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This paper describes the effects of coolant-to-mainstream density ratio and specific heat capacity flux ratio (the product of blowing ratio and specific heat capacity ratio) on the overall cooling effectiveness of high pressure (HP) turbine vanes. Experimental measurements have been conducted at correct engine-matched conditions of Mach number, Reynolds number, turbulence intensity, and coolant-to-mainstream momentum flux ratio. Vanes tested were fully cooled production parts from an engine currently in service. A foreign gas mixture of SF6 and Ar was selected for injection as coolant in the facility so that density and blowing ratios were also matched to the engine situation. The isentropic exponent of the foreign gas mixture coincides with that of air. Full-coverage surface maps of overall cooling effectiveness were acquired by an infrared (IR) thermography technique at a range of mainstream-to-coolant temperature ratios. Measurements were subsequently scaled to engine conditions by employing a new theory based on the principle of superposition and a recovery and redistribution temperature demonstrated in previous papers. It is shown that the two aerodynamically matched situations of air- and foreign-gas-cooled experiments give virtually the same effectiveness trends and patterns. Actual levels differ, however, on account of specific heat capacity flux ratio differences. The effect is described and quantified by a one-dimensional analytical model of the vane wall. Differences in Biot number with respect to engine conditions are discussed as they also influence the scaling of turbine metal temperatures. [ABSTRACT FROM AUTHOR]

Experimental measurements of overall cooling effectiveness conducted on a high-pressure turbine vane in a warm rig flow are scaled to engine conditions in this paper. A new theory for the scaling of turbine metal temperatures in cooled compressible flows has been applied, based on the principle of superposition, and demonstrated analytically and numerically in a previous paper. The analysis employs a definition of overall cooling effectiveness based on a new recovery and redistribution temperature, which makes it independent of the temperature boundary conditions of the hot and cold flow streams. This enables the vane external wall temperatures to be scaled to engine conditions by varying, in a fixed aerodynamic field, the mainstream-to-coolant temperature ratio. Experimental validation of the theory is provided in this article. Measurements were conducted in the Annular Sector Heat Transfer Facility, which employs fully cooled nozzle guide vanes, production parts of a civil aviation engine currently in service. Mainstream Mach and Reynolds numbers, inlet turbulence intensity, and coolant-to-mainstream total pressure ratio (and thus momentum flux ratio) are all matched to engine conditions. Full-coverage overall cooling effectiveness distributions, acquired by infrared thermography, are presented for a range of mainstream-to-coolant temperature ratios between 1.05 and 1.22 and subsequently scaled to engine conditions by an iterative procedure. In reducing to practice the principles of the new scaling theory, it is demonstrated that direct validation of turbine cooling system performance is possible in experiments at lower than engine temperatures. [ABSTRACT FROM AUTHOR]

A new experimental technique for the accurate measurement of steady-state metal temperature surface distributions of modern heavily film-cooled turbine vanes has been developed and is described in this paper. The technique is analogous to the thermal paint test but has been designed for fundamental research. The experimental facility consists of an annular sector cascade of high pressure (HP) turbine vanes from a current production engine. Flow conditioning is achieved by using an annular sector of deswirl vanes downstream of the test section, being both connected by a three-dimensionally contoured duct. As a result, a transonic and periodic flow, highly representative of the engine aerodynamic field, is established: Inlet turbulence levels, mainstream Mach and Reynolds numbers, and coolant-to-mainstream total pressure ratio are matched. Since the fully three-dimensional nozzle guide vane (NGV) geometry is used, the correct radial pressure gradient and secondary flow development are simulated and the cooling flow redistribution is engine-realistic. To allow heat transfer measurements to be performed, a mainstream-to-coolant temperature difference (up to 33.5°C) is generated by using two steel-wire mesh heaters, operated in series. NGV surface metal temperatures are measured (between 20°C and 40°C) by wide-band thermochromic liquid crystals. These are calibrated in situ and on a per-pixel basis against vane surface thermocouples, in a heating process that spans the entire color play and during which the turbine vanes can be assumed to slowly follow a succession of isothermal states. Experimental surface distributions of overall cooling effectiveness are presented in this paper. By employing resin vanes of the same geometry and cooling configuration (to implement adiabatic wall thermal boundary conditions) and the transient liquid crystal technique, surface distributions of external heat transfer coefficient and film cooling effectiveness can be acquired. By combining these measurements with those from the metal vanes, the results can be scaled to engine conditions with a good level of accuracy. [ABSTRACT FROM AUTHOR]

Nozzle guide vane (NGV) flow capacity is perhaps the most important parameter for engine optimization. Inaccurate evaluation of capacity would lead to incorrect performance evaluation, and unmatched stages. A new semi-transient technique has been developed and demonstrated that will allow turbine designers to measure experimentally the effective throat area of an annular cascade of nozzle guide vanes at engine-representative Mach number, Reynolds number, and mainstream-to-cooling-flow pressure ratio. The technique allows NGV capacity to be measured to bias and precision uncertainties to 95% confidence of ±0.546% and ±0.028%, which compares well to large scale industrial facilities. Order-of-magnitude cost savings are offered over typical continuously running industrial facilities by running in blowdown mode from a receiver tank, thus removing the need for large scale compressor plant. To demonstrate the technique, a high mass flow rate blowdown tunnel was developed and commissioned at the University of Oxford, and the capacity of a high-pressure NGV from a large civil aircraft engine was experimentally determined. Experimental results are presented, which allow the precision error to be accurately calculated. A detailed uncertainty analysis is given from which the bias error is computed. It is shown that the low precision error the new technique offers means that it is ideally suited to investigations in which secondary influences on capacity are the subject of the investigation. The technique is of industrial significance because methods to compute engine capacity analytically or computationally do not yet provide sufficient accuracy for engine optimization, and the new technique offers equivalent accuracy at a much reduced cost over conventional experimental techniques. By performing an uncertainty analysis using experimental data it is shown that the increase in uncertainty due to the semi-transient (as opposed to steady state) nature of the technique is approximately 0.004% (to 95% confidence), and is entirely negligible. The experimentally measured trend of capacity against pressure ratio is compared with simple 1D, 2D, and 3D inviscid models, and an analytical correction for total pressure loss is performed. It is shown that while a simple 3D model is better than a 1D model (up to 1.5% improvement) for crude estimates of engine capacity, experimental trends are poorly predicted by such simple techniques. An analytical correction for total pressure loss increases the difference between 3D prediction and experiment. The experimental data demonstrate the complex nature of the process by which nozzle capacity is determined and the need for accurate, low-cost experimental techniques for capacity measurement. Correction to engine conditions is discussed. [ABSTRACT FROM AUTHOR]

A new hot-streak (combustor) simulator has been designed and implemented in a turbine test facility at QinetiQ Farnborough (the QinetiQ Isentropic Light Piston Facility, ILPF) to study the impact of temperature distortion on high pressure (HP) turbine efficiency, aerodynamics, and heat transfer. The ILPF is an engine scale, short duration, rotating transonic turbine test facility, in which M, Re, Tg/Tw, and The hot-streak simulator is a second-generation design, in which cold gas is introduced into a hot mainstream though radial and circumferential slots upstream of the turbine stage. The simulator is rotatable, so the effect of clocking (relative circumferential position of hot streak and nozzle guide vane leading edge) can easily be investigated. An emphasis was placed on accurate measurement of turbine inlet enthalpy flux so that the impact of hot streaks on turbine efficiency could be investigated. The hot-streak simulator differs from all previous systems in that a pronounced radial and circumferential temperature profile has been generated, with a hot-streak to vane count of 1:1. The profile is very well matched (non-dimensionally) to the target profile, which is a combustor temperature profile measured in a modern operating engine at the most extreme point in the cycle. The most accurate area survey of a simulated temperature profile has been conducted to date, and this demonstrates that the simulator offers an exceptionally high degree of circumferential symmetry and run-to-run repeatability. The design and commissioning of the simulator is described, and the measured temperature profiles are compared with the target profile. [ABSTRACT FROM AUTHOR]