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This paper presents a method for the robustness analysis of the bladed disk with bounded random mistuning. The robust stability and performance are evaluated by the upper bound of the structured singular value. The robust control model of the bladed disk is established in virtue of linear fractional transformation. The influences of intentional stiffness mistuning in harmonic patterns on the robustness of the mistuned bladed disk are investigated. The numerical results indicate that the robust performance of the mistuned bladed disk could be effectively enhanced by appropriate harmonic intentional mistuning. The proposed method can help us design a bladed disk, which is insensitive to dangerous random mistuning. [DOI: 10.1115/1.3018942]

A study is presented on the fluid-dynamic pulsations and the corresponding dynamic forces generated in a centrifugal pump with single suction and vaneless volute due to blade-volute interaction. Four impellers with different outlet diameters, obtained from progressive cutbacks (trimmings) of the greatest one, were successively considered in the test pump, so that the radial gap between the impeller and the volute ranged from 8.8% to 23.2% of the impeller radius. The study was based on the numerical computation of the unsteady flow through the machine for a number of flow rates by means of the FLUENT code, solving the 3D unsteady Reynolds-averaged Navier-Stokes equations. Additionally, an experimental series of tests was conducted for the pump with one of the impellers, in order to obtain pressure fluctuation data along the volute front wall that allowed contrasting the numerical predictions. The data collected from the numerical computations were used to estimate the dynamic radial forces and torque at the blade-passing frequency, as a function of flow rate and blade-tongue radial gap. As expected, for a given impeller diameter, the dynamic load increases for off-design conditions, especially for the low range of flow rates, whereas the progressive reduction of the impeller-tongue gap brings about corresponding increments in dynamic load. In particular, varying the bladetongue gap within the limits of this study resulted in multiplying the maximum magnitude of the blade-passing frequency radial force by a factor of about 4 for low flow rates (i.e., below the nominal flow rate) and 3 for high flow rates. [DOI: 10.1115/1.2969273] Keywords: centrifugal pump, volute casing, blade-tongue interaction, unsteady flow simulation, pressure fluctuation measurements, blade-passing frequency load, effect of impeller-tongue radial gap

The investigation focuses on the analysis of the airfoil segment performances along rotor blades in the parked configuration. In this research, wind tunnel experiments on two twisted blade geometries with different airfoils played a dominant role. These measurements were carried out by the Swedish Aeronautical Research Institute, former FFA, and by the American National Renewable Energy Laboratories (NREL) during the Unsteady Aerodynamic Experiment. The spans of the blades were 2.375 m and 5 m, the STORK 5 WPX and the NREL Phase VI blade, respectively. Five span locations (inboard, midspan, outboard, and tip regions) were considered and compared with the 2D airfoil characteristics. Wing model experiments with similar blade aspect ratio were included in the research. Furthermore, the commercial computational fluid dynamics code FLUENT was used for the validation and analysis of the spanwise lift and drag coefficients at four different pitch settings, 20 deg, 30 deg, 45 deg, and 60 deg. The computed pressure distributions compared reasonably well, but the derived lift and drag showed quite some differences with the blade measurements. The lift coefficients for the sections beyond the leading-edge stall angle of the STORK blade were larger than for the NREL blade and were close to that of a wing model with similar airfoil and aspect ratio. Lift and drag coefficients for the sections of the two blades were always much smaller than the 2D results. The drag values for both blades showed quite some agreement, and airfoil and blade dependency seemed to be small. [DOI: 10.1115/1.2931507]

This paper analyzes the structure of turbulence and secondary, flows at the exit of an axial flow fan with variable pitch blades. The influence of changing the blades' pitch angle over the turbulent structures is assessed by means of turbulence intensity values and integral length scales, obtained by using hot-wire anemometry for several test conditions. Since total unsteadiness is composed of both periodic and random unsteadiness, it is necessary to filter deterministic unsteadiness from the raw velocity traces in order to obtain turbulence data. Consequently, coherent flow structures were decoupled and thus, levels of turbulence--rms values of random fluctuations--were determined using a filtering procedure that removes all the contributions stemming from the rotational frequency, the blade passing frequency, and its harmonics. The results, shown in terms of phase-averaged distributions in the relative frame of reference, revealed valuable information about the transport of the turbulent structures in the unsteady, deterministic flow patterns. The anisotropic turbulence generated at the shear layers of the blade wakes was identified as a major mechanism of turbulence generation, and significant links between the blade pitch angle and the wake turbulent intensity were established. In addition, the autocorrelation analysis of random fluctuations was also used to estimate integral length scales--larger eddy sizes--of turbulence, providing useful data for computational fluid dynamics applications based on large eddy simulation algorithms. Finally, contours of radial vorticity and helicity gave a detailed picture of the vortical characteristics of the flow patterns, and the definition of secondary flow as the deviation of the streamwise component from the inviscid kinematics was introduced to determine the efficiency of the blade design in the energy exchange of the rotor. [DOI: 10.1115/1.2903523]

At present, optimization is an enabling technology in innovation. Multi-objective and multidisciplinary optimization tools are essential in the design process for real-world applications. In turbomachinery design, these approaches give insight into the design space and identify the tradeoffs between the competing performance measures. This paper describes the application of a novel multi-objective variant of the tabu search algorithm to the aerodynamic design optimization of turbomachinery blades. The aim is to improve the performance of a specific stage and eventually of the whole engine. The integrated system developed for this purpose is described. It combines the optimizer with an existing geometry parameterization scheme and a well-established computational fluid dynamics package. Its performance is illustrated through a case study in which the flow characteristics most important to the overall performance of turbomachinery blades are optimized.

The present study is focused on the analysis of the dynamic and periodic interaction between both fixed and rotating blade rows in a single stage, low-speed axial fan with inlet guide vanes. The main goal is placed on the characterization of the unsteady flow structures involved in an axial flow fan of high reaction degree, relating them to working point variations and axial gap modifications. For that purpose, an experimental openloop facility has been developed to obtain a physical description of the flow across the turbomachine. Using hot-wire anemometry, measurements of axial and tangential velocities were carried out in two transversal sectors: one between the rows and the other downstream of the rotor, covering the whole span of the stage for a complete stator pitch. Ensemble- and time-averaging techniques were introduced to extract deterministic fluctuations from raw data, both of which are essential to understand flow mechanisms related to the blade passing frequency. An exhaustive analysis of the measured wakes has provided a comprehensive description of the underlying mechanisms in both waketransport phenomena and stator-rotor interaction. In addition, unmixed stator wakes, observed at the rotor exit, have been treated in terms of dispersion and angular displacement to indicate the influence of the blades loading on the transport of the stator wake fluid. The final aim of the paper is to highlight a complete picture of the unsteady flow patterns inside industrial axial fans. [DOI: 10.1115/1.2746920]

A health-monitoring and life-estimation strategy for composite rotor blades is developed in this work. The crosssectional stiffness reduction obtained by physics-based models is expressed as a function of the life of the structure using a recent phenomenological damage model. This stiffness reduction is further used to study the behavior of measurable system parameters such as blade deflections, loads, and strains of a composite rotor blade in static analysis and forward flight. The simulated measurements are obtained using an aeroelastic analysis of the composite rotor blade based on the finite element in space and time with physics-based damage modes that are then linked to the life consumption of the blade. The model-based measurements are contaminated with noise to simulate real data. Genetic fuzzy systems are developed for global online prediction of physical damage and life consumption using displacement- and force-based measurement deviations between damaged and undamaged conditions. Furthermore, local online prediction of physical damage and life consumption is done using strains measured along the blade length. It is observed that the life consumption in the matrix-cracking zone is about 12-15% and life consumption in debonding/delamination zone is about 45-55% of the total life of the blade. It is also observed that the success rate of the genetic fuzzy systems depends upon the number of measurements, type of measurements and training, and the testing noise level. The genetic fuzzy systems work quite well with noisy data and are recommended for online structural health monitoring of composite helicopter rotor blades.

To link to full-text access for this article, visit this link: http://dx.doi.org/10.1016/j.apergo.2005.12.007 Byline: Jacques Marsot (a), Laurent Claudon (b), Marc Jacqmin (a) Keywords: Knife; Ergonomics; Cutting; Force; MSD Abstract: Following a brief description of the problem of musculoskeletal disorders (MSDs) in the meat industry and use of knives, this paper presents a study of the influence of the main intrinsic technical characteristics of a knife on its cutting performance. This study prompted design of a specific system for measuring initial cutting capacity and cutting edge retention. This design process preceded research into test conditions (cutting speed, sample nature and thickness, knife inclination angle, etc.) offering optimum test bench operation. This equipment was then used to study the influence of blade inclination angle, steel grade and sharpening angle on cutting performance. It has been shown that cutting force varies with blade inclination. Use of knives with curved blades and/or a blade inclined with respect to the knife handle is therefore preferred in relation to our cutting force reduction objective. It has also been shown that choices are in fact governed by compromises in relation to other parameters (steel grade and edge angle). These observations confirm the need to set up suitable training of knife users to achieve best possible cutting performance. Author Affiliation: (a) French National Research and Safety Institute (INRS), Work Equipment Engineering Department, B.P. 27-54501 Vandoeuvre-Les-Nancy Cedex, France (b) French National Research and Safety Institute (INRS), Working Life Department, B.P. 27-54501 Vandoeuvre-Les-Nancy Cedex, France Article History: Received 31 January 2005; Accepted 24 December 2005

The secondary, flows, including passage and other vortices in a turbine cascade, cause significant aerodynamic losses and thermal gradients. Leading edge modification of the blade has drawn considerable attention as it has been shown to reduce the secondary flows. However, the heat transfer performance of a leading edge modified blade has not been investigated thoroughly. Since a fillet at the leading edge blade is reported to reduce the aerodynamic loss significantly, the naphthalene sublimation technique with a fillet geometry is used to study local heat (mass) transfer performance in a simulated turbine cascade. The present paper compares Sherwood number distributions on an endwall with a simple blade and a similar blade having a modified leading edge by adding a fillet. With the modified blades, a horseshoe vortex is not observed and the passage vortex is delayed or not observed for different turbulence intensities. However, near the blade trailing edge the passage vortex has gained as much strength as with the simple blade for low turbulence intensity. Near the leading edge on the pressure and the suction surface, higher mass transfer regions are observed with the fillets. Apparently the corner vortices are intensified with the leading edge modified blade.

To link to full-text access for this article, visit this link: http://dx.doi.org/10.1016/j.apergo.2004.08.002 Byline: Raymond W. McGorry (a), Peter C. Dowd (b), Patrick G. Dempsey (a) Keywords: Knife; Sharpening; Honing Abstract: Meat cutting causes a much greater incidence of upper extremity disorders than industry averages. Grip forces and cutting moments in meat cutting operations can be very high. Recently published studies indicate that the blade sharpness can significantly affect force exposure during meat cutting. Two attributes that have not been investigated with respect to force exposure during meat cutting are the effect of blade edge angle and blade finishing after sharpening (honing and polishing). This study investigated the effect of three blade edge angles and three post-sharpening finishing protocols on the grip forces and cutting moments exerted by professionals during two different meatpacking operations. A fine polish finishing protocol significantly reduced cutting time by 25.3%, mean grip force by 21.2%, and mean cutting moment by 28.4% over a coarser finishing protocol during one of the operations. No significant differences in the exposure variables were found between the blade edge angle conditions. Author Affiliation: (a) Liberty Mutual Research Institute for Safety, 71 Frankland Road, Hopkinton, MA 01748, USA (b) Anago Ltd., Hamilton, New Zealand Article History: Received 18 September 2003; Revised 20 June 2004; Accepted 5 August 2004

A quasi-static acoustic mapping method has been developed to predict rotorcraft external noise. This method takes advantage of an expansion, to first order, about a solution to the helicopter nondimensional steady-state trim equations to include the effects of acceleration parallel to the flight path and X-force control changes on the radiated noise. Application of the new method to predict helicopter blade--vortex interaction (BVI) noise has shown that choice of flight-path angle, X-force, and vehicle acceleration all have an important influence on ground noise exposure during approach and landing. The effect of constant flight-path-angle approaches on BVI ground noise, with and without X-force control, is compared with decelerating approaches at the same flight-path angle. Two different quiet flight trajectories are suggested that use a combination of these controls to minimize BVI noise exposure on the ground during a helicopter approach to a landing.

The purpose of this study was to analyze the impact of blades with slots in a rotary compressor. The equations for contact forces were derived from Newton's law. Updated moving Lagrangian descriptions with the hybrid technique and the Newark time integration method were used to determine the nodal displacement vectors. The finite element method was used to analyze the contact forces and the stresses of blades. The maximum blade contact force of 2.3 KN was larger than the maximum static force of 1 KN. The maximum blade impact Von Mises stress of 12.3 Mpa that occurred at the contact point was greater than the previous calculated static stress of 5.15 Mpa. Hence, the impact effect was significant. Experimental modal analysis was also used to determine the first and second natural frequencies of blades.

This paper describes a knowledge-based method for the automatic design of more efficient turbine blades. An Artificial Neural Network (ANN) is used to construct an approximate model (response surface) using a database containing Navier-Stokes solutions for all previous designs. This approximate model is used for the optimization, by means of Simulated Annealing (SA), of the blade geometry, which is then analyzed by a Navier-Stokes solver. This procedure results in a considerable speed-up of the design process by reducing both the interventions of the operator and the computational effort. It is also shown how such a method allows the design of more efficient blades while satisfying both the aerodynamic and mechanical constraints. The method has been applied to different types of two-dimensional turbine blades, of which three examples are presented in this paper.

The practical utility of a three-dimensional inverse viscous method is demonstrated by carrying out a design modification of a first-stage rotor in an industrial compressor. In this design modification study, the goal is to improve the efficiency of the original blade while retaining its overall aerodynamic, structural, and manufacturing characteristics. By employing a simple modification to the blade pressure loading distribution (which is the prescribed flow quantity in this inverse method), the modified blade geometry is predicted to perform better than the original design over a wide range of operating points, including an improvement in choke margin.

This study uses a nonlinear optimization method coupled with a vortex lattice cavitating propeller analysis method to design efficient propeller blades. Different constraints are imposed to improve propeller design. Several advancements in the method are shown, including the option for quadratic skew, user specified skew distribution, and a constraint limiting the minimum pressure in wetted regions of the blade. Results for a series of fully wetted runs demonstrate the effectiveness of the constraint on minimum pressure in preventing the onset of bubble or mid-chord cavitation. A comparison of a design in uniform inflow with a design in non-axisymmetric inflow indicates that a propeller designed by the present method in non-axisymmetric inflow has more favorable cavitating flow characteristics than a propeller design assuming uniform inflow. Results are also shown for a series of runs utilizing the cavity constraints. These results indicate that the present method can be used to improve on propeller designs by imposing constraints on the cavity area and cavity volume velocity harmonics, as well as by using a quadratic skew distribution.

An iterative procedure for three-dimensional blade design is presented, in which the three-dimensional blade shape is modified using a physical algorithm, based on the transpiration model. The transpiration flux is computed by means of a modified Euler solver, in which the target pressure distribution is imposed along the blade surfaces. Only a small number of modifications is needed to obtain the final geometry. The method is based on a high-resolution three-dimensional Euler solver. An upwind biased evaluation of the advective fluxes allows for a very low numerical entropy generation, and sharp shock capturing. Non-reflecting boundary conditions are applied along the inlet/outlet boundaries. The capabilities of the method are illustrated by redesigning a transonic compressor rotor blade, to achieve, for the same mass flow and outlet flow angle, a shock-free deceleration along the suction side. The second example concerns the design of a low aspect ratio turbine blade, with a positive compound lean to reduce the intensity of the passage vortices. The final blade is designed for an optimized pressure distribution, taking into account the forces resulting from the blade lean angle.

The flow field in a preswirled cooling air supply to a turbine rotor has been investigated by means of CFD simulations. Coefficients for system efficiency are derived. The influences of various geometric parameters for different configurations have been correlated with the help of appropriate coefficients. For some of the most important geometric parameters of the coverplate receiver, design recommendations have been made. For the preswirl nozzles, the potential of efficiency improvement by contour design is highlighted.

This paper describes a program of work, largely experimental, which was undertaken with the objective of developing an improved blade profile for the low-pressure turbine in aero-engine applications. Preliminary experiments were conducted using a novel technique. An existing cascade of datum blades was modified to enable the pressure distribution on the suction surface of one of the blades to be altered. Various means, such as shaped inserts, an adjustable flap at the trailing edge, and changing stagger were employed to change the geometry of the passage. These experiments provided boundary layer and lift data for a wide range of suction surface pressure distributions. The data were then used as a guide for the development of new blade profiles. The new blade profiles were then investigated in a low-speed cascade that included a set of moving bars upstream of the cascade of blades to simulate the effect of the incoming wakes from the previous blade row in a multistage turbine environment. Results are presented for two improved profiles that are compared with a datum representative of current practice. The experimental results include loss measurements by wake traverse, surface pressure distributions, and boundary layer measurements. The cascades were operated over a Reynolds number range from O. 7 x [10.sup.5] to 4.0 x [10.sup.5]. The first profile is a 'laminar flow' design that was intended to improve the efficiency at the same loading as the datum. The other is a more highly loaded blade profile intended to permit a reduction in blade numbers. The more highly loaded profile is the most promising candidate for inclusion in future designs. It enables blade numbers to be reduced by 20 percent, without incurring any efficiency penalty. The results also indicate that unsteady effects must be taken into consideration when selecting a blade profile for the low-pressure turbine.

A blade design and optimization procedure is presented in this work, which is based on viscous flow considerations. This procedure concerns the design of optimum rotating arbitrary compressible high subsonic compressor and turbine blade shapes. It takes into account the effects of wall curvature and Coriolis force on turbulence, while it allows the variation of stream surface radius, along which the blade shape is placed, as well as streamtube width, with meridional distance. In order to establish the inverse part of the viscous optimization procedure, aspects such as laminar stability, transition, optimum deceleration and, more generally, the behaviour of compressible attached and separated shear layers are discussed. A plane on which all the general properties of the compressible laminar and turbulent shear layers appear, is constructed and the generation of optimum shear layers for the critical side of the blade shape is established. The complete optimization (design) procedure is then described and discussed, while various designs realized by the present procedure are presented at the end of this paper.

Background information is presented on an arbitrary blade section design method which is outlined in a joint paper. This information concerns the assumptions, the development, and the predictive capabilities of the viscous flow calculation tool used in the design procedure. General properties of laminar and turbulent, unseparated or separated compressible shear layers, necessary for the blade optimization procedure, are discussed.

A time-marching algorithm is proposed for designing turbomachine blades with high aerodynamic efficiency. The algorithm uses a three-dimensional inverse methodology to take into account the effect of the circumferential mass-averaged tangential velocity on the aerodynamic performance of the blade geometry. Its applicability is validated in the design of a compressor blade in the subsonic-flow regime.

A new sensitivity analysis method that employs a chain rule differentiation approach helps compute the sensitivities of hub loads and rotating blade root loads to variations in blade design parameters. The algorithm in the new method utilizes the features of fuselage equations of motion, blade formulations and solution technique to determine the sensitivities. Elucidation of the fuselage by nonlinear Euler equations and inclusion of nonlinearities in the mathematical model for blade yield a model that is suitable for turning and straight flight.

The aerodynamic characteristics of an S-cambered profile are studied under forward and reversed flow conditions. The profile chord is cut by 3, 6, and 9 percent of the chord at the sharp trailing edge end and the performances of these profiles are compared. It is found that with increase in length of cutting the lift coefficient increases in forward direction and decreases in reverse direction of flow. Cutting off the sharp trailing edge improves the lift-drag characteristics in forward mode and deteriorates in the reverse mode.

This paper describes the experimental results of output power augmentation of a horizontal axis wind turbine with a tip vane. In order to find the relationship between the performance of the turbine and the configuration of the tip vane, various types and sizes were used. It was found that V-type and S-type tip vanes can improve turbine performance. Also, the dimensions of V- and S-type tip vanes were investigated. The maximum improvement achieved was a 25 percent increase in power in an existing wind turbine without a tip vane.

A three-dimensional design method for the design of the blade geometry of centrifugal compressor impellers is presented. In this method the blade shape is computed for a specified circulation distribution, normal (or tangential) thickness distribution, and meridional geometry. As the blade shapes are computed by using an inviscid slip (or flow tangency) condition, the viscous effects are introduced indirectly by using a viscous/inviscid procedure. The three-dimensional Navier-Stokes solver developed by Dawes is used as the viscous method. Two different approaches are described for incorporating the viscous effects into the inviscid design method. One method is based on the introduction of an aerodynamic blockage distribution throughout the meridional geometry, while in the other approach a vorticity term directly related to the entropy gradients in the machine is introduced. The method is applied to redesign the blade geometry of Eckardt's 30 deg backswept impeller as well as a generic high pressure ratio (transonic) impeller. The results indicate that the entropy gradient approach can fairly accurately represent the viscous effects in the machine.

An optimization procedure has been developed to address the complex and conflicting requirements associated with the design of high-speed prop-rotor aircraft. Since the key technical issues are maintenance of propulsive efficiency and aeroelastic stability in high-speed cruise without deteriorating figure of merit in hover, rotor aerodynamic performance and aeroelastic analysis are coupled, inside a closed loop, to the optimizer. The discipline couplings provide actual blade air-loads, during hover and cruise, and also provide realistic blade designs. The propulsive efficiency in high-speed cruise is used as the objective function. Constraints are also imposed on the aeroelastic stability in axial flight and the rotor figure of merit in hover. Both structural and planform design variables are used. The optimization procedure yields significant improvements in the aerodynamic characteristics of the rotor. Off-design performance studies, conducted with the optimum blade, show overall design improvements.

The direct and inverse blade-design iterations for the selection of isolated airfoils and gas turbine blade cascades are enormously reduced if the initial blade shape has performance characteristics near the desirable ones. This paper presents the hierarchical development of three direct blade-design methods of increasing utility for generating two-dimensional blade shapes. The methods can be used to generate inputs to the direct- or inverse-blade-design sequences for subsonic or supersonic airfoils for compressors and turbines, or isolated airfoils. The examples included for illustration are typical modern turbine cascades, and they have been designed by the direct method exclusively. The first method specifies the airfoil shapes with analytical polynomials. It shows that continuous curvature and continuous slope of curvature are necessary conditions to minimize the possibility of flow separation, and to lead to improved blade designs. The second method specifies the airfoil shapes with parametric fourth-order polynomials, which result in continuous-slope-of-curvature airfoils, with smooth Mach number and pressure distributions. This method is time consuming. The third method specifies the airfoil shapes by using a mixture of analytical polynomials and mapping the airfoil surfaces from a desirable curvature distribution. The third method provides blade surfaces with desirable performance in very few direct-design iterations. In all methods the geometry near the leading edge is specified by a thickness distribution added to a construction line, which eliminates the leading edge overspeed and laminar-separation regions. The blade-design methods presented in this paper can be used to improve the aerodynamic and heat transfer performance of turbomachinery cascades, and they can result in high-performance airfoils in very few iterations.

Blade surfaces with continuous curvature and continuous slope of curvature minimize the possibility of flow separation, lead to improved blade designs, and reduce the direct and inverse blade-design iterations for the selection of isolated airfoils and gas-turbine-blade cascades. A method for generating two-dimensional blade shapes is presented. The geometry near the trailing edge is specified by an analytic polynomial, the main portion of the blade surface is mapped using as input a prescribed surface-curvature distribution, and the leading edge is specified as a thickness distribution added to a construction line. This procedure is similar for the suction and pressure surfaces, and by specification it constructs continuous slope-of-curvature surfaces that result in smooth surface-Mach-number and surface-pressure distributions. The method can be used to generate subsonic or supersonic airfoils for compressors and turbines, or isolated airfoils. The resulting geometric shapes can be used as inputs to various blade-design sequences. It is shown that, with other cascade-design parameters being equal, increasing the stagger angle of turbine blades results in more front-loaded and thinner blades, and that there is an optimum stagger angle resulting in minimum wake thickness. The subsonic axial-turbine blade rows included for discussion in this paper have been designed by iterative modifications of the blade geometry to obtain a desirable velocity distribution. The blade-design method can be used to improve the aerodynamic and heat transfer performance of turbine cascades, and it can result in high-performance airfoils, even if using the direct method exclusively, in very few iterations.

This paper presents a procedure to extend a recently developed fully three-dimensional inverse method for highly loaded turbomachine blades into the transonic-flow regime. In this inverse method, the required three-dimensional blade profile to produce a prescribed swirl schedule is determined iteratively using the blade boundary conditions. In the present implementation, the flow is assumed to be inviscid and the blades are assumed to be infinitely thin. The relevant equations are solved in the conservative forms and are discretized in all three directions using a finite-volume technique. Calculations are carried out for the design of high-pressure axial- and centrifugal-compressor rotors. These examples include prescribed swirl schedules corresponding to blade designs that are shock-free and blade designs that have rapid compression regions in the blade passage.

This paper describes the effect of tip clearance changes on the pressure at the case wall of a second-stage rotor. Wall shear distributions under the rotor tip are also presented. The results show low-pressure areas extending along the rotor suction side but lying away from the blade. Pressure contours indicate the tangential loading at the tip is lower than predicted by two-dimensional calculations; however, the predicted loading is observed between the lowest pressure's path in the passage and the blade pressure side. The results suggest that a viscous or shearing layer, due to blade-to-wall relative motion, is generated on the blade side of the tip gap, which modifies the inviscid relative flow field and produces an unloading on the blade tip.

The paper presents further results from a continuing study on tip leakage in axial turbines. Rotation has been simulated in a linear cascade test section by using a moving-belt tip wall. Measurements were made inside the tip gap with a three-hole pressure probe for a clearance size of 3.8 percent of the blade chord. Two wall speeds are considered and the results are compared with the case of no rotation. As in other experiments, significant reduction in the gap mass flow rate is observed due to the relative motion. The detailed nature of the measurements allows the dominant physical mechanism by which wall motion affects the tip gap flow to be identified. Based on the experimental observations, an earlier model for predicting the tip gap flow field is extended to the case of relative wall motion. Part II of the paper examines the effect of the relative motion on the downstream flow field and the blade loading.

The effect of difference in rotor tip clearance on the mean flow fields and unsteadiness and mixing across a stator blade row were investigated using hot-wire anemometry, pressure probes, flow visualization, and the ethylene tracer-gas technique on a single-stage axial flow compressor. The structure of the three-dimensional flow fields was discussed based on results of experiments using the 12-orientation single slanted hot-wire technique and spectrum analysis of velocity fluctuation. High-pass filtered measurements of turbulence were also carried out in order to confirm small-scale velocity fluctuation, which is more realistically referred to as turbulence. The spanwise distribution of ethylene gas spreading, estimated by the measured small-scale velocity fluctuation at the rotor exit, agreed quite well with that which was experimentally measured. This fact suggests the significant role of turbulence, generated within the rotor, in the mixing process across the downstream stator. The value of the maximum mixing coefficient in the tip region was found to increase linearly as the tip clearance became enlarged, starting from the value at midspan.

This paper and its companion paper present experimental results on the effects of simulated rotation on the tip leakage in a linear turbine cascade test. Part II examines the downstream flow field. For clearance sizes of 2.4 and 3.8 percent of the blade chord, measurements were made in two planes downstream of the trailing edge using a seven-hole pressure probe. Significant changes in the tip leakage vortex and passage vortex structures are observed with the introduction of relative motion. The effects of clearance size and rotation on the relationship between bound circulation and tip-vortex circulation are discussed. The validity of a previously developed tip-vortex model for the case of rotation is examined in the light of the measurements. Finally, for clearances of 1.5, 2.4, and 3.8 percent of the blade chord, the effects of rotation on blade loading are studied through static pressure measurements on the blade surfaces. The distortion of the surface pressure field near the tip is found to be reduced with increasing wall speed. This is consistent with the reduced strength of the tip-leakage vortex as wall speed is increased. For all measurements two wall speeds are considered and the results are compared with the case of no rotation.

The phenomenon of tip leakage has been studied in two linear cascades of turbine blades. The investigation includes an examination of the performance of the cascades with a variety of tip geometries. The effects of using plain tips, suction side squealers, and pressure side squealers are reported. Traverses of the exit flow field were made in order to determine the overall performance. A method of calculating the tip discharge coefficients for squealer geometries is put forward. In linking the tip discharge coefficient and cascade losses, a procedure for predicting the relative performance of tip geometries is developed. The model is used to examine the results obtained using the different tip treatments and to highlight the important aspects of the loss generation process.

The paper describes a method of computing nonequilibrium, steady flows of wet steam in two- and quasi-three-dimensional turbine cascades. The mixture conservation equations are solved in an Eulerian reference frame using an inviscid time-marching method that includes the effects of the centrifugal and Coriolis acceleration terms in rotating blade rows. Nucleation and growth of water droplets are computed by integrating the relevant equations along true streamlines in a Lagrangian reference frame. Steam properties are computed using equations that display commercial steam table accuracy for pressures below 10 bar. Special procedures for grouping the range of droplet sizes present are described that allow an accurate representation of the droplet size distribution to be retained without requiring a large increase in CPU time. All types of wet-steam flow, including those involving secondary nucleation, can be computed. Examples are presented that display the sensitivity of the calculation procedure in computing nucleation affected by the shock- and expansion-wave structure in the region of a turbine blade trailing edge. Typical CPU time requirements for nonequilibrium solutions involving primary or secondary nucleations are about three times those for perfect gas calculations.

A numerical method for solving the three-dimensional aerothermodynamic design problem with some type of Mach number distributions on the blade surfaces is presented. In the usual aerothermodynamic design of a turbomachine, the three-dimensional coordinates of the blade are attained through stacking of the cascade profiles and may not ensure the desired velocity distribution. To avoid this problem, the present method will give new coordinates of the blade according to the required Mach number distribution. The method is based on the pseudostream function formulation and the treatment of the boundary conditions in the design problem is given. The numerical results show that the method is simple and useful in design.

Improving the performance of high-speed axial compressors through low-speed model compressor testing has proved to be economical and effective (Wisler, 1985). The key to this technique is to design low-speed blade profiles that are aerodynamically similar to their high-speed counterparts. The conventional aerodynamic similarity transformation involves the small disturbance potential flow assumption; therefore, its application is severely limited and generally not used in practical design. In this paper, a set of higher order transformation rules are presented that can accommodate large disturbances at transonic speed and are therefore applicable to similar transformations between the high-speed high-pressure compressor and its low-speed model. Local linearization is used in the nonlinear equations and the transformation is obtained in an iterative process. The transformation gives the global blading parameters such as camber, incidence, and solidity as well as the blade profile. Both numerical and experimental validations of the transformation show that the nonlinear similarity transformations do retain satisfactory accuracy for highly loaded blades up to low transonic speeds. Further improvement can be made by only slightly modifying profiles numerically without altering the global similarity parameters.

A iteratine procedure for blade design, using a time marching procedure to solve the unsteady Euler equations in the blade-to-blade plane, is presented. A flow solver, which performs the analysis of the flow field for a given geometry, is transformed into a design method. This is done by replacing the classical slip condition (no normal velocity component) by other boundary conditions, in such a way that the required pressure or Mach number distribution may be imposed directly on the blade. The unknowns are calculated on the blade wall using the so-called compatibility relations. Since the blade shape is not compatible with the required pressure distribution, a nonzero velocity component normal to the blade wall evolves from the new flow calculation. The blade geometry is then modified by resetting the wall parallel to the new flow field, using a transpiration technique, and the procedure is repeated until the calculated pressure distribution has converged to the required one. Examples for both subsonic and transonic flows are presented and show a rapid convergence to the geometry required for the desired Mach number distribution. An important advantage of the present method is the possibility to use the same code for the design and the analysis of a blade.

An analysis of the static response of a rotor blade with extension-twist coupling as a result of variations in rpm, layer angles and composite material laminations. The blade is mathematically simulated as a D-shaped spar using discrete Kirchoff theory and finite elements. Results are given in the form of carpet plots and the coupling is calculated through the angular motion of the blade tip. A new technique is also given to obtain six rotor blade arrangements made of three laminated shells.

Combustion turbine blade design criteria can generally be classified as either temperature or fatigue related. Since less is usually known about the factors influencing the fatigue phenomenon, it is considered the more challenging. In addition, as analytical and experimental techniques became more sophisticated and more accurate, the natural tendency was to replace archaic guidelines' or rules' with less conservative approaches that at times led to the discovery of new high-cycle fatigue 'thresholds.' This paper presents the evolution of the combustion turbine blade high cycle fatigue design criteria for free-standing blades. It also presents the analysis and corrective actions taken to resolve several unique combustion turbine blade fatigue problems, all encountered over a 35-year period while the author has been employed at Westinghouse Electric Corporation. Included are high-cycle fatigue problems due to cooling air leakage, seal pin friction, and combustion temperature maldistribution, as well as flow-induced nonsynchronous vibration., The evolution of fatigue design criteria for turbine blades is presented. It was shown that the development of different design criteria began during the 1950s for steam turbines and jet engines. These design criteria led to the development of the combustion turbine blade high-cycle fatigue criteria. As techniques developed, changes were made in design criteria to improve accuracy. Various failure modes were also presented. These include cooling air leakage, seal pin effects, combustion system effects and flow-induced nonsynchronous vibration.

For widely used multistage centrifugal pumps, their former structures are so bulky that nowadays growing interest has been shifted to the development of more compact structures. Following this trend, a compact pump structure is provided and analysed. To maintain the pump's pressure recovery, as well as to meet the water flow from the impeller, a circumferential twisted return guide vane (RGV) is proposed. To validate this design method, the instantaneous CFD simulations are performed to investigate the rotor-stator interventions. Within the impeller, the pressure fluctuation is cyclic symmetry, where the impeller frequency dominates. At the zone where flow leaves impeller for RGV, the pressure fluctuation is nonperiodic, the impeller frequency is major, and the rotation frequency is secondary. Within RGV, the periodic symmetric fluctuation is recovered, where the rotation frequency is governing. The fluctuation decreases from seven cycles within impeller to two cycles within RGV, indicating that the flow from impeller is well handled by RGV. To examine the pump's performance, a prototype multistage pump is designed. The testing shows that the pump efficiency is 57.5%, and the stage head is 9 m, which is comparable to former multistage centrifugal pumps. And this design is more advantageous in developing compact multistage centrifugal pumps., 1. Introduction To meet variable flow angles into gas turbines, hydraulic runners, and so forth, inlet guide vane (IGV) is adopted to keep them operating at peak performance [1-5]. For [...]

A constraint that restricts the sensitivity of a design to mistuning, small differences in blade properties of engine rotors, is developed. The constraint is dependent on the properties of only the tuned system, it requires no knowledge of the actual mistuning pattern and the mistuning statistics can be estimated for a population of manufactured rotors. The constraint is suitable for computation-intensive optimization of engine.

This is the second part of a two-part paper. First, the design-optimization system based on the adjoint gradient solution approach as described in Part I is introduced. Several test cases are studied for further validation and demonstration of the methodology and implementation. The base-line adjoint method as applied to realistic 3D configurations is demonstrated in the redesign of the NASA rotor 67 at a near-choke condition, leading to a 1.77% efficiency gain. The proposed adjoint mixing plane is applied to the redesign of a transonic compressor stage (DLR compressor stage) and an IGV-rotor-stator configuration of a Siemens industrial compressor at a single-operating point, both producing measurably positive efficiency gains. An examination on the choice of the operating mass flow condition as the basis for the performance optimization, however, highlights' the limitation of the single-point approach for practical applications. For the three-row compressor configuration, a near peak-efficiency point based redesign leads to a measurable reduction in the choke mass flow, while a near-choke point based redesign leads to a significant performance drop in other flow conditions. Subsequently, a parallel multipoint approach is implemented. The results show that a two-point design optimization can produce a consistently better performance over a whole range of mass flow conditions compared with the original design. In the final case, the effectiveness of the present method and system is demonstrated by a redesign applied to a seven-row industrial compressor at the design point, leading to a remarkable 2.4% efficiency gain. [DOI: 10.1115/1.3103928]

The adjoint method for blade design optimization will be described in this two-part paper. The main objective is to develop the capability of carrying out aerodynamic blading shape design optimization in a multistage turbomachinery environment. To this end, an adjoint mixing-plane treatment has been proposed. In the first part, the numerical elements pertinent to the present approach will be described. Attention is paid to the exactly opposite propagation of the adjoint characteristics against the physical flow characteristics, providing a simple and consistent guidance in the adjoint method development and applications. The adjoint mixing-plane treatment is formulated to have the two fundamental features of its counterpart in the physical flow domain: conservation and nonreflectiveness across the interface. The adjoint solver is verified by comparing gradient results with a direct finite difference method and through a 2D inverse design. The adjoint mixing-plane treatment is verified by comparing gradient results against those by the finite difference method for a 2D compressor stage. The redesign of the 2D compressor stage further demonstrates the validity of the adjoint mixing-plane treatment and the benefit of using it in a multi-bladerow environment. [DOI: 10.1115/1.3072498]