Axial-Flow Compressor Turning Angle and Loss by Inviscid-Viscous Interaction Blade-to-Blade Computation

1980 ◽  
Vol 102 (1) ◽  
pp. 28-34 ◽  
Author(s):  
E. C. Hansen ◽  
G. K. Serovy ◽  
P. M. Sockol
Keyword(s):  
Boundary Layers ◽  
Axial Flow ◽  
Inviscid Flow ◽  
Interactive System ◽  
Surface Boundary ◽  
Blade Surface ◽  
Axially Symmetric ◽  
Viscous Interaction ◽  
Interaction Scheme ◽  
Flow Compressor

A method for computation of the flow field around an arbitrary airfoil cascade on an axially symmetric blade-to-blade surface was developed which takes into account the development and separation of the blade surface boundary layers and mixing in the wake. The method predicts the overall fluid turning and total pressure loss in the context of an inviscid-viscous interaction scheme. The inviscid flow solution is obtained from a compressible flow matrix method. The viscous flow is obtained from a differential boundary layer method which calculates laminar, transitional and turbulent boundary layers. Provisions for the calculation of laminar and turbulent separation regions were added to the viscous scheme. The combined inviscid-viscous interaction scheme described yields results which are quantitatively consistent with experimental data. This suggests that the physical basis for the interactive system is correct and justifies continued exploration and use of the method.

scholarly journals Comparison of Calculated and Experimental Cascade Performance for Controlled-Diffusion Compressor Stator Blading

1986 ◽  
Vol 108 (1) ◽  
pp. 42-50 ◽  
Author(s):  
N. L. Sanger ◽  
R. P. Shreeve
Keyword(s):  
Aspect Ratio ◽  
Flow Field ◽  
Boundary Layers ◽  
Inviscid Flow ◽  
Reynolds Numbers ◽  
Surface Boundary ◽  
Performance Parameters ◽  
Blade Surface ◽  
Controlled Diffusion ◽  
Inlet Angle

The midspan section of a previously reported controlled-diffusion compressor stator has been experimentally evaluated in cascade. Measurements were taken over a range of incidence angles for blade chord Reynolds numbers from 470,000 to 690,000. Blade chord length was 12.7 cm, aspect ratio was 2.0, and solidity was 1.67. Measurements included conventional cascade performance parameters as well as blade surface pressures. Computations were made for the inviscid flow field, surface boundary layers, and loss for several of the blade inlet angle conditions, and compared against corresponding data.

Boundary Layer Measurement on the Blade Surface of a Multi-Stage Axial Flow Compressor

2003 ◽  
Author(s):  
Y. H. Shin ◽  
R. L. Elder ◽  
I. Bennett
Keyword(s):  
Boundary Layer ◽  
Boundary Layer Flow ◽  
Axial Flow ◽  
Velocity Profiles ◽  
Surface Boundary ◽  
Blade Surface ◽  
Suction Surface ◽  
Axial Flow Compressor ◽  
Layer Flow ◽  
Flow Compressor

This study presents experimental investigations into blade suction surface boundary layer flow in a multistage axial flow compressor. The experiments were focused on the third stage of the 4-stage Low Speed Research Compressor (LSRC) at Cranfield University. Measurements within the boundary layer were obtained using a hot wire probe. This was traversed normal to the blade surface at small increments, capturing the unsteady velocity profile within the boundary layer. Detailed boundary layer flow measurements covering most of the stator suction surface were taken and are described using time mean and ensemble averaged velocity profiles. Turbulence intensity in the boundary layer flow on the blade suction surface is also discussed. A strong wake-induced strip zone due to passing wake disturbances are generated at midspan near the blade leading edge at rotor blade passing frequency. Corner separation was observed at the tip region near the trailing edge. Normalized velocity profiles in this region show no variation in time.

scholarly journals Computation of Secondary Flows in an Axial Flow Compressor Including Inviscid and Viscous Flow Computation

1984 ◽  
Author(s):  
Francis Leboeuf
Keyword(s):  
Boundary Layer ◽  
Viscous Flow ◽  
Axial Flow ◽  
Inviscid Flow ◽  
Secondary Flows ◽  
Computational Method ◽  
Flow Computation ◽  
Axial Flow Compressor ◽  
Flow Compressor ◽  
Highly Loaded

A computational method for secondary flows in a compressor has been extended to treat stalled flows. An integral equation is used which simulates the inviscid flow at the wall, under the viscous flow influence. We present comparisons with experimental results for a 2D stalled boundary layer, and for the secondary flow in a highly loaded stator of an axial flow compressor.

Off-Design Transition and Separation Behavior of a CDA Cascade

1996 ◽  
Vol 118 (2) ◽  
pp. 204-210 ◽  
Author(s):  
W. Steinert ◽  
H. Starken
Keyword(s):  
Three Dimensional ◽  
Axial Flow ◽  
Experimental Information ◽  
Complete Separation ◽  
Surface Boundary ◽  
Compressor Cascade ◽  
Adiabatic Wall ◽  
Design Data ◽  
Flow Compressor ◽  
Separation Behavior

The design of modern axial flow compressor blade sections as well as the code validation require experimental information about the transition and separation behavior of blade surface boundary layers. The experience has shown in the past that such information has to be obtained on the whole surface and not only by point measurements because both transition and separation may be of a three-dimensional nature even in a straight cascade. Therefore, a new visualization technique based on Liquid Crystals (LC), showing the adiabatic wall temperature, has been developed. With this method, transition, local separation, and complete separation can be detected. Design and off-design data of a subsonic (M1 = 0.62) Controlled Diffusion Airfoil (CDA) compressor cascade measured in a wind tunnel are presented. The LC results are supplemented by ink-injection tests and overall performance data.

A Computation and Comparison With Measurements of Transonic Flow in an Axial Compressor Stage With Shock and Boundary Layer Interaction

1982 ◽  
Vol 104 (2) ◽  
pp. 510-515 ◽  
Author(s):  
U. K. Singh
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Axial Flow ◽  
Surface Boundary ◽  
Compressor Stage ◽  
Blade Surface ◽  
Layer Interaction ◽  
Flow Adjustment ◽  
Displacement Thickness ◽  
Boundary Layer Interaction

The flow field within a transonic axial flow compressor stage has been computed using a three-dimensional time-marching technique. Limited viscous effects are considered by including a calculation of the blade surface boundary layers. The boundary layer calculation forms an integral part of the whole computation scheme, which consists of, respectively: (i) inviscid Mach number calculations, (ii) blade surface boundary layer displacement thickness calculations, (iii) inviscid Mach number calculations with mass flow adjustment (based on the calculated displacement thicknesses) on the blade surfaces. The boundary layer computation is done by using integral calculation methods and has specifically been developed to account for a shock and boundary layer interaction (should one exist). Comparisons are made with measured results obtained with an advanced laser velocimeter. The calculated Mach number contours are in extremely good agreement with the experimental results. It is concluded that the calculation technique is a useful tool in the design of transonic axial flow turbomachines.

Numerical Investigations on Influence of Uniform Blade Surface Roughness on the Performance Characteristics of a Transonic Axial Flow Compressor Stage

2017 ◽  
Author(s):  
Ravi J. Chotalia ◽  
Dilipkumar Bhanudasji Alone
Keyword(s):  
Surface Roughness ◽  
Steady State ◽  
Axial Flow ◽  
Performance Degradation ◽  
Aviation Industry ◽  
Compressor Stage ◽  
Blade Surface ◽  
Axial Flow Compressor ◽  
Flow Compressor ◽  
The Impact

Application of surface roughness to rotating mechanical bodies will result into performance degradation. In Aviation Industry, one of the most affecting causes for performance or efficiency degradation of gas turbine engine is the blade surface roughness. The aerosols which are very small particles in the atmosphere having diameters in the microns, impinges to the compressor blade inside the aircraft engine at higher altitudes. The aerosols damages surfaces of the compressor blades. Despite of having small dimensions, due to higher velocity of the aircraft, aerosol’s impinging creates roughened surfaces and fouling. This paper is an attempt to numerically evaluate the performance degradation of the single stage transonic axial flow compressor due to uniform roughness created by the aerosols. Various cases with different roughness on various sections of the blades are analyzed to study and identify which section of the blade is more influenced by roughness. The transonic axial flow compressor has a capability of producing 1.36 stage total pressure ratio, swallowing air mass flow rate of 23 kg/s at rated design speed of 12930 rpm is used for the steady state numerical analysis. A systematic steady state 3-dimensional numerical study using solver with SST k-ω turbulence model has been carried out to evaluate the impact of blade surface roughness on the performance of compressor stage. Moreover, cases with the aerosols having different dimensions and their resulting effect is also studied to find out how performance varies when the aircraft enters into atmosphere having big aerosols from the atmosphere having smaller one and vice-e-versa.

Effect of Incoming Wakes on the Stator Performance in a Single-Stage Low Speed Axial Flow Compressor Operating at Design and Near Stall Conditions

2016 ◽  
Author(s):  
Guillaume Pallot ◽  
Dai Kato ◽  
Wataru Kanameda ◽  
Yutaka Ohta
Keyword(s):  
Boundary Layers ◽  
Total Pressure ◽  
Axial Flow ◽  
Operating Conditions ◽  
Single Stage ◽  
Generation Process ◽  
Rotor Wake ◽  
Low Speed ◽  
Pressure Losses ◽  
Flow Compressor

Unsteady flow phenomena can significantly influence the performance of turbomachines. The convection of the wake coming from a rotor into a downstream stator is one of these phenomena. In the case of compressors, when the rotor wake is transported through a downstream stator, it undergoes viscous mixing and stretching (Smith 1966), which are two mechanisms responsible for its attenuation. The flow field of a low speed single-stage compressor comprising a rotor and a downstream stator is computed using unsteady CFD simulations at design and near stall conditions. Simulations results are compared to steady and unsteady data obtained from yawmeter and hotwire measurements at both rotor and stator exit. The study focuses on the rotor wake attenuation and the related unsteady total pressure loss generated in the stator passage. The loss due to viscous mixing of the rotor wake is calculated analytically using a wake dissipation model. Based on experimental, numerical and analytical results, a break-down of the unsteady total pressure losses is performed for the two operating conditions. Unsteady total pressure losses are classified into two categories. The first category is the loss generated by viscous mixing of the rotor wake and the second one the loss generated by the interactions between the rotor wake and the stator pressure and suction surfaces boundary layers (interaction loss). Results show that the interactions between the rotor wake and the stator surfaces boundary layers play an important part in the unsteady loss generation process and that the contribution of this interaction loss increases from design to near stall condition.

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