M. L. G. Oldfield

An Examination of the Contributions to Loss on a Transonic Turbine Blade in Cascade

Experiments to measure losses of a linear cascade of transonic turbine blades are reported. Detailed measurements of the boundary layer at the rear of the suction surface of a blade and examination of wake traverse data enable the individual components of boundary layer, shock and mixing loss to be determined. Results indicate that each component contributes significantly to the overall loss in different Mach number regimes. Traverses in the near wake of the blade indicate the way in which the wake develops and facilitate examination of the development of the mixing loss.

Instantaneous Exhaust Temperature Measurements Using Thermocouple Compensation Techniques

This paper discusses a method of measuring the instantaneous exhaust gas temperature by thermocouples. Measuring the exhaust gas temperature is useful for a better understanding of engine processes. Thermocouples do not measure the instantaneous exhaust gas temperature because of their limited dynamic response. A thermocouple compensation technique has been developed to estimate the time constant in situ. This method has been commissioned in a simulation study and a controlled experiment with a reference temperature. The studies have shown that the signal bandwidth has to be restricted, since noise will be amplified in the temperature reconstruction. The technique has been successfully applied to some engine exhaust measurements. A comparison between two independent pairs of thermocouples has shown that temperature variations at frequencies up to 80Hz can be recovered. The medium load results agree with a previous study, which used fast response thermometers with a bandwidth of about 50 Hz. However, the results at low load and two different speeds have highlighted the need to do some 1-D unsteady flow simulations, in order to gain more insight into the exhaust process.

PIV Maps of Tip Leakage and Secondary Flow Fields on a Low-Speed Turbine Blade Cascade With Moving End Wall

New, detailed flow field measurements are presented for a very large low-speed cascade representative of a high-pressure turbine rotor blade with turning of 110 deg and blade chord of 1.0 m. Data were obtained for tip leakage and passage secondary flow at a Reynolds number of 4.0×10 5 , based on exit velocity and blade axial chord. Tip clearance levels ranged from 0% to 1.68% of blade span (0% to 3% of blade chord). Particle image velocimetry was used to obtain flow field maps of several planes parallel to the tip surface within the tip gap, and adjacent passage flow. Vector maps were also obtained for planes normal to the tip surface in the direction of the tip leakage flow. Secondary flow was measured at planes normal to the blade exit angle at locations upstream and downstream of the trailing edge. The interaction between the tip leakage vortex and passage vortex is clearly defined, revealing the dominant effect of the tip leakage flow on the tip end-wall secondary flow. The relative motion between the casing and the blade tip was simulated using a motor-driven moving belt system. A reduction in the magnitude of the undertip flow near the end wall due to the moving wall is observed and the effect on the tip leakage vortex examined.

PIV Maps of Tip Leakage and Secondary Flow Fields on a Low Speed Turbine Blade Cascade With Moving Endwall

New, detailed flow field measurements are presented for a very large low-speed cascade representative of a high-pressure turbine rotor blade with turning of 110 degrees and blade chord of 1.0 m. Data was obtained for tip leakage and passage secondary flow at a Reynolds number of 4.0 × 105 , based on exit velocity and blade axial chord. Tip clearance levels ranged from 0% to 1.68% of blade span (0% to 3% of blade chord). Particle Image Velocimetry (PIV) was used to obtain flow field maps of several planes parallel to the tip surface within the tip gap, and adjacent passage flow. Vector maps were also obtained for planes normal to the tip surface in the direction of the tip leakage flow. Secondary flow was measured at planes normal to the blade exit angle at locations upstream and downstream of the trailing edge. The interaction between the tip leakage vortex and passage vortex is clearly defined, revealing the dominant effect of the tip leakage flow on the tip endwall secondary flow. The relative motion between the casing and the blade tip was simulated using a motor-driven moving belt system. A reduction in the magnitude of the under-tip flow near the endwall due to the moving wall is observed and the effect on the tip leakage vortex examined.Copyright

Effect of Free-Stream Turbulence on Flat-Plate Heat Flux Signals: Spectra and Eddy Transport Velocities

An experimental study of the eddy structure in a flat-plate turbulent boundary layer with significant levels of free-stream turbulence is presented. This is relevant to the enhancement of turbomachinery heat transfer by turbulence and should lead to more realistic CFD modeling. Previous measurements showed that Nusselt numbers may be increased by up to 35%, and that this increase depended on turbulence integral length scale as well as intensity. The new results described here provide an insight into the mechanism responsible. Thin film gages and hot wires were used to take simultaneous high-frequency measurements of fluctuating heat transfer rates to the flat plate and the fluctuating flow velocity in the free stream and boundary layer. Spectra and correlation analysis shows that the turbulent eddy structure of the boundary layer is dominated by the free-stream turbulence at intensities of 3% and above. Eddies in the boundary layer mimicked those in the free stream and convected at the free-stream velocity U, rather than the {approximately} 0.8U characteristic of boundary layers. The main heat transfer enhancing mechanism is due to the penetration of free-stream turbulent eddies deep into the boundary layer, rather than enhancement of existing boundary layer turbulence.

Measurements of Hot Combustor Turbulence Spectra

This paper presents measurements of turbulence spectra at the exit plane of aircraft turbine combustors running at atmospheric exit pressure. These measurements are important in providing realistic input conditions for both experimental measurements and CFD predictions of first stage turbine heat transfer.A transient technique in which a pitot probe was only briefly exposed to the flow allowed uncooled, flush mounted sub-miniature pressure transducers to be used for measuring the turbulence spectra of combustor exhaust gases at temperatures up to 1500 K.Three different burner configurations were tested at fuel : air ratios from 0 to 0.02. It was found that the combustion process makes little difference to the turbulence power spectrum at wavenumbers between 100 and 1200 m−1. The effect of using two fuels with different burning rates ( paraffin and diesel ) was also studied.A brief description of the probe calibration is included. The response was found to be a function of both frequency and wavenumber.Copyright

Blade Tip Heat Transfer and Aerodynamics in a Large Scale Turbine Cascade With Moving Endwall

High resolution Nusselt number distributions were measured on the blade tip surface of a large, 1.0 m chord, low-speed cascade representative of a high-pressure turbine. Data were obtained at a Reynolds number of 4.0×105 based on exit velocity and blade axial chord. Tip clearance levels ranged from 0.56% to 1.68% design span or equally from 1% to 3% of the blade chord. An infrared camera, looking through the hollow blade, made detailed temperature measurements on a constant heat flux tip surface. The relative motion between the endwall and the blade tip was simulated by a moving belt. The moving belt endwall significantly shifts the region of high Nusselt number distribution and reduces the overall averaged Nusselt number on the tip surface by up to 13.3%. The addition of a suction side squealer tip significantly reduced local tip heat transfer and resulted in a 32% reduction in averaged Nusselt number. Analysis of pressure measurements on the blade airfoil surface and tip surface along with particle image velocimetry velocity flow fields in the gap gives an understanding of the heat transfer mechanism.

Calibration of a four-hole pyramid probe and area traverse measurements in a short-duration transonic turbine cascade tunnel

A four-hole pyramid probe has been calibrated for use in a short-duration transonic turbine cascade tunnel. The probe is used to create area traverse maps of total and static pressure, and pitch and yaw angles of the flow downstream of a transonic annular cascade. This data is unusual in that it was acquired in a short-duration (5 s of run time) annular cascade blowdown tunnel. A four-hole pyramid probe was used which has a 2.5 mm section head, and has the side faces inclined at 60° to the flow to improve transonic performance.The probe was calibrated in an ejector driven, perforated wall transonic tunnel over the Mach number range 0.5–1.2, with pitch angles from -20° to + 20° and yaw angles from-23° to +23°. A computer driven automatic traversing mechanism and data collection system was used to acquire a large probe calibration matrix (∼ 10,000 readings) of non dimensional pitch, yaw, Mach number, and total pressure calibration coefficients. A novel method was used to transform the probe calibration matrix of the raw coefficients into a probe application matrix of the physical flow variables (pitch, yaw, Mach number etc.). The probe application matrix is then used as a fast look-up table to process probe results. With negligible loss of accuracy, this method is faster by two orders of magnitude than the alternative of global interpolation on the raw probe calibration matrix.The blowdown tunnel (mean nozzle guide vane blade ring diameter 1.1 m) creates engine representative Reynolds numbers, transonic Mach numbers and high levels (≈ 13%) of inlet turbulence intensity. Contours of experimental measurements at three different engine relevant conditions and two axial positions have been obtained. An analysis of the data is presented which includes a necessary correction for the finite velocity of the probe. Such a correction is non trivial for the case of fast moving probes in compressible flow.

A Method for Correlating the Influence of External Crossflow on the Discharge Coefficients of Film Cooling Holes

An empirical means of predicting the discharge coefficients of film cooling holes in an operating engine has been developed. The method quantifies the influence of the major dimensionless parameters, namely hole geometry, pressure ratio across the hole, coolant Reynolds number, and the free stream Mach number. The method utilizes discharge coefficient data measured on both a first-stage high-pressure nozzle guide vane from a modern aero-engine and a scale (1.4 times) replica of the vane. The vane has over 300 film cooling holes, arranged in 14 rows. Data was collected for both vanes in the absence of external flow. These noncrossflow experiments were conducted in a pressurized vessel in order to cover the wide range of pressure ratios and coolant Reynolds numbers found in the engine. Regrettably, the proprietary nature of the data collected on the engine vane prevents its publication, although its input to the derived correlation is discussed. Experiments were also conducted using the replica vanes in an annular blowdown cascade which models the external flow patterns found in the engine. The coolant system used a heavy foreign gas (SF 6 /Ar mixture) at ambient temperatures which allowed the coolant-to-mainstream density ratio and blowing parameters to be matched to engine values. These experiments matched the mainstream Reynolds and Mach numbers and the coolant Mach number to engine values, but the coolant Reynolds number was not engine representative (Rowbury, D. A., Oldfield, M. L. G., and Lock, G. D., 1997, Engine-Representative Discharge Coefficients Measured in an Annular Nozzle Guide Vane Cascade, ASME Paper No. 97-GT-99, International Gas Turbine and Aero-Engine Congress & Exhibition, Orlando, Florida, June 1997; Rowbury, D. A., Oldfield, M. L. G., Lock, G. D., and Dancer, S. N., 1998, Scaling of Film Cooling Discharge Coefficient Measurements to Engine Conditions, ASME Paper No. 98-GT-79, International Gas Turbine and Aero-Engine Congress & Exhibition, Stockholm, Sweden, June 1998). A correlation for discharge coefficients in the absence of external crossflow has been derived from this data and other published data. An additive loss coefficient method is subsequently applied to the cascade data in order to assess the effect of the external crossflow. The correlation is used successfully to reconstruct the experimental data. It is further validated by successfully predicting data published by other researchers. The work presented is of considerable value to gas turbine design engineers as it provides an improved means of predicting the discharge coefficients of engine film cooling holes.

Detailed boundary layer measurements on a transonic turbine cascade

The boundary layers of a transonic turbine blade have been measured in detail. The full velocity profiles have been measured at a number of stations on both the suction and pressure surfaces, at conditions representative of engine operation, using a Pitot traverse technique and a large-scale (300 mm chord) linear cascade. This information has made it possible to follow the development of the boundary layers, initially laminar, through a region of natural transition to a fully developed turbulent layer. Comparisons with other, less detailed, measurements on the same profile using Pitot traverse and surface-mounted thin films confirm the essential features of the boundary layers.