Transonic Turbine Blade Research Articles

Endwall Heat Transfer Measurements in a Transonic Turbine Cascade

Turbine blade endwall heat transfer measurements are presented for a range of Reynolds and Mach numbers. Data were obtained for Reynolds numbers based on inlet conditions of 0.5 and 1.0 × 106, for isentropic exit Mach numbers of 1.0 and 1.3, and for free-stream turbulence intensities of 0.25 and 7.0 percent. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136 deg of turning and an axial chord of 12.7 cm. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for CFD code and model verification. The flow field in the cascade is highly three dimensional as a result of thick boundary layers at the test section inlet. Endwall heat transfer data were obtained using a steady-state liquid crystal technique.

Differential interferometry with adjustable spatial carrier fringes for turbine blade cascade flow investigations

Digital evaluated differential interferometry using adjustable spatial carrier fringes was applied to flow measurements in a transonic turbine blade cascade. The interferograms were evaluated using a two-dimensional Fourier Analysis. This evaluation provided density gradient maps of the flow field in a digital form. A specially designed interferometer was used allowing adjustment of sensitivity and superposition of carrier-fringe system separately. This type of interferometer is also highly insensitive to vibration noise. For a turbine blade cooling film a comparison of this type of differential interferometry with holographic interferometry and a Schlieren visualisation is also given. As a result differential interferometry using adjustable spatial carrier fringes and digital fringe analysis gave density-gradient maps similar to Schlieren recordings but of quantitative nature, thus enabling compensation of wave front distortions. Integration of these density gradient maps resulted in density maps which were then compared to recordings done with pulsed holographic interferometry.

Comparison of predicted and experimental Nusselt number for a film-cooled rotating blade

The predictions from a three-dimensional (3-D) Navier-Stokes code have been compared to the Nusselt number data obtained on a film-cooled, rotating turbine blade. The blade chosen is the ACE rotor with five rows containing 93 film-cooling holes covering the entire span. This is the only film-cooled rotating blade on which experimental heat transfer data are available for the present comparison. Over 2.25 million grid points are used to compute the flow over the blade. Usually, in a film cooling computation on a stationary blade, the computational domain is just one spanwise pitch of the film-cooling holes, with periodic boundary conditions in the span direction. However, for a rotating blade, the computational domain consists of the entire blade span from hub to tip, as well as the tip clearance region. As far as the authors know, the present work is the first comparison of the prediction of surface heat transfer using a 3-D Navier-Stokes code with film injection and the measured heat flux on a fully film-cooled rotating transonic turbine blade. A reasonably good comparison with the measured data is obtained on the suction surface, particularly near the hub section. On the pressure surface, however, the comparison between the data and the prediction is poor. A potential reason for the discrepancy on the pressure surface could be the presence of unsteady effects caused by stator-rotor interaction in the experiments, which are not modeled in the present numerical computations.

Comparison of predicted and experimental Nusselt number for a film-cooled rotating blade

The predictions from a three-dimensional (3-D) Navier-Stokes code have been compared to the Nusselt number data obtained on a film-cooled, rotating turbine blade. The blade chosen is the ACE rotor with five rows containing 93 film-cooling holes covering the entire span. This is the only film-cooled rotating blade on which experimental heat transfer data are available for the present comparison. Over 2.25 million grid points are used to compute the flow over the blade. Usually, in a film cooling computation on a stationary blade, the computational domain is just one spanwise pitch of the film-cooling holes, with periodic boundary conditions in the span direction. However, for a rotating blade, the computational domain consists of the entire blade span from hub to tip, as well as the tip clearance region. As far as the authors know, the present work is the first comparison of the prediction of surface heat transfer using a 3-D Navier-Stokes code with film injection and the measured heat flux on a fully film-cooled rotating transonic turbine blade. A reasonably good comparison with the measured data is obtained on the suction surface, particularly near the hub section. On the pressure surface, however, the comparison between the data and the prediction is poor. A potential reason for the discrepancy on the pressure surface could be the presence of unsteady effects caused by stator-rotor interaction in the experiments, which are not modeled in the present numerical computations.

Three-Dimensional Navier–Stokes Analysis and Redesign of an Imbedded Bellmouth Nozzle in a Turbine Cascade Inlet Section

Experimental measurements in the inlet of a transonic turbine blade cascade showed unacceptable pitchwise flow nonuniformity. A three-dimensional, Navier–Stokes computational fluid dynamics (CFD) analysis of the imbedded bellmouth inlet in the facility was performed to identify and eliminate the source of the flow nonuniformity. The blockage and acceleration effects of the blades were accounted for by specifying a periodic static pressure exit condition interpolated from a separate three-dimensional Navier–Stokes CFD solution of flow around a single blade in an infinite cascade. Calculations of the original inlet geometry showed total pressure loss regions consistent in strength and location to experimental measurements. The results indicate that the distortions were caused by a pair of streamwise vortices that originated as a result of the interaction of the flow with the imbedded bellmouth. Computations were performed for an inlet geometry that eliminated the imbedded bellmouth by bridging the region between it and the upstream wall. This analysis indicated that eliminating the imbedded bellmouth nozzle also eliminates the pair of vortices, resulting in a flow with much greater pitchwise uniformity. Measurements taken with an installed redesigned inlet verify that the flow nonuniformity has indeed been eliminated.

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.

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 Pilot 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 Pilot traverse and surface-mounted thin films confirm the essential features of the boundary layers.

Navier-Stokes analysis of flow and heat transfer inside high-pressure-ratio transonic turbine blade rows

A detailed numerical study of the flow and the heat transfer inside high-pressure-ratio transonic/supersonic turbine blade rows is described. A high-order upwinding relaxation method is used to solve the Reynoldsaveraged Navier-Stokes equation inside three different turbine blade rows. A two-equation turbulence model with a low-Reynolds-n umber modification is used to describe turbulent shear stress and turbulent heat flux. The numerical results are compared with available experimental data on two-dimensional geometries. It is found that very complex flow phenomena near the trailing edge of transonic turbine blade rows can be very well predicted numerically, and the method can be used in the development of new designs for HP turbines.

Effect of Incidence on Wall Heating Rates and Aerodynamics on a Film-Cooled Transonic Turbine Blade

This study investigates the influence of incidence on convective heat transfer to highly curved surfaces of a film-cooled turbine rotor blade. A computational study of free-stream inviscid aerodynamics without cooling at various incidences is followed by well-documented measured heat transfer data sets. The heat transfer experiments are discussed for cases with and without film cooling, performed under realistic gas turbine flow conditions in the short-duration heat transfer facility of the von Karman Institute for Fluid Dynamics. The precise location of the stagnation point and the iso-Mach number contours in the passage for each incidence (−10, 0, 10, +15 deg) are presented for a nominal exit Mach number of 0.94. The free-stream mass flow rate was kept constant for each experiment at different incidence levels. Three rows of compound angled discrete cooling holes are located near the leading edge in a showerhead configuration. Two rows of staggered discrete cooling holes are located on the suction side and a single row of cooling holes is located on the pressure side. The short-duration measurements of quantitative wall heat fluxes on nearly isothermal blade surfaces both in the presence and absence of coolant ejection are presented. The study indicated that the change of the position of the stagnation point strongly altered the aerodynamic behavior and convective heat transfer to the blade in approximately the first 30 percent of both the pressure side and the suction side in the presence and absence of film cooling. The immediate vicinity of the stagnation point was not significantly affected by changing incidence without cooling. Transitional behavior both on the suction surface and on the pressure surface was significantly influenced by the changes in approaching flow direction. Flow separation associated with incidence variations was also observed. Extremely low levels of the convective heat transfer coefficients were experienced near the regions where small separation bubbles are located.

An inviscid model for the base pressure of transonic turbine cascade

An inviscid base pressure model for transonic turbine blade has been presented. It has been shown that for a given back pressure the base pressure at the trailing edge, and the profile loss of a turbine blade are fixed according to the model and the base pressure can be calculated with the help of an inviscid numerical scheme. A parameteric study on the model shows that a blade profile with positive curvature downstream of the throat is advantageous for generating less loss, whilst the worst situation is when the exit flow reaches the sonic condition.

The Trailing Edge Loss of Transonic Turbine Blades

Trailing edge loss is one of the main sources of loss for transonic turbine blades, contributing typically 1/3 of their total loss. Transonic trailing edge flow is extremely complex, the basic flow pattern is understood but methods of predicting the loss are currently based on empirical correlations for the base pressure. These correlations are of limited accuracy. Recent findings that the base pressure and loss can be reasonably well predicted by inviscid Euler calculations are justified and explained in this paper. For unstaggered choked blading, it is shown that there is a unique relationship between the back pressure and the base pressure and any calculation that conserves mass, energy and momentum should predict this relationship and the associated loss exactly. For realistic staggered blading, which operates choked but with subsonic axial velocity, there is also a unique relationship between the back pressure and the base pressure (and hence loss) but the relationship cannot be quantified without knowing a further relationship between the base pressure and the average suction surface pressure downstream of the throat. Any calculation that conserves mass, energy and momentum and also predicts this average suction surface pressure correctly will again predict the base pressure and loss. Two-dimensional Euler solutions do not predict the suction surface pressure exactly because of shock smearing but nevertheless seem to give reasonably accurate results. The effects of boundary layer thickness and trailing edge coolant ejection are considered briefly. Coolant ejection acts to reduce the mainstream loss. It is shown that suction surface curvature downstream of the throat may be highly beneficial in reducing the loss of blades with thick trailing edges operating at high subsonic or low supersonic outlet Mach numbers.

Measurement and Modeling of the Gas Turbine Blade Transition Process as Disturbed by Wakes

Heat transfer measurements have been made on a transonic turbine blade undergoing natural transition and with a simulation of the effect of NGV wake interactions. The use of wide bandwidth heat transfer instrumentation permits the tracking of individual unsteady events that were identified as being due to either the impinging wakes or to the turbulent spots occurring within the transition process. Trajectories of these events as seen by the blade surface instrumentation have been measured. Numerical models have been developed for the effects of both types of turbulent activity. The convection of the wake through the passage is predicted, allowing for estimations of the expected times for which the boundary layer is disturbed by the wake fluid. The new model for the random generation and subsequent growth and convection of the turbulent spots produces a time-resolved prediction of the intermittent heat transfer signals by use of a time-marching procedure. By superimposing the two numerical models it is possible to simulate the measured instantaneous heat transfer characteristics.

Aerodynamic Loss Penalty Produced by Film Cooling Transonic Turbine Blades

Aerodynamic loss measurements are presented for a state-of-the-art film cooled transonic gas turbine rotor blade tested in a two-dimensional cascade. A mixture of carbon dioxide and air, which correctly simulated engine coolant-to-mainstream density ratio and blowing rate, was ejected from each of five individual cooling hole rows in the aerofoil suction surface. The temperature of the coolant was equal to the cascade inlet stagnation temperature. The dependence of blade row efficiency and turning on outlet Mach number, blowing rate, and coolant-to-mainstream density ratio was investigated. Measured surface static pressure distributions were compared with time-marching predictions for both the datum aerofoil and film cooled blades. Detailed suction surface boundary layer measurements both upstream and downstream of a cooling film were compared with available differential calculation procedures. Unexpectedly, films downstream of the throat, even at blowing rates near unity, did not generate significantly higher losses compared to prethroat suction surface films on this aerofoil.

Shock Wave/Boundary Layer Interaction in Dry and Wet Transonic Steam

As part of the continuing technological advance towards the design of transonic steam turbine blades by analysis, rather than by experience based on time-consuming field testing, a series of experiments in a one-dimensional Laval nozzle was conducted to determine the nature of self-excited fluctuations due to shock wave/boundary layer interaction. Results in superheated steam flow showed peak-to-peak pressure fluctuations reaching 80 percent of the ideal across a normal shock wave, with a frequency spectrum having a fairly flat response to 150 Hz. Results in wet flow showed a distinctly different spectrum with decreased activity up to 200 Hz, but a shock wave interacting with incipient condensation near the nozzle throat produced additional broad spectral strength at 500 and 700 Hz. In addition, shock wave fluctuation in supercooled flow could trigger condensation in advance of the natural Wilson point.

The Influence of Trailing Edge Ejection on the Base Pressure in Transonic Turbine Cascades

This paper summarizes the results of base pressure studies on transonic turbine blades in the presence of an ejection of coolant flow from a slot in the trailing edge. The first part of the paper reports on tests carried out on a enlarged model of the overhang section of a typical transonic cascade. This model provides valuable information about the detailed trailing edge pressure distribution and points to an asymmetric evolution of the base pressure on both sides of the slot in the presence of a bleed. The second part of the paper presents experimental results from cascade tests covering an outlet Mach number range M2, is = 0.5 to 1.35. These experiments underline the importance of the coolant flow impact on the base pressure and confirm the asymmetry of the base pressure with respect to the cooling slot. Tests with different coolant flow gases point to the significance of a proper simulation of the density ratio between coolant flow and main flow.

Compatibility of Transonic Flows at Thick Trailing Edge in Turbine Blade Row

In transonic turbine blades with thick trailing edge, the base pressure strongly influences the cascade performances. This paper presents two methods for predicting the base pressure, which are extension of the analyses on the backward facing step by Korst (neglecting upstream boundary layer) and McDonald (taking account of upstream boundary layer). The compatibility is considered between separated flows (jets) from suction and pressure surfaces in each method. The validity of both methods are discussed in comparison with the experiment of a two-dimensional turbine cascade with high stagger.

Peculiarities of magnetoacoustic resonance of ferrites (YIG nand YGdIG specimens): Burkov, V.D., Vashkovkii, A.V. and Kildishev V.N. Soviet Physics-Solid State, vol 10 no 12 (June 1969) pp 2975–2976 (2119)

Transonic turbine blade Over-Tip Leakage (OTL) flow has significant effects on turbine efficiency and blade thermal load. At present, most of the numerical studies on OTL flow are based on Reynolds-Averaged Navier-Stokes (RANS) method. However, it is difficult to characterize the physical mechanisms of transonic OTL flow accurately due to the inherent shortcomings of RANS method, especially for the unsteady characteristics. To understand the underlying loss mechanisms and aerothermal characteristics of the transonic OTL flow with tip gap variations, the present study applies a Very Large Eddy Simulation (VLES) method based on Wilcox k–ω model to study a transonic cascade OTL flow. The feasibility and accuracy of the VLES turbulence model in OTL flow simulations are confirmed by numerical assessments comparison with the experimental results. The detailed flow structure and its effect on the losses of tip leakage flow are studied for different cases with varying tip gap heights. It is observed that the flow structures including separation vortex and shock waves in the gap become more complex with increasing the tip gap. The mass flow rate, pressure loss, entropy generation, and aerothermal performance for different tip gap heights are further discussed.