Pre-Deformation Method for Manufactured Compressor Blade Based on Load Incremental Approach

2020 ◽  
Vol 37 (3) ◽  
pp. 259-265
Author(s):  
Kang Da ◽  
Wang Yongliang ◽  
Zhong Jingjun ◽  
Liu Zihao
Keyword(s):  
Compressor Blade ◽  
Aerodynamic Load ◽  
Deformation Method ◽  
Transonic Compressor ◽  
Incremental Approach ◽  
Compressor Rotor ◽  
Centrifugal Load ◽  
Blade Shape ◽  
Stationary Shape ◽  
Deformation Procedure

AbstractThe blade deformation caused by aerodynamic and centrifugal loads during operating makes blade configurations different from their stationary shape. Based on the load incremental approach, a novel pre-deformation method for cold blade shape is provided in order to compensate blade deformation under running. Effect of nonlinear blade stiffness is considered by updating stiffness matrix in response to the variation of blade configuration when calculating deformations. The pre-deformation procedure is iterated till a converged cold blade shape is obtained. The proposed pre-deformation method is applied to a transonic compressor rotor. Effect of load conditions on blade pre-deformation is also analyzed. The results show that the pre-deformation method is easy to implement with fast convergence speed. Neither the aerodynamic load nor centrifugal load can be neglected in blade pre-deformation.

Parametric Bird Strike Study of a Transonic Rotor Using Isogeometric Analysis

2016 ◽  
Author(s):  
Sam Duckitt ◽  
Chiara Bisagni ◽  
Shahrokh Shahpar
Keyword(s):  
Isogeometric Analysis ◽  
Control Points ◽  
Bird Strike ◽  
Transonic Compressor ◽  
Mechanical Constraints ◽  
Compressor Rotor ◽  
Centrifugal Load ◽  
Design Variables ◽  
Particle Hydrodynamics ◽  
The Impact

This paper investigates the use of isogeometric analysis (IGA) to study high velocity impact on a transonic compressor rotor resulting from a bird strike. An approach is developed for creating volumetric NURBS blade models which are suitable for IGA. A newly implemented 3D solid NURBS element within the development version of LS-Dyna is validated against finite elements for the NASA rotor 37 under a steady centrifugal load. The smoothed particle hydrodynamics (SPH) method is then used to simulate impact from a bird strike. As a preliminary assessment for multi-disciplinary optimisation (MDO), with the objective to improve aerodynamic performance whilst satisfying mechanical constraints from impact, a number of different blade designs are created by modifying the NURBS control points directly. Hence the control points used in analysis can also be used in the design space. This approach eliminates the need for re-meshing, highlighting the advantages that IGA can bring to design optimisation, since without filtering, moving finite element nodes can result in non-smooth geometries. NURBS parametrisations are also more efficient resulting in fewer design variables, thereby accelerating the optimisation process. The effect of blade sweep, lean, twist and thickness on the impact response are investigated. The results in this paper show the promise that IGA holds in this field but some limitations of the current LS-Dyna implementation are also discussed.

Development of Hub Corner Stall and Its Influence on the Performance of Axial Compressor Blade Rows

1999 ◽  
Vol 121 (1) ◽  
pp. 67-77 ◽  
Author(s):  
C. Hah ◽  
J. Loellbach
Keyword(s):  
Numerical Solutions ◽  
Stokes Equations ◽  
Flow Direction ◽  
Three Dimensional ◽  
Suction Side ◽  
Compressor Blade ◽  
Suction Surface ◽  
Simple Diffusion ◽  
Transonic Compressor ◽  
Compressor Rotor

A detailed investigation has been performed to study hub corner stall phenomena in compressor blade rows. Three-dimensional flows in a subsonic annular compressor stator and in a transonic compressor rotor have been analyzed numerically by solving the Reynolds-averaged Navier–Stokes equations. The numerical results and the existing experimental data are interrogated to understand the mechanism of compressor hub corner stall. Both the measurements and the numerical solutions for the stator indicate that a strong twisterlike vortex is formed near the rear part of the blade suction surface. Low-momentum fluid inside the hub boundary layer is transported toward the suction side of the blade by this vortex. On the blade suction surface near the hub, this vortex forces fluid to move against the main flow direction and a limiting stream surface is formed near the hub. The formation of this vortex is the main mechanism of hub corner stall. When the aerodynamic loading is increased, the vortex initiates further upstream, which results in a larger corner stall region. For the transonic compressor rotor studied in this paper, the numerical solution indicates that a mild hub corner stall exists at 100 percent rotor speed. The hub corner stall, however, disappears at the reduced blade loading, which occurs at 60 percent rotor design speed. The present study demonstrates that hub corner stall is caused by a three-dimensional vortex system and that it does not seem to be correlated with a simple diffusion factor for the blade row.

Development of Hub Corner Stall and its Influence on the Performance of Axial Compressor Blade Rows

1997 ◽  
Author(s):  
Chunill Hah ◽  
James Loellbach
Keyword(s):  
Numerical Solutions ◽  
Stokes Equations ◽  
Flow Direction ◽  
Three Dimensional ◽  
Suction Side ◽  
Compressor Blade ◽  
Suction Surface ◽  
Simple Diffusion ◽  
Transonic Compressor ◽  
Compressor Rotor

A detailed investigation has been performed to study hub corner stall phenomena in compressor blade rows. Three-dimensional flows in a subsonic annular compressor stator and in a transonic compressor rotor have been analyzed numerically by solving the Reynolds-averaged Navier-Stokes equations. The numerical results and the existing experimental data are interrogated to understand the mechanism of compressor hub corner stall. Both the measurements and the numerical solutions indicate that a strong twister-like vortex is formed near the rear part of the blade suction surface. Low momentum fluid inside the hub boundary layer is transported toward the suction side of the blade by this vortex. On the blade suction surface near the hub, this vortex forces fluid to move against the main flow direction and a limiting stream surface is formed near the hub. The formation of this vortex is the main mechanism of hub corner stall. When the aerodynamic loading is increased, the vortex initiates further upstream, which results in a larger corner stall region. For the transonic compressor rotor studied in this paper, the numerical solution and the measured data indicate that a mild hub corner stall exists at 100 percent rotor speed. The hub corner stall, however, disappears at the reduced blade loading which occurs at 60 percent rotor design speed. The present study demonstrates that hub corner stall is caused by a three-dimensional vortex system and that it does not seem to be correlated with a simple diffusion factor for the blade row.

An Entropy Based Evaluation of Efficiency of a Transonic Compressor Rotor With Water Injection

2008 ◽  
Author(s):  
Istvan Szabo ◽  
Mark G. Turner
Keyword(s):  
Viscous Dissipation ◽  
Entropy Generation ◽  
Dissipation Function ◽  
Entropy Generation Rate ◽  
Thermodynamic Efficiency ◽  
Compression Process ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Wet Compression ◽  
Isentropic Efficiency

Defining the thermodynamic efficiency of the wet compression process in a compressor is not trivial, since the flow in this case has multiple phases present which interact with each other. In this paper, an approach is presented that calculates the overall entropy creation and thus the isentropic efficiency of a wet compression process in a transonic compressor rotor. The viscous dissipation function is calculated everywhere in the domain in the post-processing phase of the CFD simulation and integrated to the wall, with special treatment in the near-wall regions where high rates of entropy generation occur. The isentropic efficiency of the wet compression is then determined from the entropy generation rate. Analytical integration of wall functions and numerical integration of the viscous dissipation function allows for reasonable results even with relatively coarse grids and can be applied for single-phase flows. The methodology presented is also useful to quantify the efficiency of thermodynamic processes in devices that introduce streams into the flow path, such as cooled turbines and compressors with flow control.

Experimental Investigation During Stall and Surge in a Transonic Fan Stage and Rotor-Only Configuration

2006 ◽  
Author(s):  
A. J. Gannon ◽  
G. V. Hobson ◽  
R. P. Shreeve ◽  
I. J. Villescas
Keyword(s):  
High Speed ◽  
Fast Response ◽  
Pressure Measurements ◽  
Post Processing ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
The Difference ◽  
Data Signal ◽  
Transonic Fan ◽  
Casing Pressure

High-speed pressure measurements of a transonic compressor rotor-stator stage and rotor-only configuration during stall and surge are presented. Rotational speed data showed the difference between the rotor-only case and rotor-stator stage. The rotor-only case stalled and remained stalled until the control throttle was opened. In the rotor-stator stage the compressor surged entering a cyclical stalling and then un-stalling pattern. An array of pressure probes was mounted in the case wall over the rotor for both configurations of the machine. The fast response probes were sampled at 196 608 Hz as the rotor was driven into stall. Inspection of the raw data signal allowed the size and speed of the stall cell during its growth to be investigated. Post-processing of the simultaneous signals of the casing pressure showed the development of the stall cell from the point of inception and allowed the structure of the stall cell to be viewed.

Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor

1996 ◽  
Author(s):  
Chunill Hah ◽  
Douglas C. Rabe ◽  
Thomas J. Sullivan ◽  
Aspi R. Wadia
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Viscous Flow ◽  
Total Pressure ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Aerodynamic Loss ◽  
Transonic Compressor ◽  
Inlet Distortion ◽  
Compressor Rotor

The effects of circumferential distortions in inlet total pressure on the flow field in a low-aspect-ratio, high-speed, high-pressure-ratio, transonic compressor rotor are investigated in this paper. The flow field was studied experimentally and numerically with and without inlet total pressure distortion. Total pressure distortion was created by screens mounted upstream from the rotor inlet. Circumferential distortions of 8 periods per revolution were investigated at two different rotor speeds. The unsteady blade surface pressures were measured with miniature pressure transducers mounted in the blade. The flow fields with and without inlet total pressure distortion were analyzed numerically by solving steady and unsteady forms of the Reynolds-averaged Navier-Stokes equations. Steady three-dimensional viscous flow calculations were performed for the flow without inlet distortion while unsteady three-dimensional viscous flow calculations were used for the flow with inlet distortion. For the time-accurate calculation, circumferential and radial variations of the inlet total pressure were used as a time-dependent inflow boundary condition. A second-order implicit scheme was used for the time integration. The experimental measurements and the numerical analysis are highly complementary for this study because of the extreme complexity of the flow field. The current investigation shows that inlet flow distortions travel through the rotor blade passage and are convected into the following stator. At a high rotor speed where the flow is transonic, the passage shock was found to oscillate by as much as 20% of the blade chord, and very strong interactions between the unsteady passage shock and the blade boundary layer were observed. This interaction increases the effective blockage of the passage, resulting in an increased aerodynamic loss and a reduced stall margin. The strong interaction between the passage shock and the blade boundary layer increases the peak aerodynamic loss by about one percent.

Flow Field Unsteadiness in the Tip Region of a Transonic Compressor Rotor

1997 ◽  
Vol 119 (1) ◽  
pp. 122-128 ◽  
Author(s):  
S. L. Puterbaugh ◽  
W. W. Copenhaver
Keyword(s):  
Flow Field ◽  
High Performance ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Vortex Interaction ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Operating Points

An experimental investigation concerning tip flow field unsteadiness was performed for a high-performance, state-of-the-art transonic compressor rotor. Casing-mounted high frequency response pressure transducers were used to indicate both the ensemble averaged and time varying flow structure present in the tip region of the rotor at four different operating points at design speed. The ensemble averaged information revealed the shock structure as it evolved from a dual shock system at open throttle to an attached shock at peak efficiency to a detached orientation at near stall. Steady three-dimensional Navier Stokes analysis reveals the dominant flow structures in the tip region in support of the ensemble averaged measurements. A tip leakage vortex is evident at all operating points as regions of low static pressure and appears in the same location as the vortex found in the numerical solution. An unsteadiness parameter was calculated to quantify the unsteadiness in the tip cascade plane. In general, regions of peak unsteadiness appear near shocks and in the area interpreted as the shock-tip leakage vortex interaction. Local peaks of unsteadiness appear in mid-passage downstream of the shock-vortex interaction. Flow field features not evident in the ensemble averaged data are examined via a Navier-Stokes solution obtained at the near stall operating point.

Characterization of Shock Wave–Endwall Boundary Layer Interactions in a Transonic Compressor Rotor

1988 ◽  
Vol 110 (3) ◽  
pp. 386-392 ◽  
Author(s):  
D. C. Rabe ◽  
A. J. Wennerstrom ◽  
W. F. O’Brien
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Air Force ◽  
Leading Edge ◽  
Transonic Compressor ◽  
Laser Measurements ◽  
Compressor Rotor ◽  
Boundary Layer Interaction ◽  
Air Force Base

The passage shock wave–endwall boundary layer interaction in a transonic compressor was investigated with a laser transit anemometer. The transonic compressor used in this investigation was developed by the General Electric Company under contract to the Air Force. The compressor testing was conducted in the Compressor Research Facility at Wright-Patterson Air Force Base, OH. Laser measurements were made in two blade passages at seven axial locations from 10 percent of the axial blade chord in front of the leading edge to 30 percent of the axial blade chord into the blade passage. At three of these axial locations, laser traverses were taken at different radial immersions. A total of 27 different locations were traversed circumferentially. The measurements reveal that the endwall boundary layer in this region is separated from the core flow by what appears to be a shear layer where the passage shock wave and all ordered flow seem to end abruptly.

scholarly journals A Three-Dimensional Shock Loss Model Applied to an Aft-Swept, Transonic Compressor Rotor

1996 ◽  
Author(s):  
Steven L. Puterbaugh ◽  
William W. Copenhaver ◽  
Chunill Hah ◽  
Arthur J. Wennerstrom
Keyword(s):  
Flow Rate ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Numerical Results ◽  
Design Flow ◽  
Design System ◽  
Loss Model ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Shock Loss

An analysis of the effectiveness of a three-dimensional shock loss model used in transonic compressor rotor design is presented. The model was used during the design of an aft-swept, transonic compressor rotor. The demonstrated performance of the swept rotor, in combination with numerical results, is used to determine the strengths and weaknesses of the model. The numerical results were obtained from a fully three-dimensional Navier-Stokes solver. The shock loss model was developed to account for the benefit gained with three-dimensional shock sweep. Comparisons with the experimental and numerical results demonstrated that shock loss reductions predicted by the model due to the swept shock induced by the swept leading edge of the rotor were exceeded. However, near the tip the loss model under-predicts the loss because the shock geometry assumed by the model remains swept in this region while the numerical results show a more normal shock orientation. The design methods and the demonstrated performance of the swept rotor is also presented. Comparisons are made between the design intent and measured performance parameters. The aft-swept rotor was designed using an inviscid axisymmetric streamline curvature design system utilizing arbitrary airfoil blading geometry. The design goal specific flow rate was 214.7 kg/sec/m2 (43.98 lbm/sec/ft2), the design pressure ratio goal was 2.042, and the predicted design point efficiency was 94.0. The rotor tip sped was 457.2 m/sec (1500 ft/sec). The design flow rate was achieved while the pressure ratio fell short by 0.07. Efficiency was 3 points below prediction, though at a very high 91 percent. At this operating condition the stall margin was 11 percent.

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