Sweep in a Transonic Fan Rotor: Part 1 — 3D Geometry Package

1998 ◽  
Author(s):  
Hazem F. Abdelhamid ◽  
Raymond P. Shreeve
Keyword(s):  
Axial Compressor ◽  
Compressor Blade ◽  
Design Parameters ◽  
Control Points ◽  
Manufacturing Optimization ◽  
3D Geometry ◽  
Blade Shape ◽  
Two Parameters ◽  
Transonic Fan ◽  
Blade Element

A geometry package was developed which uses six Bezier surfaces to describe an axial compressor blade. The blade is defined by 32 control points and two parameters, which determine the leading and trailing edge extensions. The package was used to represent a reference transonic fan rotor to within machining tolerances, and then to introduce forward and backward sweep holding blade-element design parameters fixed. Blade lean and point geometry manipulations were also demonstrated. All geometries produced by the package are machinable without approximation. The Bezier-surface representation was chosen in order to minimize the number of control points required to specify the blade shape and eventually enable aero-structural-manufacturing optimization.

A Family of Surface Velocity Distributions for Axial Compressor Blading and Their Theoretical Performance

1976 ◽  
Vol 98 (2) ◽  
pp. 229-238 ◽  
Author(s):  
G. J. Walker
Keyword(s):  
High Performance ◽  
Axial Compressor ◽  
Compressor Blade ◽  
Surface Velocity ◽  
Design Parameters ◽  
Suction Surface ◽  
Wide Range ◽  
Using Data ◽  
High Level ◽  
Theoretical Performance

The influence of free stream disturbances on transition is discussed and it is noted that significant regions of laminar flow may exist on axial turbomachine blades despite the high level of disturbance to which they are subjected. A family of surface velocity distributions giving unseparated flow on the suction surface of an axial compressor blade is derived using data from detailed boundary layer measurements on the blading of a single-stage machine. The distributions are broadly similar to those adopted by Wortmann in designing high performance isolated aerofoil sections for operation at much higher Reynolds numbers. The theoretical performance of blades having the specified surface velocity distributions is computed for a wide range of conditions, and the effects of varying Reynolds number and other design parameters are analyzed. The results suggest the possibility of obtaining useful improvements in performance over that of conventional compressor blade sections. The computed performance values show an almost unique relation between the blade losses and the suction surface diffusion ratio. However the correlation of losses with the equivalent diffusion ratio is found to break down at high values of the latter parameter.

Multiple Surrogate Modeling for Axial Compressor Blade Shape Optimization

2008 ◽  
Vol 24 (2) ◽  
pp. 301-310 ◽  
Author(s):  
Abdus Samad ◽  
Kwang-Yong Kim ◽  
Tushar Goel ◽  
Raphael T. Haftka ◽  
Wei Shyy
Keyword(s):  
Shape Optimization ◽  
Axial Compressor ◽  
Surrogate Modeling ◽  
Compressor Blade ◽  
Blade Shape

scholarly journals Numerical Simulation of Rotating Stall in Axial Compressor Blade Rows and Stages

1986 ◽  
Author(s):  
H. G. Neuhoff ◽  
K. G. Grahl
Keyword(s):  
Axial Compressor ◽  
Rotating Stall ◽  
Compressor Blade ◽  
Time Step ◽  
Time Constants ◽  
Compressor Rotor ◽  
Cascade Flow ◽  
Flow Through ◽  
Transonic Fan ◽  
Highly Loaded

Rotating stall is simulated by a time step integration procedure of the Euler equations. The prescribed compressors consist of finite inlet and outlet ducts, blade rows of finite chord lengths and a throttle without impedance. Due to this compressor model the net mass flow through the compressor remains constant during the transient to fully developed rotating stall. Results are presented for a highly loaded rotor and a transonic fan stage. Contrary to former nonlinear approaches, the presented theory indicates that the fully developed rotating stall in a single compressor rotor is not affected by the number of lobes of an initial circumferential disturbance. For a highly loaded stage the dependence of the stall parameters on the time constants of the cascade flow is demonstrated.

Flow Characteristics of an Optimized Axial Compressor Rotor Using Smooth Design Parameters

2016 ◽  
Author(s):  
Syed Moez Hussain Mahmood ◽  
Mark G. Turner ◽  
Kiran Siddappaji
Keyword(s):  
Thickness Distribution ◽  
Axial Compressor ◽  
Leading Edge ◽  
Design Tool ◽  
Smooth Transition ◽  
Design Parameters ◽  
Spanwise Direction ◽  
Aerodynamic Optimization ◽  
Blade Shape ◽  
Smooth Design

Blade designs have evolved from NACA series and free vortex assumptions to detailed meanline and forced vortex definitions. A design process is presented with numerous parametric options to explore a large design space. Smoothness in turbomachinery blade shapes is critical to an effective design. A cubic B-spline is used to control spanwise variations in the curvature definition of airfoil camber, thickness distribution, leading edge definition, inlet angles and outlet angles as parameters with a small number of control points. Varying parameters of individual blade sections requires more control variables that increases the parameter space and adds kinks in the 3D blade shape. Benefits of this smooth spanwise capability are demonstrated by linking the blade design tool with an aerodynamic optimization system. A single subsonic rotor (rotor 6 of a 10 stage axial compressor derived from the GE EEE design) has been considered as the baseline for the optimization process. Optimization is performed by varying curvature of the airfoil camberline as well as inlet and outlet angles in the spanwise direction. A single objective optimization was performed to optimize isentropic efficiency. An improvement in efficiency of 0.83% from 91.87% to 92.63% was obtained. The optimized blade geometry has a smooth transition from a traditional airfoil shape at the hub section to an S-shaped airfoil at the mid and tip sections. This unique blade shape was obtained because the airfoil camber curvature definition was allowed to vary smoothly spanwise. An S-shaped blade near the mid and tip section promotes flow to move radially downwards which allows for a reduction in entropy generation due to tip leakage flows. Entropy is used to quantify losses and improvement in efficiency.

Novel Compressor Blade Shaping Through a Free-Form Method

2017 ◽  
Vol 139 (8) ◽  
Author(s):  
Alistair John ◽  
Shahrokh Shahpar ◽  
Ning Qin
Keyword(s):  
Axial Compressor ◽  
Optimization Method ◽  
Aerodynamic Performance ◽  
Compressor Blade ◽  
Free Form ◽  
The Novel ◽  
Surface Sensitivity ◽  
Free Form Deformation ◽  
Blade Shape ◽  
Novel Design

This paper describes the use of the free-form-deformation (FFD) parameterization method to create a novel blade shape for a highly loaded, transonic axial compressor. The novel geometry makes use of precompression (via an S-shaping of the blade around midspan) to weaken the shock and improve the aerodynamic performance. It is shown how free-form-deformation offers superior flexibility over traditionally used parameterization methods. The novel design (produced via an efficient optimization method) is presented and the resulting flow is analyzed in detail. The efficiency benefit is over 2%, surpassing other results in the literature for the same geometry. The precompression effect of the S-shape is analyzed and explained, and the entropy increase across the shock (along the midblade line) is shown to be reduced by almost 80%. Adjoint surface sensitivity analysis of the datum and optimized designs is presented, showing that the S-shape is located in the region predicted to be most significant for changes in efficiency. Finally, the off-design performance of the blade is analyzed across the rotor characteristics at various speeds.

Axial Compressor Performance Restoration by Blade Profile Control

1984 ◽  
Author(s):  
William B. Roberts
Keyword(s):  
Axial Compressor ◽  
Integral Boundary ◽  
Profile Control ◽  
Profile Losses ◽  
Layer Technique ◽  
Blade Shape ◽  
Compressor Performance ◽  
Performance Penalty ◽  
Nominal Performance ◽  
Blade Element

One major source of increased compressor losses is imperfection in blade shape. These imperfections exist for new blades, and as an engine wears in service, the blade shapes become further distorted due to erosion. These imperfections cause an aerodynamic performance penalty in the form of decreased engine capability and increased fuel consumption. A simple numerical blade element panel method, combined with a integral boundary layer technique, has been used to make an aerodynamic analysis of the effect of imperfections in blade shape for compressor blading. This analysis indicated what blade shapes would give increased profile losses, and what kind of reshaping allows restoration of nominal performance. The results of this study were tested by a series of engine cell tests with JT8D engines. These tests indicate that new or used blades that have been sorted, properly re-shaped, and matched allow a superior recovery of performance compared with traditional blade refurbishment practices.

Reliable FE Modeling of Chord-Wise Blade Modes in Compressor Design Process

2010 ◽  
Author(s):  
Ilya Fedorov ◽  
Jaroslaw Szwedowicz ◽  
Wolfgang Kappis ◽  
Igor Putchkov
Keyword(s):  
Design Process ◽  
Axial Compressor ◽  
Compressor Blade ◽  
Design Parameters ◽  
Concept Design ◽  
Compressor Blades ◽  
Mechanical Integrity ◽  
Resonance Frequencies ◽  
Local Flexibility ◽  
Compressor Efficiency

To meet the highest compressor efficiency and resonance free operation, the design process of a modern compressor blade requires several iterations between the aerodynamic and mechanical integrity disciplines. The 1D beam theories, usually used in the concept design process, do not consider the local flexibility of a flat, tapered and twisted geometry of an axial compressor airfoil. Therefore, chord-wise bending resonances of the compressor blade, excited by flow field upstream and downstream, cannot be predicated in a reliable manner. In the paper, firstly the sensitivity of compressor blade vibrations is analysed in terms of airfoil design parameters, rotor coupling effects, and mistuning phenomena. Owing to high bending stiffness of a welded shaft, a numerical CWB tool is developed mainly for reliable predictions of chord-wise bending resonances of the compressor blade in the design process. Finally, the tool reliability is demonstrated by a good agreement of the numerical and experimental resonance frequencies, which have been measured with the tip-timing system at the front stage of the axial compressor in the field. Regarding the measured compressor bladed disc, the numerical sensitive study is carried out to determine an impact of contact uncertainties in the blade root on the computed resonance frequencies. The paper shows how physical uncertainties of the root contact and airfoil mistuning are involved in practical manner into the design process of compressor blades. In the design process, the presented CWB tool allows for fast and reliable mitigation of chord-wise bending resonances, which requires the collective solution between the aerodynamics and mechanical integrity disciplines, as it is illustrated in this paper.

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.

Numerical Efforts of Aerodynamic Re-Design in a Single-Stage Transonic Axial Compressor: Part 1 — Stator Design

2017 ◽  
Author(s):  
Justin (Jongsik) Oh
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Axial Flow ◽  
Single Stage ◽  
Design Flow ◽  
Design Parameters ◽  
Original Design ◽  
Circular Arc ◽  
Surge Margin ◽  
Flow Compressor

In many aerodynamic design parameters for the axial-flow compressor, three variables of tailored blading, blade lean and sweep were considered in the re-design efforts of a transonic single stage which had been designed in 1960’s NASA public domains. As Part 1, the re-design was limited to the stator vane only. For the original MCA (Multiple Circular Arc) blading, which had been applied at all radii, the CDA (Controlled Diffusion Airfoil) blading was introduced at midspan as the first variant, and the endwalls of hub and casing (or tip) were replaced with the DCA (Double Circular Arc) blading for the second variant. Aerodynamic performance was predicted through a series of CFD analysis at design speed, and the best aerodynamic improvement, in terms of pressure ratio/efficiency and operability, was found in the first variant of tailored blading. It was selected as a baseline for the next design efforts with blade lean, sweep and both combined. Among 12 variants, a case of positively and mildly leaned blades was found the most attractive one, relative to the original design, providing benefits of an 1.0% increase of pressure ratio at design flow, an 1.7% increase of efficiency at design flow, a 10.5% increase of the surge margin and a 32.3% increase of the choke margin.

scholarly journals Observations of Wake-Induced Turbulent Spots on an Axial Compressor Blade

1993 ◽  
Author(s):  
G. J. Walker ◽  
W. J. Solomon ◽  
J. P. Gostelow
Keyword(s):  
Surface Film ◽  
Axial Compressor ◽  
Adverse Pressure Gradient ◽  
Shear Layers ◽  
Compressor Blade ◽  
Ensemble Average ◽  
Transitional Flow ◽  
Instability Wave ◽  
Random Disturbance ◽  
Turbulent Spots

Measurements of transitional flow in regions of strong adverse pressure gradient on an axial compressor stator are reported. The range of observations covers separating laminar flow at transition onset, and reattachment of intermittently turbulent periodically separated shear layers. Transition was characterised by the regular appearance of turbulent spots in association with the rotor blade wake disturbances. However, the initial breakdown did not coincide with the wake passage as has usually been observed by other workers. The spots rather evolved from the growth of instability wave packets which lagged the wake passage. Data presented from the compressor blade measurements include: mean and ensemble-average velocities and associated integral parameters; distributions of total, periodic and random disturbance components; typical individual velocity fluctuation records; contours of ensemble-average random disturbance level; and boundary layer intermittency distributions. Measurements of turbulent intermittency showed a significant fall in this quantity near the wall in the reattaching flow. This has significant implications for the interpretation of transition data from surface film gage observations.

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