Impact of Surface Roughness on Compressor Cascade Performance

2010 ◽  
Vol 132 (6) ◽  
Author(s):  
Seung Chul Back ◽  
June Hyuk Sohn ◽  
Seung Jin Song
Keyword(s):  
Gas Turbines ◽  
Velocity Ratio ◽  
Compressor Blade ◽  
Compressor Cascade ◽  
Roughness Effects ◽  
Flow Angle ◽  
Angle Deviation ◽  
Blade Row ◽  
Rough Region ◽  
Linear Compressor Cascade

This paper presents an experimental investigation of roughness effects on aerodynamic performance in a low-speed linear compressor cascade. Equivalent sandgrain roughnesses of 12 μm, 180 μm, 300 μm, 425 μm, and 850 μm have been tested. In nondimensional terms, these roughnesses represent compressor blade roughnesses found in actual gas turbines. Downstream pressure, velocity, and angle have been measured with a five-hole probe at 0.3 chord downstream of the blade trailing edge. For the tested roughnesses of 180 μm, 300 μm, 425 μm, and 850 μm, the axial velocity ratio across the blade row decreases by 0.1%, 2.1%, 2.5%, and 5.4%, respectively. For the same cases, the exit flow angle deviation increases by 24%, 38%, 51%, and 70%, respectively. Finally, the mass-averaged total pressure loss increases by 12%, 44%, 132%, and 217%, respectively. Also, the loss increases more rapidly in the transitionally rough region. Thus, among the three parameters, the loss responds most sensitively to changes in compressor blade roughness.

scholarly journals Effects of Transient Heat Transfer on Compressor Stability

2018 ◽  
Vol 140 (12) ◽  
Author(s):  
A. Kiss ◽  
Z. Spakovszky
Keyword(s):  
Heat Transfer ◽  
Pressure Ratio ◽  
Strong Dependence ◽  
Transient Heat Transfer ◽  
Compressor Blades ◽  
Flow Angle ◽  
Stall Margin ◽  
Point Reduction ◽  
Angle Deviation ◽  
Blade Row

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.

Vortex Structures for Highly-Loaded Subsonic Compressor Cascades With Slot Injection

2017 ◽  
Author(s):  
Huanlong Chen ◽  
Huaping Liu ◽  
Dongfei Zhang ◽  
Linxi Li
Keyword(s):  
Suction Side ◽  
Compressor Blade ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Blade Passage ◽  
Vortex Control ◽  
Analysis Theory ◽  
Numerical Simulation Methods ◽  
Linear Compressor Cascade ◽  
Highly Loaded

A promising flow analytical way to offset the respective shortcomings for the experimental measure and numerical simulation methods is presented. First, general topological rules which are applicable to the skin-friction vector lines on the passage surface, to the flow patterns in the cross-section of the cascade as well as on the blade-to-blade surface were deduced for the turbomachinery cascades with/without suction/blowing slots in this paper. Second, the qualitative analysis theory of the differential equation was used to investigate the distribution feature of the flow singular points for the limiting streamlines equation. The topological structure of the flow pattern on the cascade passage surfaces was discussed in detail. Third, the experiment and numerical simulations results for a linear compressor cascade passage with highly-loaded compound-lean slotted blade, which were combined to topologically examine the flow structure with penetrating slot injections through the blade pressure side and suction side. The results showed that the general topological rules are applicable and effective for flow diagnosis in highly-loaded compressor blade passage with slots. Finally, an integrated vortex control model, in which the blade compound-lean effect and the injection flow through the slots were coupled, was presented. The model shows that reasonable slot injection configurations can effectively control the concentrated shedding vortices from the suction surface of a highly-loaded compressor cascades passage, thereby the aerodynamic performance for the blade passage is remarkably improved. The present work provides a novel theoretical analysis method and insights of the flow for the turbine blade passage with cooling structures, aspirated compressor blade passage and other applications with new flow control configurations in turbomachinery field.

The Use of a Finite Difference Technique to Predict Cascade, Stator, and Rotor Deviation Angles and Optimum Angles of Attack

1973 ◽  
Vol 95 (3) ◽  
pp. 185-190 ◽  
Author(s):  
P. R. Dodge
Keyword(s):  
Experimental Data ◽  
Finite Difference ◽  
Velocity Ratio ◽  
Three Dimensional ◽  
Compressor Blade ◽  
Deviation Angle ◽  
Kutta Condition ◽  
Meridional Velocity ◽  
Finite Difference Technique ◽  
Blade Row

This paper develops a method of using a quasi-three-dimensional finite difference (blade to blade) flow program to predict deviation angle. The work discussed herein was accomplished in support of Contract NAS3-15324 with Lewis Research Center of NASA, Cleveland, Ohio. An appropriate approximation to the Kutta Condition is developed. Results are compared with experimental data from cascades including the effects of meridional velocity ratio and compressibility. Predictions are extended to cases with radius changes. In addition, the same program is used to predict optimum angle of attack for a compressor blade row.

scholarly journals Investigation of Wash Fluid Preheating on the Effectiveness of Online Compressor Washing in Industrial Gas Turbines

2020 ◽  
Author(s):  
Roupa Agbadede ◽  
Biweri Kainga
Keyword(s):  
Gas Turbines ◽  
Pressure Loss ◽  
Total Pressure ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Flow Angle ◽  
Pressure Loss Coefficient ◽  
Industrial Gas Turbines ◽  
Total Pressure Loss Coefficient

Abstract This study presents an investigation of wash fluid preheating on the effectiveness of online compressor washing in industrial gas turbines. Crude oil was uniformly applied on the compressor cascade blades surfaces using a roller brush, and carborundum particles were ingested into the tunnel to create accelerated fouled blades. Demineralized water was preheated to 500C using the heat coil provided in the tank. When fouled blades washed with preheated demineralized and the one without preheating were compared, it was observed that there was little or no difference in terms of total pressure loss coefficient and exit flow angle. However, when the fouled and washed cases were compared, there was a significant different in total pressure loss coefficient and exit flow angle.

Study of Relative Endwall Motion Effects in a Compressor Cascade Through Direct Numerical Simulations

2020 ◽  
Vol 143 (1) ◽  
Author(s):  
Jordi Ventosa-Molina ◽  
Martin Lange ◽  
Ronald Mailach ◽  
Jochen Fröhlich
Keyword(s):  
Numerical Simulations ◽  
Secondary Flow ◽  
Relative Motion ◽  
Direct Numerical Simulations ◽  
Compressor Cascade ◽  
Flow Angle ◽  
Pressure Losses ◽  
Wake Effects ◽  
Secondary Flow Structure ◽  
Linear Compressor Cascade

Abstract Linear cascades are commonly used as surrogate geometries when performing fundamental studies of turbomachinery blading. Several effects are not accounted for in linear cascades, such as the relative motion between blade and endwall. In this study, three different relative endwall velocities are analyzed. The effect of the relative motion between endwall and blade in a linear compressor cascade is studied through direct numerical simulations. Results show a significant change in the secondary flow structure within the passage. Most notably, the tip leakage vortex is displaced away from the blade. Still, the blade spanwise range affected by the secondary flow field is similar to the case without relative endwall motion. At the outlet plane, a stratification of the total pressure losses and the exit flow angle is found, which overshadows any blade wake effects near the endwall.

scholarly journals Design, Optimization, and Analysis of a High-Turning Transonic Tandem Compressor Cascade

1997 ◽  
Author(s):  
Anton Weber ◽  
Wolfgang Steinert
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Side Wall ◽  
Design Flow ◽  
Basic Question ◽  
Compressor Cascade ◽  
Flow Angle ◽  
Transonic Compressor ◽  
Blade Row ◽  
Transonic Cascade

As a feasibility study for a stator guide vane a highly loaded transonic compressor stator blade row was designed, optimized, and tested in a transonic cascade facility. The flow entering the turning device with an inlet Mach number of 1.06 has to be turned by more than 60° and diffused extremely to leave the blade row without swirl. Therefore, the basic question was: Is it feasible to gain such a high amount of flow turning with an acceptable level of total pressure losses? The geometric concept chosen is a tandem cascade consisting of a transonic blade row with a flow turning of 10° followed by a subsequent high-turning subsonic cascade. The blade number ratio of the two blade rows was selected to be 1:2 (transonic: subsonic). Design and optimization have been performed using a modern Navier-Stokes flow solver under 2D assumptions by neglecting side wall boundary-layer effects. In the design process it was found to be necessary to guide the wake of the low turning transonic blade near the suction surface of the subsonic blade. Furthermore, it is advantageous to enlarge the blade spacing of the ‘wake’ passage in relation to the neighbouring one of the high turning part. The optimized design geometry of the tandem cascade was tested in the transonic cascade windtunnel of the DLR in Cologne. At design flow conditions the experiments confirmed the design target in every aspect. A flow turning of more than 60°, a static pressure ratio of 1.75, and a total pressure loss coefficient of 0.15 was measured. The working range at design inlet Mach number of 1.06 is about 3.5° in terms of the inlet flow angle. A viscous analysis of various operating points showed excellent agreement with the experimental results.

Paper 32: The Effects of Reynolds Number, Turbulence Intensity, and Axial Velocity Ratio on Compressor Blade Performance

1969 ◽  
Vol 184 (7) ◽  
pp. 93-100
Author(s):  
R. Shaw
Keyword(s):  
Reynolds Number ◽  
Axial Velocity ◽  
Free Stream ◽  
Velocity Ratio ◽  
Compressor Blade ◽  
Free Stream Turbulence ◽  
Blade Row ◽  
Reynolds Number Effects ◽  
Side Walls ◽  
Series Of Experiments

The paper describes a series of experiments that were conducted on a compressor blade of conventional geometry, in a low-speed research compressor, to determine the effect of Reynolds number on the performance of (1) an isolated rotor blade row when the free-stream turbulence is low, (2) the isolated rotor at a higher level of turbulence, and (3) the rotor row when unsteadiness is introduced by a set of inlet guide vanes of zero camber and zero stagger. In each case the axial velocity ratio at mid-span was measured and tests were carried out under similar conditions on a two-dimensional cascade of blades using a wind tunnel with porous side walls. The compressor results indicate that there are significant Reynolds number effects under these conditions, and the comparison with the cascade results demonstrates that there is a difference in performance between the rotating row and the cascade.

The Effect of Reynolds Number and Turbulence Intensity on the Performance of a Compressor Cascade with Prescribed Velocity Distribution

1977 ◽  
Vol 19 (3) ◽  
pp. 93-100 ◽  
Author(s):  
J. Citavy ◽  
J. F. Norbury
Keyword(s):  
Reynolds Number ◽  
Turbulence Intensity ◽  
Axial Velocity ◽  
Separation Bubble ◽  
Velocity Ratio ◽  
Aerodynamic Performance ◽  
Substantial Effect ◽  
Compressor Cascade ◽  
Flow Angle ◽  
Outlet Flow

Experimental results are presented on the effect of Reynolds number ( Re) and turbulence intensity ( Tu) on the aerodynamic performance of a PVD compressor cascade at design incidence. The pressure distribution, outlet flow angle and losses were measured within the ranges Re = 0.6 times 105 to 2 times 105 and Tu = 0.35 to 4.4 per cent. In some experiments, the effect of axial velocity ratio ( AVR) was investigated. A substantial effect of the Reynolds number and turbulence intensity on the growth and bursting of the separation bubble was observed, with consequent effects on the aerodynamic performance of the cascade. The bursting of the bubble also gave rise to a hysteresis effect with Reynolds number.

The Influence of Multimode Transition Initiated by Periodic Wakes on the Profile Loss of a Linear Compressor Cascade

2000 ◽  
Author(s):  
Reinhold Teusch ◽  
Stefan Brunner ◽  
Leonhard Fottner ◽  
Marius Swoboda
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Separation Bubble ◽  
Compressor Cascade ◽  
Linear Compressor ◽  
Laminar Separation ◽  
Blade Row ◽  
Linear Compressor Cascade ◽  
Profile Loss ◽  
Undisturbed Flow

This paper presents results of boundary layer and loss measurements in a high speed cascade wind tunnel on a linear compressor cascade under the influence of unsteady, periodic wakes. The wakes of an upstream blade row were simulated by cylindrical bars moved by a belt mechanism upstream of the cascade. Extensive hot-film array, hot-wire and pressure measurements with variation of steady and unsteady inlet flow conditions have been performed for a better understanding of the transition and loss mechanisms on a blade row interacting with wakes. The incoming wakes are inducing early forced transition in the boundary layer followed in time by calmed regions. Due to its higher shear stress level and its fuller velocity profile, the calmed flow is able to suppress laminar separation bubbles and to delay transition in the region with undisturbed flow between wakes, playing a significant role in the loss generation process. At the investigated low Reynolds number, where the measurements for the steady flow case showed a well-developed laminar separation bubble, reductions of profile loss up to 20% were observed for the measured configuration. In the case of the high Reynolds number, where in undisturbed flow only a small separation bubble was detected, a profile loss rise up to 30% was measured. Beside a better understanding of unsteady flow physics the goal of these basic investigations of unsteady transition is to create a wide database for the improvement of transition modeling in unsteady CFD codes.

scholarly journals Effects of Transient Heat Transfer on Compressor Stability

2018 ◽  
Author(s):  
A. Kiss ◽  
Z. Spakovszky
Keyword(s):  
Heat Transfer ◽  
Pressure Ratio ◽  
Strong Dependence ◽  
Transient Heat Transfer ◽  
Compressor Blades ◽  
Flow Angle ◽  
Stall Inception ◽  
Stall Margin ◽  
Angle Deviation ◽  
Blade Row

The effects of heat transfer between the compressor structure and primary gas path flow on compressor stability are investigated during hot engine re-acceleration transients, or so called “Bodie” 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. Transient calculations with heat transfer 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 and the remainder is attributed to stage re-matching. It is found that heat transfer increases loading in the front stages and destabilizes the front block. Furthermore, the stage re-matching due to heat transfer alters the slope of the compressor characteristic and promotes modal-type stall inception. 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, such as NPSS, only capture 1.2 points of the 8.0 point stall margin reduction. This implies that, using this new capability, opportunities exist early in the design process to address potential stability issues due to transient heat transfer.

Sign in / Sign up

Export Citation Format

Share Document