Transonic Compressor Blade Research Articles

The effect of Reynolds number on transonic compressor blade rotor section

In this paper, the effect of Reynolds number on transonic compressor blade rotor section is investigated. After passing through the first transonic compressor stages , the flow becomes remarkably compressed. In the present work, it is intended to numerically investigate the effects of the inflow Reynolds number on the unique incidence, flow losses, deviation angle, and shock position, at three different important points of “Minimum Loss” and “Choked Flow” in started conditions and “Stall Operation” in un-started conditions.

Shock wave boundary layer interaction on suction side of compressor profile in single passage test section

The shock wave boundary layer interaction on the suction side of transonic compressor blade is one of the main objectives of TFAST project (Transition Location Effect on Shock Wave Boundary Layer Interaction). In order to investigate the flow structure on the suction side of a profile, a design of a generic test section in linear transonic wind tunnel was proposed. The experimental and numerical results for the flow structure investigations are shown for the flow conditions as the existing ones on the suction side of the compressor profile. Near the sidewalls the suction slots are applied for the corner flow structure control. It allows to control the Axial Velocity Density Ratio (AVDR), important parameter for compressor cascade investigations. Numerical results for Explicit Algebraic Reynolds Stress Model with transition modeling are compared with oil flow visualization, schlieren and Pressure Sensitive Paint. Boundary layer transition location is detected by Temperature Sensitive Paint.

Investigations of shock wave boundary layer interaction on suction side of compressor profile

The shock wave boundary layer interaction on the suction side of transonic compressor blade is one of main objectives of TFAST project (Transition Location Effect on Shock Wave Boundary Layer Interaction). In order to look more closely into the flow structure on the suction side of a blade, a design of a generic test section in linear transonic wind tunnel was proposed. The test section which could reproduce flow structure, shock wave location, pressure distribution and boundary layer development similar to the obtained on a cascade profile is the main objective of the presented here design. The design of the proposed test section is very challenging, because of shock wave existence, its interaction with boundary layer and its influence on the 3-D flow structure in the test section.

Multipoint Design Optimization of a Transonic Compressor Blade by Using an Adjoint Method

This paper presents the application of a viscous adjoint method to the multipoint design optimization of a rotor blade through blade profiling. The adjoint method requires about twice the computational effort of the flow solution to obtain the complete gradient information at each operating condition, regardless of the number of design parameters. NASA Rotor 67 is redesigned through blade profiling. A single point design optimization is first performed to verify the effectiveness and feasibility of the optimization method. Then in order to improve the performance for a wide range of operating conditions, the blade is redesigned at three operating conditions: near peak efficiency, near stall, and near choke. Entropy production through the blade row combined with the constraints of mass flow rate and total pressure ratio is used as the objective function. The design results are presented in detail and the effects of blade profiling on performance improvement and shock/tip-leakage interaction are examined.

Study of Shock Wave Control by Suction & Blowing on a Highly-loaded Transonic Compressor Cascade

A numerical investigation of shock wave control over a highly-loaded transonic compressor blade is presented in the current work. Two types of active flow control methods, namely blowing and suction, are employed. Their effectiveness in controlling the shock wave boundary layer interaction in the intended cascade passage has been investigated numerically. The three-dimensional RANS equations in Cartesian coordinate system are solved using the cell-centred control volume approach. Blowing and suction at four locations along the blade are simulated. Aerodynamic performance is evaluated for each case in terms of total pressure recovery coefficient, deviation angle, etc. Results show that blowing is not capable of providing desired control outcome, as the effects can only be seen in the neighbourhood of the blowing slot. The reconfiguration of shock waves in cascade is realized by suction through the slots located downstream after shock waves. Both the right branch of leading edge shock and the passage shock wave move downstream toward the suction slot. Compared with the baseline case, the suction with a suction ratio of 1.2% from slot D (at the 51% Chord) increased the total pressure recovery coefficient by 10%, and decreased the deviation angle by 5°.

Prediction of Turbomachinery Aeroelastic Behavior From a Set of Representative Modes

A method is proposed for the determination of the aeroelastic behavior of a system responding to mode-shapes which are different from the tuned in vacuo ones, due to mistuning, mode family interaction, or any other source of mode-shape perturbation. The method is based on the generation of a data base of unsteady aerodynamic forces arising from the motion of arbitrary modes and uses least square approximations for the prediction of any responding mode. The use of a reduced order technique allows for mistuning analyses and is also applied for the selection of a limited number of arbitrary modes. The application of this method on a transonic compressor blade shows that the method captures the aeroelastic properties well in a wide frequency range. A discussion of the influence of the mode-shapes and frequency on the final stability response is also provided.

A New Reduced Order Modeling for Stability and Forced Response Analysis of Aero-Coupled Blades Considering Various Mode Families

This paper presents the description and application of a new method for stability and forced response analyses of aerodynamically coupled blades considering the interaction of various mode families. The method, here referred as multimode least square, considers the unsteady forces due to the blade motion at different modes shape families and calculates the aerodynamic matrixes by means of a least square (L2) approximations. This approach permits the prediction of mode families’ interaction with capabilities of structural, aerodynamic and force mistuning. A projection technique is implemented in order to reduce the computational domain. Application of the method on tuned and structural mistuned forced response and stability analyses is presented on a highly loaded transonic compressor blade. When considering structural mistuning the forced response amplitude magnification is highly affected by the change in aerodynamic damping due to mistuning. Analyses of structural mistuning without aerodynamic coupling might result in over-estimated or under-estimated response when the source of damping is mainly aerodynamic. The frequency split due to mistuning can cause that mode families’ interact due to reducing their frequencies separation. The advantage of the present method is that the effect of mode family interaction on aerodynamic damping and forced response is captured not being restricted to single mode families.

An improvement on the efficiency of a single rotor transonic compressor by reducing the shock wave strength on the blade suction surfaces

It is well known that increasing the rotational velocity is an effective way to increase the total pressure ratio. With increasing flow velocity especially under the condition of transonic flow in the supersonic region, where exist strong shock waves, the shock wave loss becomes main and important. Simultaneously, there occurs boundary layer separation due to the shock wave / boundary layer interaction. In the present paper the transonic compressor blades were studied and analyzed to find a proper and simple way to reduce the shock wave loss by optimizing the suction surface configuration or controlling the gradient of isentropic Mach number on the suction surface. A Navier-Stokes solver combined with a modified design algorithm was developed and used. The NASA single rotor for transonic flow compressor was served as a numerical example to show the effectiveness of this method. Two cases for both original and modified rotors were analyzed and compared.

가중평균대리모델을 사용한 천음속 압축기 블레이드 최적화

The main purpose of the present study is to perform shape optimizations of transonic compressor blade in order to enhance its performance. In this study, the Latin hypercube sampling of design of experiments and the weighted average surrogate model with the help of a gradient based optimization algorithm are used within design space by the lower and upper limits of each design variable and for finding optimum designs, respectively. 3-D Reynolds-averaged Navier-Stokes solver is used to evaluate the objective functions of adiabatic efficiency and pressure ratio. Six variables from lean and airfoil thickness profile are selected as design variables. The results show that the adiabatic efficiency is nhanced by 1.43% by efficiency optimization while the pressure ratio is increased very small, and <BR>pressure ratio is increased by 0.24% by pressure ratio optimization.

Aerodynamic Inverse Design for Viscous Flow in Turbomachinery Blading

An inverse design method for turbomachinery blading based on a time-accurate solution of the compressible viscous flow equations on a time-varying geometry is presented. The blade pressure loading and thickness distributions are the prescribed design parameters. The blade profile is modified as it moves at a virtual velocity distribution that would make the momentum flux on the blade surfaces equal to the flux corresponding to the prescribed loading distribution. The unsteady flow due to the blade motion is simulated by solving the Reynolds-averaged Navier-Stokes equations that are discretized using a cell-vertex finite volume method in which an arbitrary Lagrangian-Eulerian formulation is used to account for the mesh movement and deformation during the design procedure. The method is first verified by inversely designing an existing blade using its loading distribution as the design target and starting from a profile that has a different camberline. The robustness, flexibility, and usefulness of this design method are demonstrated by redesigning a subsonic turbine and a transonic compressor blade for which, for the latter case, the conventional quasi-steady approach failed. The redesign cases demonstrate that the blade aerodynamic performance can be improved by carefully tailoring the target loading distribution.

0411 遷音速圧縮機翼の翼形状3次元化による性能向上研究(S47 次世代ガスタービン,S47 次世代ガスタービン)

0411 遷音速圧縮機翼の翼形状3次元化による性能向上研究(S47 次世代ガスタービン,S47 次世代ガスタービン)

Aerostructural Optimization of a Transonic Compressor Rotor

This paper presents a framework for multi-objective and multidisciplinary design optimization using highfidelity analysis tools. In this framework the aerodynamic performance is evaluated based on a Navier‐Stokes equation solver, and the structure dynamics is computed using commercially available finite element software. We employ a genetic algorithm as a robust design optimization tool to facilitate the multi-objective optimization. We also use the response surface approach to tackle the difficulties associated with the organizational complexity and computational burden inherent in the multidisciplinary optimization. The coupling between the fluid solver and structural solver is realized through a thin-plate spline interpolation algorithm. The proposed approach is then used to perform aerostructural optimization of a three-dimensional transonic compressor blade. Our numerical results show that this method can improve the existing design and reduce the required computational time by orders of magnitude.

Multi-Objective Optimization of Transonic Compressor Blade Using Evolutionary Algorithm

In this work we perform multi-objective optimization of the NASA rotor67 transonic compressor blade. Our objectives are to maximize the stage pressure ratio as well as to minimize the compressor weight. The backbones of the optimization approach consist of a genetic algorithm, a gradient-based method, and a response surface model. The genetic algorithm is used to facilitate the multi-objective optimization and to flnd the global optima of high-dimensional problems. The gradient-based method accelerates the optimization convergence rate. The response surface model, constructed to replace the computationally expensive analysis tool, reduces the computational cost. Representative solutions are selected from the Pareto-optimal front to verify against the CFD tool. Comparing with the baseline design some optimal solutions increase the stage pressure ratio by 1.8% and decrease the weight by 5.4%. A detailed study of ∞ow structure near peak e‐ciency is presented by means of pressure distribution and streamlines inside boundary layers. Our results show that the optimized blade favors a lighter weight by a thinner blade shape. The stage pressure rise is attributed to a reduced separation zone and a weaken shock wave.

Multiobjective Optimization Using Coupled Response Surface Model and Evolutionary Algorithm.

In this work we develop an efficient approach for computationally expensive multiobjective design optimization problems. In this approach we bring together design of experiment, a response surface model, a genetic algorithm, and computational-fluid-dynamics analysis tools to provide an integrated optimization system. We use an improved hypercube sampling to preselect an array of design points on which the computational-fluid-dynamics code will run. Then a computationally cheap surrogate model is constructed based on response surface approximation. A real-coded genetic algorithm is then applied on the surrogate model to perform multiobjective optimization. Representative solutions are chosen from the Pareto-optimal front to verify against the computational-fluid-dynamics code. This proposed method is used in the redesign of a single-stage turbopump, a two-stage turbopump, and the NASA rotor67 transonic compressor blade

Aerodynamic Damping Predictions in Turbomachines Using a Coupled Fluid-Structure Model

Flow-induced vibration of turbomachine blade rows is a coupled fluid-structure problem. Thus, rather than separate fluid and structural models, a coupled interacting fluid-structures analysis is needed. This need is addressed by extending the finite element code ALE3D that solves the three-dimensional Euler equations to model the unsteady aerodynamics of turbomachine blade rows. The same finite element model is applied to both the blading and the fluid, which results in consistency between the fluid and structure. Such a coupled interacting fluid-structure analysis enables the aerodynamic damping of multiple vibration modes to be predicted from two time-domain simulations: one with the blading in a vacuum and one with the blading in flow. This novel approach to predict aerodynamic damping is demonstrated by the consideration of a modern transonic compressor blade row. The blading is first impulsed in its first bending and first torsion modes in a vacuum. It is then immersed in the design-point flowfield and impulsed in its first bending and first torsion modes again. Signal processing tools applied to the predicted blade response time history extract the difference in the decay rate of both modes.

Shape Optimization of Transonic Compressor Blade Using Quasi-Three-Dimensional Flow Physics

Shape Optimization of Transonic Compressor Blade Using Quasi-Three-Dimensional Flow Physics

Shape optimization of transonic compressor blade using quasi-three-dimensional flow physics

Shape optimization of transonic compressor blade using quasi-three-dimensional flow physics

The Impact of Viscous Effects on the Aerodynamic Damping of Vibrating Transonic Compressor Blades—A Numerical Study

A parametric study which investigates the influence of viscous effects on the damping behavior of vibrating compressor cascades is presented here. To demonstrate the dependence of unsteady aerodynamic forces on the flow viscosity, a computational study was performed for a transonic compressor cascade of which the blades underwent tuned pitching oscillations while the flow conditions extended from fully subsonic to highly transonic flow. Additionally, the reduced frequency and Reynolds number were varied. In order to check the linear behavior of the aerodynamic forces, all calculations were carried out for three different oscillation amplitudes. Comparisons with inviscid Euler results helped identify the influence of viscous effects. The computations were performed with a Navier-Stokes code, the basic features of which are the use of an AUSM upwind scheme, an implicit time integration, and the implementation of the Baldwin-Lomax turbulence model. In order to demonstrate the possibility of this code to correctly predict the unsteady behavior of strong shock-boundary layer interactions, the experiment of Yamamoto and Tanida on a self-induced shock oscillation due to shock-boundary layer interaction was calculated. A significant improvement in the prediction of the shock amplitude was achieved by a slight modification of the Baldwin Lomax turbulence model. An important result of the presented compressor cascade investigations is that viscous effects may cause a significant change in the aerodynamic damping. This behavior is demonstrated by two cases in which an Euler calculation predicts a damped oscillation whereas a Navier-Stokes computation leads to an excited vibration. It was found that the reason for these contrary results are shock-boundary-layer interactions which dramatically change the aerodynamic damping.

The Design of an Advanced Civil Fan Rotor

A recent ASME paper by the authors described a quasi-three-dimensional calculation system for transonic compressor blade rows. The system predicts both the internal flow field and the overall performance of the blade row. It therefore enables the compressor engineer to optimize the blade shapes in order to improve the design point efficiency. This is explored in the present paper. A new type of blade profile has been developed to allow sufficient freedom for the optimization. Application to the design of a high-efficiency, transonic civil fan rotor is discussed.

The role of a mixed design and analysis method in a turbomachinery blade design system

AbstractIn this paper the requirements of a design method are considered, with particular emphasis given to the needs of designers of multi‐stage compressors. It is concluded that what is required is a quasi three‐dimensional compatible design and analysis method. The extension of the finite element blade‐to‐blade method FINSUP to include a design calculation compatible with the analysis calculation is presented. In this method, changes to the blade shape are modelled using a surface transpiration model. By displacing every point on the blade surface individually, the relationship between the blade shape and the surface velocity is determined. It is shown that with a Newton–Raphson procedure adopted in the method, a very efficient manner of introducing the design option is possible. As a consequence the resulting program is fast and completely interactive. An example of the method is shown in which a transonic compressor blade with a suction surface shock is redesigned to give a blade with controlled suction surface diffusion.