces. The verification and validation of the thermal designs of the turbine components was confirmed via extensive dedicated testing in high-speed cascades for the film cooling performances, and in scaled perspex models for the internal heat transfer coefficients and local flow distributions.The complete turbine was further tested and validated in the GT26 Test Power Plant in Birr, Switzerland via a dedicated turbine thermal paint test run and a subsequent performance and mapping testing phase.

This paper describes a non-invasive, non-destructive, transient inverse measurement technique that allows one to determine internal heat transfer coefficients and rib positions of real gas turbine blades from outer surface temperature measurements after a sudden flow heating. The determination of internal heat transfer coefficients is important during the design process to adjust local heat transfer to spatial thermal load. The detection of rib positions is important during production to fulfill design and quality requirements. For the analysis the one-dimensional transient heat transfer problem inside of the turbine blade's wall was solved. This solution was combined with the Levenberg-Marquardt method to estimate the unknown boundary condition by an inverse technique. The method was tested with artificial data to determine uncertainties with positive results. Then experimental testing with a reference model was carried out. Based on the results, it is concluded that the presented inverse technique could be used to determine internal heat transfer coefficients and to detect rib positions of real turbine blades.

A successful implementation of a cooled turbine blade design for a heavy duty gas turbine engine is a technology challenge that requires a stringent engineering approach. The increased spread of hot gas versus metal temperature, the flatter temperature profiles for reduced emissions and the aerodynamic 3D-profile shape requirement and all at a reduced cooling air consumption place the specification of a new turbine blade, that is put forward to the aerothermal engineers, as a technical challenge. It is also desired to reduce the available development time to be able to introduce new technology features faster into the market. The paper aims to demonstrate turbine blade cooling and heat transfer design process enhancements that allow: increased thermal predictability, increased capturing of three dimensional effects and reduced engineering development cycle time from initial design to full engine validation. Selected items will be shown for demonstration. One example is the use of symmetry and parameterization to move CFD from an analysis tool into an available design tool to capture effects as rotation or three-dimensionality. Another example is the use of heat sinks within a finite element model to represent individual cooling holes instead of hole geometry.

One of the most challenging aspects of gas turbine cooling is the cooling of the first stages of turbine blades. Here the highest external heat load is seen at the leading edge of the blade. The present study investigates the internal cooling in a triangular channel with a rounded edge as a model of a leading edge cooling channel for a gas turbine blade. A transient liquid crystal method is used to measure the heat transfer. Experimental results are reported for a number of new 3D rib configurations for Reynolds numbers between 50 000 and 200 000. From the experimental results it has been found that 60 deg. ribs provide in general higher heat transfer enhancements than 45 deg. ribs. However, this results in extremely high friction factors for the 60 deg. ribs. Taking the local and mean distributions of the heat transfer coefficients (as well as the increase in friction factors) into consideration, it was found that the most promising rib arrangement for leading edge cooling is a 3D rib with 45 deg. angle and double-sided fully overlapped ribs in the arc area. These ribs provide uniform heat transfer in the arc area as well as a high level of the heat transfer coefficients in the channel. The resulting friction factors are in an acceptable range for these ribs. (c) 2006 Elsevier Masson SAS. All rights reserved.

The reduction of 2-bromo-1,3,2-diazaborole tBuNCH=CH(tBu)BBr (4c) with a potassium mirror in 1,2-dimethoxyethane afforded a non-separable 1 : 2 : 1 mixture of the compounds tBuNCH=CHN(tBu)BH (5), tBuNCH=CHN(tBu)BOCH3 (6) and {tBuNCH=CHN(tBu)B}(2)O (7). Reaction of 4c with a potassium-sodium alloy in N,N,N',N'-tetramethylethylenediamine led to 5 as the major product. 1,3,2-Diazaboroles tBuNCH=CHN(tBu)BN-Me-2 (8) and {tBuNCH=CHN(tBu)B}(2) (9) were spectroscopically identified as minor products. The treatment of 4c with potassium-sodium alloy in toluene solution in the presence of [15]crown-5 yielded a 1 : 1 mixture of 5 and the benzyl derivative tBuNCH=CHN( tBu)BCH2Ph (12). The same reaction in toluene-d(8) produced the deuterated species 5-d(1) and 12-d(7). tBuNCH=CHN(tBu)BCH3 (17) and 1,4-diazabutadiene (tBuN=CH)(2) (18) resulted from the treatment of tBuNCH=CHN(tBu)BSCH3 (15) with potassium-sodium alloy in n-hexane. In contrast to this, compound 9 was obtained as the main product of the reduction of tBuNCH=CHN(tBu)BStBu (16) under similar conditions. The reduction of the 1-bromo-2-tert-butyl-1,2-dihydro[1,3,2]diazaborolo[1,5-a]pyridine (19) smoothly produced the respective diborane(4) derivative (20) which was subjected to X-ray diffraction analysis.

Reaction of a pyridine carbaldimine 2-(BuN)-Bu-t=CHC5H4N 1a with a molar equivalent of boron tribromide afforded a bicyclic 1,3,2-diazaborolium bromide 2a as an orange solid, whereas 1a and the corresponding 2-(2,6-Me2C6H3)N=CHC5H4N 1b with two equivalents of boron trifluoride gave non-ionic yellow adducts 3a and 3b. The reduction of compounds 2a, 3a and 3b with an excess of sodium amalgam in a hexane slurry led to the formation of 1-X-2-R-1,2-dihydro[1,3,2]diazaborolo[1,5-a]pyridines 4 (X=Br; R=Bu-t), 5a (F; Bu-t) and 5b (F; 2,6-Me2C6H3) as yellow oils (4a, 5a) or a yellow wax (5b), respectively. Treatment of 4 with an excess of chlorotrimethylsilane caused a Br/Cl exchange to afford chloro derivative 6. The addition of a methyl group to the boron atom was effected by reaction of heterocycle 4 with methyllithium. The BCN derivative 8 resulted from treatment of 4 with silver cyanide. Reduction of 4 with lithium aluminium hydride gave the 1-hydro-derivative 9, whereas the (BSBu)-Bu-t compound was obtained from the reaction of 4 with (KSBu)-Bu-t. Compound 8 was subjected to an X-ray diffraction analysis.

Reaction of equimolar amounts of diphenylketene with a series of 1,3-di-tert-butyl-2,3-dihydro-1H-1,3,2-diazaboroles [X = Br (1a), F (1b), NH2 (1c), NMe2 (1d), Me (1e), SnMe3 (1f), CHC(SnMe3)C6H4-4-Cl (1g)] regioselectively afforded good yields of the 1,3,2-oxazaborolidines (2a−g). The X-ray structure analysis of 2d revealed an essentially planar five-membered heterocycle with a long B−O bond and a strong exocyclic BN−π bond.

3,4-Dihydro-2,4-diphenyl-2H-1,2,4,3-triazaboroles 3a, 3b, and 4 were synthesized by cyclocondensation of N1,N3-diphenylformamidrazone (1) with dibromophenylborane, dibromomethylborane, and boron trichloride. 3-Chloro-3,4-dihydro-2,4-diphenyl-2H-1,2,4,3-triazaborole (4) was converted into 3,4-dihydro-2,4-diphenyl-2H-1,2,4,3-triazaborole (5) by treatment with LiAlH4. The corresponding 3-cyanato and 3-cyano derivatives 6 and 7 resulted from the reaction of 4 with AgOCN and AgCN, respectively. Compound 7 was transformed into the bis(1,2,4,3-triazaborolyl)oxane 8 by silver oxide. Compounds 1–8 were characterized by elemental analyses and spectroscopic methods (1H, 11B, and 13C NMR; IR; MS). The molecular structure of 8 was established by single-crystal X-ray diffraction analysis.

A detailed study of the electronic structure of Nb3Te43As, for x = 0, 0.5 and 1.0 using core level X-ray photoelectron spectroscopy and valence band ultraviolet photoelectron spectroscopy is presented. Using He I and He II radiation, pronounced changes in the spectra upon replacing Te by As are found in the region close to the Fermi energy. These changes are explained on the basis of the electronic band structure calculations using the linear muffin tin orbital method. In addition, these calculations clearly demonstrate the change from quasi-one- to three-dimensional behavior upon going from Nb3Te4 to Nb3Te3As. This is in line with results of resistivity measurements presented, as well as previous X-ray single crystal investigations. Down to 1.5 K no transition into the superconducting state is observed.