Cooling Air Injection Research Articles

Investigation on the differences in unsteady film cooling behaviors of gas turbine blades between mainstream and cooling air pulsations for a cylindrical hole

In practice, film cooling on gas turbine blade inevitably works in an unsteady environment, which is introduced by periodical rotor/stator interaction and unsteady combustion. Previous studies have introduced two methods to simulate the realistic unsteady condition: 1) the pulsation of mainstream velocity; and 2) the pulsation of coolant injection. However, up to this point, the differences in instantaneous film cooling behaviors between these two methods remain unclear. This work presents a series of large eddy simulations to exhibit the unsteady flow and film cooling behaviors under steady and the two unsteady flow conditions. The numerical strategy is validated against our time-resolved experimental data. Time-averaged results show that the difference between the two pulsations is not significant if the averaged blowing ratio remains the same. However, the pulsation type plays a dominant role on the transient mode of film coverage. Under the steady condition, film coverage instability is induced by the unsteady trajectory of near-wall vortex structure; but with pulsed environments, the unsteadiness magnitude increases, and the area with high unsteadiness level enlarges. The pulsation of the mainstream velocity induces a more severe film coverage instability compared to the pulsation of the cooling air injection, because of the higher fluctuation energy of the mainstream bulk. Under mainstream pulsation, the probability distribution of instantaneous cooling effectiveness is the most scattered, and the corresponding fluctuation range is the largest.

Effect of pulsed coolant injection on unsteady film cooling characteristics with serrated trench design

To increase cooling effectiveness, film-hole designs with serrated trenches have been widely reported in the gas turbine industry. However, the unsteadiness level of the film coverage and the inevitable environmental pulsation are always neglected, which may lead to an unexpected failure of hot components by the instantaneous exposure to a high-temperature environment. In this work, a series of large eddy simulations are conducted to comprehensively evaluate the transient film cooling performance with serrated trench designs, and the effect of pulsed cooling air injection is discussed. Time-averaged results reveal that: 1) under steady condition, serrated trench design exhibits a higher film effectiveness in most regions compared to the traditional transverse trench, which supports the previously reported advantage. 2) Pulsed cooling air injection would cause a significant reduction of film effectiveness for the serrated trenches. Under the pulsed condition, film effectiveness with a serrated angle of 100° can be over 20% lower than that with a traditional transverse trench. Moreover, transient results indicate that under both steady and pulsed conditions, the serrated trenches always exhibit a lower steadiness level of film cooling, relative to the traditional transverse trench. Therefore, it should be prudent to suggest film cooling with serrated trenches in practical designs.

Experimental and Numerical Study of Transonic Cooled Turbine Blades

State-of-the-art gas turbines (GT) operate at high temperatures that exceed the endurance limit of the material, and therefore the turbine components are cooled by the air taken from the compressor. The cooling provides a positive impact on the lifetime of GT but has a negative impact on its performance. In convection-cooled turbine blades the coolant is usually discharged through the trailing edge and leads to limitations on the minimal size of the trailing edge, thereby negatively affecting the losses. Moreover, the injection of cooling air in the turbine disturbs the main flow, and may lead to an additional increase in loss. Trailing edge loss is a significant part of the overall loss in modern gas turbines. This study comprises investigations of the unguided flow angle, the trailing edge shape, and cooling air injection through the trailing edge on the base pressure and profile losses in cooled blades. Some vane and blade cascades with different unguided turning angle and two shapes of trailing edges with and without coolant injection were studied both experimentally and numerically. This analysis provides a split of losses caused by different factors, and offers opportunities for efficiency and lifetime improvements of real engine designs/upgrades. In particular, it is shown that an increase in the unguided turning angle and the use of a round trailing edge result in a reduction of loss in case of a relatively thick trailing edge. Numerical investigation showed that an increase in the unguided turning angle at the initial transonic vane with a thick and blunt trailing edge, without a change in other basic geometric parameters, allowed for a significant reduction of the profile loss by about 3–4% at the exit Mach number M2is = 0.7–1.0. Experimental investigation of four cascades with cooling air injection into the base flow through the trailing edge allowed us to validate the fact that in blades with a low level of base pressure Cpb < −0.1 at m¯ = 0 a non-monotonic dependence of the change of losses against relative cooling air mass flow m¯ is observed. Firstly, the cooling air injection into wake increases base pressure and decreases losses; then the losses start to increase with increasing cooling mass flow due to the interaction between the main flow and the cooling air (mixing losses) and, finally, due to the cooling mass flow increase and momentum increase losses are decreased. In blades with an increased level of the base pressure coefficient Cpb ≥ −0.1 at m¯ = 0 the cooling air injection results in an increase in losses right from the beginning of the injection and then, according to the cooling mass flow increase and momentum rise, losses decrease. It is also shown that injection through the trailing edge slot parallel to the main flow leads to a neutral loss impact and even a loss reduction in the subsonic range and a loss increase in the supersonic range of exit Mach numbers.

An experimental study of flow and heat transfer performance of a longitudinal corrugated liner for a combustion chamber

Experiments are conducted to investigate the flow and heat transfer performance of a longitudinal corrugated liner for a combustion chamber. The effect of cooling air passage height and blockage ratio on the heat transfer performance of the liner has been studied. In addition, the friction factors in cooling air passages are also presented and the test results are compared with previously reported studies. In the experiments, the pressure fields having the same conditions as that present in the real engine are established. The detailed temperature distributions on the hot side of the liner are measured by using the infrared imaging technique. Local film cooling effects are calculated from these temperature distributions. A range of parameter combinations of interest in cooled liner practice is covered, including combinations of variations in Reynolds number of cooling air passage, passage height and blockage ratio. The results indicate that the blockages in the cooling air passage are not beneficial for obtaining a high film cooling effect and for the uniform distribution of film cooling effect in the streamwise direction. The cooling air flow will be disturbed by the blockages located in the cooling air passage which weakens the effect of dynamic pressure on cooling air injection. The local and downstream film cooling effect will be significantly influenced. The friction factor increases with the increase of bolckage ratio. Compared with previous works, the friction factors in the present conditions are much higher (at least by 133%) than that for parallel plat passage and pipes with rough walls. As the passage height increases, the film cooling effect increases first and then decreases. However, the relative deviation decreases first and then increases. That means there is an optimum passage height for a higher film cooling effect and a uniform distribution. The passage height has an significant effect on the momentum of cooling air injection and the distribution of cooling air flow rate through the windward side of the crest and the leeward side of the crest. The effect of passage height on the friction factor is slight. The f-Re correlations are proposed based on the present experimental data.

Continuous adjoint aerodynamic optimization design for multi-stage gas turbine with cooling air

PurposeThis paper aims to conduct the optimization of the multi-stage gas turbine with the effect of the cooling air injection based on the adjoint method.Design/methodology/approachContinuous adjoint method is combined with the S2 surface code.FindingsThe optimization of the stagger angles, stacking lines and the passage can improve the attack angles and restrain the development of the boundary, reducing the secondary flow loss caused by the cooling air injection.Practical implicationsThe aerodynamic performance of the gas turbine can be improved via the optimization of blade and passage based on the adjoint method.Originality/valueThe results of the first study on the adjoint method applied to the S2 surface through flow calculation including the cooling air effect are presented.

Numerical investigations of pulsed film cooling on an entire turbine vane

This paper presents a numerical investigation of pulsed film cooling on a modified NASA C3X vane, which has five rows of film cooling hole: three rows at leading edge, and the other two at pressure side and suction side, respectively. Square and sinusoidal waves are considered to pulse cooling air injection. The performances of film cooling effectiveness are discussed at three blowing ratios (0.5, 0.75 and 1.0) and four Strouhal numbers (0.0027, 0.0054, 0.0108 and 0.0216). Previous investigations of pulsed film cooling were limited within flat plates or semicircular cylinders, therefore it is necessary to provide a comprehensive reference regarding pulsed film cooling on an entire turbine vane. Through the present numerical simulations conducted on the entire turbine vane, the following interesting phenomena are discovered: 1) at leading edge and pressure side, film cooling effectiveness decreases when blowing ratio or Strouhal number increases, but at suction side, the trend reverses; 2) sinusoidal wave pulsed flow leads to less mainstream ingestion into film hole than square wave pulsed flow; 3) in the leading edge region, sinusoidal wave pulsed flow exhibits higher film cooling effectiveness than steady flow at blowing ratios of 0.75 and 1.0.

Aerodynamic optimization design for high pressure turbines based on the adjoint approach

A first study on the continuous adjoint formulation for aerodynamic optimization design of high pressure turbines based on S2 surface governed by the Euler equations with source terms is presented. The objective function is defined as an integral function along the boundaries, and the adjoint equations and the boundary conditions are derived by introducing the adjoint variable vectors. The gradient expression of the objective function then includes only the terms related to physical shape variations. The numerical solution of the adjoint equation is conducted by a finite-difference method with the Jameson spatial scheme employing the first and the third order dissipative fluxes. A gradient-based aerodynamic optimization system is established by integrating the blade stagger angles, the stacking lines and the passage perturbation parameterization with the quasi-Newton method of Broyden–Fletcher–Goldfarb–Shanno (BFGS). The application of the continuous adjoint method is validated through a single stage high pressure turbine optimization case. The adiabatic efficiency increases from 0.8875 to 0.8931, whilst the mass flow rate and the pressure ratio remain almost unchanged. The optimization design is shown to reduce the passage vortex loss as well as the mixing loss due to the cooling air injection.

Impact of Cooling Air Injection on the Combustion Stability of a Premixed Swirl Burner Near Lean Blowout

In most dry, low-NOx combustor designs of stationary gas turbines, the front panel impingement cooling air is directly injected into the combustor primary zone. This air partially mixes with the swirling flow of premixed reactants from the burner and reduces the effective equivalence ratio in the flame. However, local unmixedness and the lean equivalence ratio are supposed to have a major impact on combustion performance. The overall goal of this investigation is to answer the question of whether the cooling air injection into the primary combustor zone has a beneficial effect on combustion stability and NOx emissions or not. The flame stabilization of a typical swirl burner with and without front panel cooling air injection is studied in detail under atmospheric conditions close to the lean blowout limit (LBO) in a full-scale, single-burner combustion test rig. Based on previous isothermal investigations, a typical injection configuration is implemented for the combustion tests. Isothermal results of experimental studies in a water test rig adopting high-speed planar laser-induced fluorescence (HSPLIF) reveal the spatial and temporal mixing characteristics for the experimental setup studied under atmospheric combustion. This paper focuses on the effects of cooling air injection on both flame dynamics and emissions in the reacting case. To reveal dependencies of cooling air injection on combustion stability and NOx emissions, the amount of injected cooling air is varied. OH*-chemiluminescence measurements are applied to characterize the impact of cooling air injection on the flame front. Emissions are collected for different cooling air concentrations, both global measurements at the chamber exit, and local measurements in the region of the flame front close to the burner exit. The effect of cooling air injection on pulsation level is investigated by evaluating the dynamic pressure in the combustor. The flame stabilization at the burner exit changes with an increasing degree of dilution with cooling air. Depending on the amount of cooling, only a specific share of the additional air participates in the combustion process.

Through-Flow Analysis of Air-Cooled Gas Turbines

This paper describes the development of a new through-flow method for the analysis of axial multistage turbines with cooling by air from compressor bleed. The method is based on a stream function approach and a finite element solution procedure. It includes a high-fidelity loss and deviation model with improved correlations. A radial distribution model of losses and a new spanwise mixing model are applied to simulate 3D flow effects. The calibration of the models is performed by calculation of a number of test cases with different configurations, with the aim of achieving high accuracy and optimum robustness for each of the test cases considered. Various types of cooling air injection were encompassed: film cooling, trailing edge injection, and disk/endwall coolant flow. There are two effects of air cooling: (i) increase in mass flow downstream of the injection surface and (ii) reduction of the gas total temperature connected with total pressure losses. For both of these effects, the appropriate 2D models were developed and applied. The code was applied to flow analysis and performance prediction of a newly developed industrial gas turbine. Comparison of the predicted results and measured test data for a number of parameters showed good agreement. The results of the validation confirmed that this method based on calibrated correlations can be considered a reliable tool for flow analysis and parameter variation during the design phase.

Impact of Cooling Air Injection on the Primary Combustion Zone of a Swirl Burner

In most dry, low NOx combustor designs, the front panel impingement cooling air is directly injected into the combustor primary zone. As this air partially mixes with the swirling flow of premixed reactants from the burner prior to completion of heat release, it reduces the effective equivalence ratio in the flame and has a beneficial effect on NOx emissions. However, the fluctuations of the equivalence ratio in the flame potentially increase heat release fluctuations and influence flame stability. Since both effects are not yet fully understood, isothermal experiments are made in a water channel, where high speed planar laser-induced fluorescence (HSPLIF) is applied to study the cooling air distribution and its fluctuations in the primary zone. In addition, the flow field is measured with high speed particle image velocimetry (HSPIV). Both mixing and flow field are also analyzed in numerical studies using isothermal large eddy simulation (LES), and the simulation results are compared with the experimental data. Of particular interest is the influence of the injection configuration and cooling air momentum variation on the cooling air penetration and dispersion. The spatial and temporal quality of mixing is quantified with probability density functions (PDF). Based on the results regarding the equivalence ratio fluctuations, regions with potential negative effects on combustion stability are identified. The strongest fluctuations are observed in the outer shear layer of the swirling flow, which exerts a strong suction effect on the cooling air. Interestingly, the cooling air dilutes the recirculation zone of the swirling flow. In the reacting case, this effect is expected to lead to a decrease of the temperature in the flame-anchoring zone below the adiabatic flame temperature of the premixed reactant, which may have an adverse effect on flame stability.

Turbine Stator Well CFD Studies: Effects of Coolant Supply Geometry on Cavity Sealing Performance

The increase of aeroengine performance through the improvement of aerodynamic efficiency of core flow is becoming more and more difficult to achieve. However, there are still some devices that could be improved to enhance global engine efficiency. Particularly, investigations on the internal air cooling systems may lead to a reduction of cooling air with a direct benefit to the overall performance. At the same time, further investigations on heat transfer mechanisms within turbine cavities may help to optimize cooling air flows, saving engine life duration. This paper presents a computational fluid dynamics (CFD) study aimed at the characterization of the effects of different geometries for cooling air supply within turbine cavities on wall thermal effectiveness and sealing mass flow rate. Several sealing air supply geometries were considered in order to point out the role of cooling air injection position, swirl number, and jet penetration on the cavities’ sealing performance. Steady state calculations were performed using two different computational domains: the first consists of a sector model of the whole turbine including the second stator well, while the second is a cut-down model of the stator well. Thanks to the simplified geometry of the test rig with respect to actual engines, the study has pointed out clear design suggestions regarding the effects of geometry modification of cooling air supply systems.

Full coverage film cooling using compound angle

An experimental and numerical study is carried out on a cooling film issuing from a multiperforated wall of a simplified combustor. The objectives of this work are to achieve a better understanding of the dynamics of the film and to construct an experimental database on a simplified geometry in order to test numerical models. A parametric study of film cooling efficiency based on the direction of the cooling air injection is presented and shows that a swirling injection greatly enhances the cooling efficiency. As accounting for multiperforated walls in numerical simulations cannot be done at the jets scale because of computing resources, in this article are presented RANS computations performed using a uniform boundary condition to provide the injection of coolant. Two injection models are applied on this boundary and numerical results are compared to experimental data in the recovery region. The standard model is shown to be totally inappropriate while the multiperforation model delivers promising results although some weaknesses appear very close to the wall. To cite this article: B. Michel et al., C. R. Mecanique 337 (2009).

Control of Rotor Tip Leakage Through Cooling Injection From the Casing in a High-Work Turbine

This paper presents an experimental investigation of a novel approach for controlling the rotor tip leakage and secondary flow by injecting cooling air from the stationary casing onto the rotor tip. It contains a detailed analysis of the unsteady flow interaction between the injected air and the flow in the rotor tip region and its impact on the rotor secondary flow structures. The experimental investigation has been conducted on a one-and-1/2-stage, unshrouded turbine, which has been especially designed and built for the current investigation. The turbine test case models a highly loaded, high pressure gas turbine stage. Measurements conducted with a two-sensor fast-response aerodynamic probe have provided data describing the time-resolved behavior of flow angles and pressures, as well as turbulence intensity in the exit plane of the rotor. Cooling air has been injected in the circumferential direction at a 30 deg angle from the casing tangent, opposing the rotor turning direction through a circumferential array of ten equidistant holes per rotor pitch. Different cooling air injection configurations have been tested. Injection parameters such as mass flow, axial position, and size of the holes have been varied to see the effect on the rotor tip secondary flows. The results of the current investigation show that with the injection, the size and the turbulence intensity of the rotor tip leakage vortex and the rotor tip passage vortex reduce. Both vortices move toward the tip suction side corner of the rotor passage. With an appropriate combination of injection mass flow rate and axial injection position, the isentropic efficiency of the stage was improved by 0.55 percentage points.

An experimental study of film cooling in a negatively bowed cascade with a large turning angle

Transverse measurements are carried out at the outlet of the straight and negatively bowed turbine cascades with a 106° turning angle respectively with and without cooling air injection. The air injection is from single or multiple rows of ten holes (26 schemes in all). Experimental results show that the negatively bowed cascade produces more energy loss than the straight one without cooling air injection. Cooling air injection from the holes on both the pressure and the suction surface near the trailing edge can reduce energy losses in both the straight and the negatively bowed cascades. The increase in loss is less in the negatively bowed cascade than that in the straight one with cooling air injection from holes of multiple rows at the leading edge, whereas cooling air injection from holes of multiple rows at the trailing edge reduces the energy loss in the negatively bowed cascade. © 2004 Wiley Periodicals, Inc. Heat Trans Asian Res, 34(1): 1–8, 2005; Published online in Wiley InterScience (www.interscience.wiley.com). DOI 10.1002/htj.20042

Étude en soufflerie thermique du refroidissement de parois poreuses par effusion de gaz

A heated air tunnel and specially its test section features, designed for studies of wall protection from high temperature gases, are described. A coolant, moving through the porous wall, flush fitted in the horizontal test section floor, achieves its thermal protection. Preliminary velocity measurements, carried out at ambient temperature and without injection, within the test section, pointed out the following aerodynamic characteristics: stationarity and turbulence intensity in the potential flow, boundary layer features over the porous wall according to the entrance test section boundary conditions, wall friction coefficient estimation. Velocity and temperature profiles, in non-isothermal conditions, over the porous wall with or without cooling air injection, are reported. They point out the injection rate influence, variable from 0 to 2 %.

Measured thrust losses associated with secondary air injection through nozzle walls

The experimental program discussed herein addressed the aerodynamic effects of auxiliary air on the thrust of a scaled two-dimensional converging/diverging nozzle. The tests were intended as a means of validating certain analytical and computational fluid dynamics (CFD) models for the thrust loss associated with cooling air injection. The ratios of injected flow rate to mainstream flow rate employed are typical of those used to cool the internal surfaces of high-performance jet aircraft nozzles. The primary purpose of the investigation was to establish a data base of nozzle performance sensitivity to various secondary air parameters. Divergent section pressure distributions obtained in the presence of across the flaps indicate that the injected flow constricts the primary flow thus reducing its effective expansion ratio. The results obtained also show a definite dependence of performance on the angle at which the secondary air is injected. The thrust coefficient changes resulting from injection, relative to the no injection situation, were reasonably predicted by a simple one-dimensional formulation corrected for certain two-dimensional flow effects. This formulation was observed to apply for injected flow rates up to about 6% of the primary flow rate. Nomenclature A = area Cd = discharge coefficient Q = coefficient used to modify one-dimensional stream thrust

Cooling-Air Injection Into Secondary Flow and Loss Fields Within a Linear Turbine Cascade

In order to understand overall performance and internal flows of air-cooled turbine blade rows, flows in a model linear cascade were surveyed with secondary air injection from various locations of the blade surfaces. The secondary air interacted with the cascade passage vortices and changed the loss distribution significantly. The cascade overall loss decreased when the air was injected along the mainstream and increased when the air was injected against the mainstream from some locations of the blade leading edge. Effects on overall kinetic energy of the secondary flows and on the cascade outlet flow angle were also discussed in this paper.

Suppressing the Infrared Signatures of Marine Gas Turbines

The exhaust plumes and visible areas of the engine exhaust ducting associated with marine gas turbines are major sources of infrared (IR) radiation on ships. These high-radiance sources make excellent targets for IR-guided threats. In recent years significant efforts have been made to reduce or eliminate these high-radiance sources to increase the survivability of naval and commercial ships when sailing in high-risk areas of the world. Typical IR signature suppression (IRSS) systems incorporate film cooling of visible metal sources, optical blockage to eliminate direct line-of-sight visibility of hot exhaust system parts, and cooling air injection and mixing for plume cooling. Because the metal surfaces radiate as near black bodies, every attempt is made to reduce the temperatures of the visible surfaces to near ambient conditions. The exhaust gases radiate selectively and therefore do not have to be cooled to the same degree as the metal surfaces. The present paper briefly describes the motivation for incorporating IRSS into the exhaust systems of marine power plants. IRSS hardware developed in Canada by the Canadian Department of National Defence and Davis Engineering Limited is presented along with details of their operating principles. A typical installation is presented and discussed. Design impacts on the ship are described with reference to engine back pressure, noise, and weight and center of gravity effects.

Investigation of the Aerodynamic Performance of Small Axial Turbines

This paper describes results of an experimental investigation of small axial turbine performance characteristics. Included are test data on the effects of the following design variables on small turbine aerodynamic efficiency: (a) blade height, (b) vane endwall contouring, (c) blade reaction, (d) blade tip clearance, (e) stage work, and (f) vane and blade airfoil row solidity. In addition, the effects of vane, blade, and disk cooling air injection on turbine efficiency are presented. The turbines evaluated were single stage, low aspect ratio configurations sized for airflows of 8 pps (3.63 kg/sec) or less and designed for inlet temperatures in the 2200-to-2500 deg F (1204-to-1371 deg C) range. The efficiency data presented in the paper cover both design and off-design velocity and pressure ratios. These data illustrate that relatively high efficiencies can be obtained in small, low aspect ratio axial turbines with an optimum design.