Flows In Blade Rows Research Articles

Effects of Transient Heat Transfer on Compressor Stability

The effects of heat transfer between the compressor structure and the primary gas path flow on compressor stability are investigated during hot engine re-acceleration transients. A mean line analysis of an advanced, high-pressure ratio compressor is extended to include the effects of heat transfer on both stage matching and blade row flow angle deviation. A lumped capacitance model is used to compute the heat transfer of the compressor blades, hub, and casing to the primary gas path. The inputs to the compressor model with heat transfer are based on a combination of full engine data, compressor test rig measurements, and detailed heat transfer computations. Nonadiabatic transient calculations show a 8.0 point reduction in stall margin from the adiabatic case, with heat transfer predominantly altering the transient stall line. 3.4 points of the total stall margin reduction are attributed to the effect of heat transfer on blade row deviation, with the remainder attributed to stage rematching. Heat transfer increases loading in the front stages and destabilizes the front block. Sensitivity studies show a strong dependence of stall margin to heat transfer magnitude and flow angle deviation at low speed, due to the effects of compressibility. Computations for the same transient using current cycle models with bulk heat transfer effects only capture 1.2 points of the 8.0 point stall margin reduction. Based on this new capability, opportunities exist early in the design process to address potential stability issues due to transient heat transfer.

Transient flow in a compressor blade row for a periodic vibration motion

The goal of this work is to conduct a transient compressor blade row flow simulation as part of blade flutter modeling. An integral step of blade flutter modeling is the calculation of the aerodynamic damping factor as a function of the possible vibration mode shapes. Using Fourier method, the number of blade passages required for transient flow analysis is kept to a minimum of two for all vibration modes. In this work, a compressor rotor blade row is considered. The vibration modes are obtained using ANSYS mechanical, then, unsteady flow is obtained for vibrating blades with a harmonic motion. Work of the flow on the blade is calculated and hence the aerodynamic damping is obtained.

J055043 航空エンジン用低圧タービン翼の高効率化に関する研究 : Wake流入条件下における二次流れ損失の調査(〔J055-04〕流体機械の研究開発におけるEFD/CFD(4))

This paper details experimental studies on the secondary flow losses in low-pressure turbine (LPT) cascades. A large portion of overall loss in LPT is due to secondary flow loss. Hence it is important to understand the mechanisms of secondary flow loss production. Moreover, strong wake generated by the upstream blade rows exists in a turbine. This causes aerodynamic interaction of the secondary flow and boundary layer separation in the downstream blade rows, resulting in a complex flow field in the secondary flow region. The objective of this study is to investigate the detail of secondary flow losses under the effect of wake-passing.

A New Streamline Curvature Throughflow Method for Radial Turbomachinery

This paper describes a newly developed streamline curvature throughflow method for the analysis of radial or mixed flow machines. The code includes curved walls, curved leading and trailing edges, and internal blade row calculating stations. A general method of specifying the empirical data provides separate treatment of blockage, losses, and deviation. Incompressible and compressible fluids are allowed, including real gases and supersonic relative flow in blade rows. The paper describes some new aspects of the code. In particular, a relatively simple numerical model for spanwise mixing is derived; the calculation method for prescribed pressure ratio in compressor blade rows is described; and the method used to redistribute the flow across the span due to choking is given. Examples are given of the use and validation of the code for many types of radial turbomachinery, and these show that it is an excellent tool for preliminary design.

Simulation of 3D flow in turbine blade rows including the effects of coolant ejection

This paper describes the numerical simulation of three-dimensional viscous flows in air-cooled turbine blade rows with the effects of coolant ejection. A TVD Navier-Stokes flow solver incorporated with Baldwin-Lomax turbulence model and multi-grid convergence acceleration algorithm are used for the simulation. The influences of coolant ejection on the main flow are accounted by volumetric coolant source terms. Numerical results for a four-stage turbine are presented and discussed.

Single-passage solution of three-dimensional unsteady flows in a transonic fan rotor

Calculations of unsteady turbomachinery flows using conventional time domain methods typically would lead to the use of multiple passages/whole annulus owing to the required direct periodic condition. This makes conventional unsteady solutions extremely time consuming and is one of the major difficulties in applying unsteady calculations in a blading design environment. A single-passage approach to three-dimensional unsteady Navier-Stokes calculations using the Fourier series based shape correction method has been developed. The shape correction method gives a single-passage solution to unsteady flows in blade rows under multiple disturbances with arbitrary interblade phase angles. Therefore the computing time can be dramatically reduced. A case study for a transonic fan rotor subject to long-scale circumferential inlet distortions shows that a factor of 5 speed-up could be achieved. Detailed comparisons in both unsteady and time-averaged flow results between the single-passage and multiple-passage solutions demonstrate the validity and effectiveness of the present single-passage approach. A further calculation of a transonic rotor simultaneously under the influences of both upstream and downstream blade rows also shows the capability of the shape correction method in dealing with multiple disturbances.

On the Development of Viscous Solvers for Computationof Transient Flows in Turbomachines

This paper focuses on development and validation of a viscous solver for the computation of unsteady flows in turbomachinery blade rows and stages consisting of rotors and stators. The code has been evolved from steady‐state single flow solvers developed by Wiedermann based on time‐marching finite difference schemes. A two‐equation eddy viscosity model is applied, and the wall boundary conditions are determined by the y+‐distance of the first grid line away from the wall. For the solution of transient flow fields the original time‐stepping algorithm is replaced by a time‐accurate scheme.The emphasis of the code validation will be on blade‐row interaction in a complete turbomachinery stage. A 3‐D example will be discussed and those parameters evaluated which are important for actual blading design.

Improved Blade Profile Loss and Deviation Angle Models for Advanced Transonic Compressor Bladings: Part II—A Model for Supersonic Flow

New blading concepts as used in modern transonic axial-flow compressors require improved loss and deviation angle correlations. The new model presented in this paper incorporates several elements and treats blade-row flows having subsonic and supersonic inlet conditions separately. The second part of the present report focuses on the extension of a well-known correlation for cascade losses at supersonic inlet flows. It was originally established for DCA bladings and is now modified to reflect the flow situation in blade rows having low-cambered, arbitrarily designed blades including precompression blades. Finally, the steady loss increase from subsonic to supersonic inlet-flow velocities demonstrates the matched performance of the different correlations of the new model.

Improved Blade Profile Loss and Deviation Angle Models for Advanced Transonic Compressor Bladings: Part I—A Model for Subsonic Flow

New blading concepts as used in modern transonic axial-flow compressors require improved loss and deviation angle correlations. The new model presented in this paper incorporates several elements and treats blade-row flows having subsonic and supersonic inlet conditions separately. In the first part of this paper two proved and well-established profile loss correlations for subsonic flows are extended to quasi-two-dimensional conditions and to custom-tailored blade designs. Instead of a deviation angle correlation, a simple method based on singularities is utilized. The comparison between the new model and a recently published model demonstrates the improved accuracy in prediction of cascade performance achieved by the new model.

Numerical simulations of flows in centrifugal turbomachinery

A quasi-three-di mensional Navier-Stokes analysis has been extended and applied to flows in centrifugal turbomachinery blade rows. The numerical procedure was validated using a radial fan geometry typical of those found in fossil fuels plants and air-handling systems. The predicted blade loading and performance characteristics showed excellent agreement with the experimental data. The numerical analysis was then applied to a low AR centrifugal impeller. Inviscid and viscous flow simulations were performed at three operating conditions. While steady solutions were obtained in the inviscid flow calculations, the appearance of an oscillating separation bubble on the pressure surface of the impeller necessitated that the viscous flow simulations be time dependent. By comparing the predicted and experimental circumferential distributions of the relative frame velocity and flow angle downstream of the impeller, it was hypothesized that in the experiments the end-wall secondary flows energized the impeller suction surface boundary layer, making the flow locally behave like an inviscid fluid. The performance curve generated from the viscous calculations showed satisfactory agreement with the experimental data, while the inviscid calculations overpredicted the performance of the impeller. It was concluded that the physics retained in quasi-three-d imensional analysis give a useful first approximation of the flow trends in certain types of centrifugal turbomachinery.

Viscous Analysis of Three-Dimensional Rotor Flow Using a Multigrid Method

A three-dimensional code for rotating blade-row flow analysis has been developed. The space discretization uses a cell-centered scheme with eigenvalue scaling for the artificial dissipation. The computational efficiency of a four-stage Runge–Kutta scheme is enhanced by using variable coefficients, implicit residual smoothing, and a full-multigrid method. An application is presented for the NASA rotor 67 transonic fan. Due to the blade stagger and twist, a zonal, nonperiodic H-type grid is used to minimize the mesh skewness. The calculation is validated by comparing it with experiments in the range from the maximum flow rate to a near-stall condition. A detailed study of the flow structure near peak efficiency and near stall is presented by means of pressure distribution and particle traces inside boundary layers.

Deforming grid variational principle for unsteady small disturbance flows in cascades

A variational method for computing unsteady subsonic flows in turbomachinery blade rows is presented. A variational principle that describes the harmonic small disturbance behavior of the full potential equations about a nonlinear mean flow is developed. Included in this variational principle is the effect of a deforming computational grid that conforms to the motion of vibrating airfoils. Bilinear isoparametric finite elements are used to discretize the variational principle, and the resulting discretized equations are solved efficiently using lower-upper decomposition. The use of a deforming computational grid dramatically improves the accuracy of the method since no error-producing extrapolation is required to apply the upwash boundary conditions or to evaluate the unsteady pressure on the airfoil surfaces

Linearized Euler predictions of unsteady aerodynamic loads in cascades

A linearized Euler solver for calculating unsteady flows in turbomachinery blade rows due to both incident gusts and vibratory blade motion is presented. Using the linearized Euler technique, one decomposes the flow into a mean (or steady) flow plus an unsteady, harmonically varying, small-disturbance flow. Linear variable coefficient equations describe the small-disturbance behavior of the flow and are solved using a pseudotime time-marching Lax-Wendroff scheme. For the blade-motion problem, a harmonically deforming computational grid that conforms to the motion of vibrating blades eliminates large error producing mean flow gradient terms that would otherwise appear in the unsteady flow tangency boundary condition

Comparison of Time-Resolved Turbine Rotor Blade Heat Transfer Measurements and Numerical Calculations

Time-resolved turbine rotor blade heat transfer data are compared with ab initio numerical calculations. The data were taken on a transonic, 4-to-1 pressure ratio, uncooled, single-stage turbine in a short-duration turbine test facility. The data consist of the time history of the heat transfer distribution about the rotor chord at midspan. The numerical calculation is a time accurate, two-dimensional, thin shear layer, multiblade row code known as UNSFLO. UNSFLO uses Ni’s Lax-Wendroff algorithm, conservative boundary conditions, and a time tilting algorithm to facilitate the calculation of the flow in multiple blade rows of arbitrary pitch ratio with relatively little computer time. The version used for this work had a simple algebraic Baldwin-Lomax turbulence model. The code is shown to do a good job of predicting the quantitative time history of the heat flux distribution. The wake/boundary layer and transonic interaction regions for suction and pressure surfaces are identified and the shortcomings of the current algebraic turbulence modeling in the code are discussed. The influence of hardware manufacturing tolerance on rotor heat transfer variation is discussed. A physical reasoning explaining the discrepancies between the unsteady measurement and the calculations for both the suction and pressure surfaces are given, which may be of use in improving future calculations and design procedures.

The Simulation of Three-Dimensional Viscous Flow in Turbomachinery Geometries Using a Solution-Adaptive Unstructured Mesh Methodology

This paper presents a numerical method for the simulation of flow in turbomachinery blade rows using a solution-adaptive mesh methodology. The fully three-dimensional, compressible, Reynolds-averaged Navier–Stokes equations with k–ε turbulence modeling (and low Reynolds number damping terms) are solved on an unstructured mesh formed from tetrahedral finite volumes. At stages in the solution, mesh refinement is carried out based on flagging cell faces with either a fractional variation of a chosen variable (like Mach number) greater than a given threshold or with a mean value of the chosen variable within a given range. Several solutions are presented, including that for the highly three-dimensional flow associated with the corner stall and secondary flow in a transonic compressor cascade, to demonstrate the potential of the new method.

The unresolved unsteady flow in multistage compressor blade rows

A series of experiments are described that are directed at the quantitative investigation of turbomachine rotor blade row unsteady unresolved velocities and blade surface pressures generated by the wakes from upstream blade and vane airfoil rows. The ensemble-averaged velocities and pressures quantify the vane wake generated periodic and aperiodic unsteadiness, with the unresolved unsteadiness defined by the variance. The vane wake generated aperiodic, periodic, and unresolved axial and tangential unsteady velocities and correlations are quantified at the inlet to each rotor row of a three-stage axial flow research compressor by means of a rotating cross hot-wire probe. The rotor blade surface unresolved unsteady pressure—time distributions and time-averaged values are also quantified for each rotor row, accomplished by means of ultraminiature high-response pressure tranducers.

Solution of the three-dimensional turbomachinery blade row flow field problem with allowance for viscosity effects

The three-dimensional viscous compressible flow in the interblade channel of a turbomachine blade row is examined. A mathematical model of steadystate separated turbulent polytropic-gas flow with allowance for wall flow effects is proposed. It is assumed that the thickness of the boundary layers and the scales of the wall flow interaction zones are small as compared with the characteristic dimension of the system. Subsonic flow calculations are made for near-design flow regimes.

3-D Axial Blade Row Flow Field Inviscid Finite Volume Prediction and Verification

Article 3-D Axial Blade Row Flow Field Inviscid Finite Volume Prediction and Verification was published on June 1, 1991 in the journal International Journal of Turbo & Jet-Engines (volume 8, issue 1-2).

LDA Measurement of the Passage Flow Field in an Annular Airfoil Cascade

Models to predict the complex three-dimensional flow through turbomachine blade rows are being developed. To verify these models and direct necessary refinements, it is necessary that predictions be correlated with data obtained in experiments which model the fundamental three-dimensional blade row flow phenomena. This paper describes a series of experiments performed in a large scale, subsonic, annular cascade facility specifically designed to provide such data. In particular, the effect of incidence angle on the three-dimensional passage flow field through an annular cascade of cambered airfoils is investigated and quantified, accomplished by obtaining detailed and expensive LDA data. These data demonstrate and quantify the development of the passage vortices through the airfoil passage and their strong interaction with the endwall boundary layers.

Calculation of Three-Dimensional, Inviscid, Rotational Flow in Axial Turbine Blade Rows

This paper describes a fully explicit, time marching, corrected viscosity, finite volume method to solve the Euler equations in a cylindrical coordinate system. The rotational character of the incoming flow can be taken into account. On the outflow boundary, a generalized radial equilibrium condition is imposed. Blade rows of complex geometry can be handled. At present, the method has been used to calculate the flow through the nozzle vanes of the VKI low-speed turbine facility. The calculated results show good agreement with the experimental data for the spatial distribution of both the static pressure and the flow angle.