A HIGH-SPEED DISK ROTOR RIG DESIGN FOR TIP AEROTHERMAL RESEARCH

2021 ◽  
pp. 1-14
Author(s):  
Shaopeng Lu ◽  
Qiang Zhang ◽  
Li He
Keyword(s):  
Mach Number ◽  
High Speed ◽  
Cfd Simulation ◽  
Optical Measurement ◽  
Design Concept ◽  
System Level ◽  
Tip Leakage Flow ◽  
Component Level ◽  
Tip Leakage ◽  
Spatially Resolved

Abstract The relative casing motion can greatly vary the Over-Tip-Leakage (OTL) flow structure and aerothermal performance. The existing tip experimental research facilities including stationary linear cascade, cascade rigs with low speed moving belt, or high-speed rotor rigs are either not capable of reproducing the high relative casing Mach number, or extremely expensive and still difficult for optical measurement. This paper presents a high-speed disk rotor design which can simulate the high casing relative speed. The unique feature of this rig design concept is that it enables full optical access of the tip surface under the engine-representative OTL flow condition. In this paper, the feasibility of the design concept is demonstrated and assessed by RANS CFD simulation, both in component level and whole rig system level. First-of-its-kind experimental results of spatially-resolved tip heat transfer coefficient distribution at high relative casing Mach number are reported. The design idea demonstrated in this paper can also be useful for other tip leakage flow studies.

A High-Speed Disk Rotor Rig Design for Tip Aerothermal Research

2020 ◽  
Author(s):  
S. Lu ◽  
Q. Zhang ◽  
L. He
Keyword(s):  
High Speed ◽  
Cfd Simulation ◽  
Optical Measurement ◽  
Design Concept ◽  
System Level ◽  
Tip Leakage Flow ◽  
Component Level ◽  
Tip Leakage ◽  
Rotor Design ◽  
Wide Range

Abstract The relative casing motion can greatly vary the Over-Tip-Leakage (OTL) flow structure and thermal performance. The existing tip experimental research facilities include stationary linear cascade, cascade rigs with low speed moving belt, or high-speed rotor rigs are either not capable of reproducing the high relative casing Mach number, or extremely expensive and still difficult for optical measurement. This paper presents a highspeed disk rotor design which can simulate the high casing relative speed. The unique feature of this rig design concept is that it enables full optical access of the tip surface under the engine-representative OTL flow condition. In this paper, the feasibility of the design concept is demonstrated and assessed by RANS CFD simulation, both in component level and whole rig system level. The design idea demonstrated in this paper can be useful for a wide range of tip leakage flow studies.

Comparison of Turbine Tip Leakage Flow for Flat Tip and Squealer Tip Geometries at High-Speed Conditions

2004 ◽  
Vol 128 (2) ◽  
pp. 213-220 ◽  
Author(s):  
Nicole L. Key ◽  
Tony Arts
Keyword(s):  
Reynolds Number ◽  
High Speed ◽  
Flow Characteristics ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Cascade Arrangement ◽  
Oil Flow ◽  
Blade Tip ◽  
Tip Velocity

The tip leakage flow characteristics for flat and squealer turbine tip geometries are studied in the von Karman Institute Isentropic Light Piston Compression Tube facility, CT-2, at different Reynolds and Mach number conditions for a fixed value of the tip gap in a nonrotating, linear cascade arrangement. To the best knowledge of the authors, these are among the very few high-speed tip flow data for the flat tip and squealer tip geometries. Oil flow visualizations and static pressure measurements on the blade tip, blade surface, and corresponding endwall provide insight to the structure of the two different tip flows. Aerodynamic losses are measured for the different tip arrangements, also. The squealer tip provides a significant decrease in velocity through the tip gap with respect to the flat tip blade. For the flat tip, an increase in Reynolds number causes an increase in tip velocity levels, but the squealer tip is relatively insensitive to changes in Reynolds number.

scholarly journals Investigation on the Flow Field Entropy Structure of Non-Synchronous Blade Vibration in an Axial Turbocompressor

Entropy ◽  
2020 ◽  
Vol 22 (12) ◽  
pp. 1372
Author(s):  
Mingming Zhang ◽  
Anping Hou
Keyword(s):  
Flow Field ◽  
High Speed ◽  
Guide Vane ◽  
Leading Edge ◽  
Inlet Guide Vane ◽  
Tip Leakage Flow ◽  
Good Correspondence ◽  
Blade Vibration ◽  
Tip Leakage ◽  
Spiral Vortex

In order to explore the inducing factors and mechanism of the non-synchronous vibration, the flow field structure and its formation mechanism in the non-synchronous vibration state of a high speed turbocompressor are discussed in this paper, based on the fluid–structure interaction method. The predicted frequencies fBV (4.4EO), fAR (9.6EO) in the field have a good correspondence with the experimental data, which verify the reliability and accuracy of the numerical method. The results indicate that, under a deviation in the adjustment of inlet guide vane (IGV), the disturbances of pressure in the tip diffuse upstream and downstream, and maintain the corresponding relationship with the non-synchronous vibration frequency of the blade. An instability flow that developed at the tip region of 90% span emerged due to interactions among the incoming main flow, the axial separation backflow, and the tip leakage vortices. The separation vortices in the blade passage mixed up with the tip leakage flow reverse at the trailing edge of blade tip, presenting a spiral vortex structure which flows upstream to the leading edge of the adjacent blade. The disturbances of the spiral vortexes emerge to rotate at 54.5% of the rotor speed in the same rotating direction as a modal oscillation. The blade vibration in the turbocompressor is found to be related to the unsteadiness of the tip flow. The large pressure oscillation caused by the movement of the spiral vortex is regarded as the one of the main drivers for the non-synchronous vibration for the present turbocompressor, besides the deviation in the adjustment of IGV.

Experimental Investigation of the Tip Clearance Flow for an Axial Flow Fan Rotor Using a PIV System

2004 ◽  
Author(s):  
Xiaocheng Zhu ◽  
Wanlai Lin ◽  
Zhaohui Du
Keyword(s):  
High Pressure ◽  
High Speed ◽  
Low Pressure ◽  
Tip Clearance ◽  
Tip Leakage Flow ◽  
Discrete Vortex ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Ventilation Fan ◽  
Good Agreement

The flow field in the tip region of an axial ventilation fan is investigated with a PIV (Particle Image Velocimeter) system at the design condition. Characteristics of a ventilation fan are an extreme low-pressure difference and a large tip clearance with a low rotating speed. Flow fields with three different tip clearances are surveyed on three different circumferential planes, respectively. The phase-locked average method is used to investigate the generation and the development of a tip leakage vortex. The result from PIV system is compared with that from a PDA (Particle Dynamics Anemometer) system. Both data are in good agreement. The structure of the tip leakage vortex for the rotor is illustrated. The characteristic of a leakage vortex is described in both velocity vectors and vortical contours. It is found that the tip leakage flow for a low speed and a low pressure ventilation fan also has a chance to roll up into a discrete vortex at three different tip clearances, which is similar to high speed and high-pressure compressors and turbines. When the tip clearance increases, the scope and the location variation for the tip leakage vortex increase. Finally, the trajectories of the tip leakage vortex by the experimental measurement are compared with predictions from the existing models for high speed and high-pressure compressors and turbines. A good agreement is obtained.

scholarly journals On Scaling Method to Investigate High-Speed Over-Tip-Leakage Flow at Low-Speed Condition

2018 ◽  
Vol 140 (6) ◽  
Author(s):  
Hongmei Jiang ◽  
Li He ◽  
Qiang Zhang ◽  
Lipo Wang
Keyword(s):  
High Speed ◽  
Turbine Blades ◽  
Flow Distribution ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Scaling Method ◽  
Low Speed ◽  
Tip Leakage ◽  
Speed Condition

Modern high-pressure turbine blades operate at high-speed conditions. The over-tip-leakage (OTL) flow can be high-subsonic or even transonic. From the consideration of problem simplification and cost reduction, the OTL flow has been studied extensively in low-speed experiments. It has been assumed a redesigned low-speed blade profile with a matched blade loading should be sufficient to scale the high-speed OTL flow down to the low-speed condition. In this paper, the validity of this conventional scaling approach is computationally examined. The computational fluid dynamics (CFD) methodology was first validated by experimental data conducted in both high- and low-speed conditions. Detailed analyses on the OTL flows at high- and low-speed conditions indicate that, only matching the loading distribution with a redesigned blade cannot ensure the match of the aerodynamic performance at the low-speed condition with that at the high-speed condition. Specifically, the discrepancy in the peak tip leakage mass flux can be as high as 22%, and the total pressure loss at the low-speed condition is 6% higher than the high-speed case. An improved scaling method is proposed hereof. As an additional dimension variable, the tip clearance can also be “scaled” down from the high-speed to low-speed case to match the cross-tip pressure gradient between pressure and suction surfaces. The similarity in terms of the overall aerodynamic loss and local leakage flow distribution can be improved by adjusting the tip clearance, either uniformly or locally.

The Influence of Tip Clearance Momentum Flux on Stall Inception in a High-Speed Axial Compressor

2013 ◽  
Vol 135 (5) ◽  
Author(s):  
Joshua D. Cameron ◽  
Matthew A. Bennington ◽  
Mark H. Ross ◽  
Scott C. Morris ◽  
Juan Du ◽  
...  
Keyword(s):  
Momentum Flux ◽  
High Speed ◽  
Axial Compressor ◽  
Leading Edge ◽  
Tip Clearance ◽  
Flow Coefficient ◽  
Axial Position ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage

Experimental and numerical studies were conducted to investigate tip-leakage flow and its relationship to stall in a transonic axial compressor. The computational fluid dynamics (CFD) results were used to identify the existence of an interface between the approach flow and the tip-leakage flow. The experiments used a surface-streaking visualization method to identify the time-averaged location of this interface as a line of zero axial shear stress at the casing. The axial position of this line, denoted xzs, moved upstream with decreasing flow coefficient in both the experiments and computations. The line was consistently located at the rotor leading edge plane at the stalling flow coefficient, regardless of inflow boundary condition. These results were successfully modeled using a control volume approach that balanced the reverse axial momentum flux of the tip-leakage flow with the momentum flux of the approach fluid. Nonuniform tip clearance measurements demonstrated that movement of the interface upstream of the rotor leading edge plane leads to the generation of short length scale rotating disturbances. Therefore, stall was interpreted as a critical point in the momentum flux balance of the approach flow and the reverse axial momentum flux of the tip-leakage flow.

Numerical Investigation of Active Flow Control Using Tip Synthetic Jet on the Performance of a High-Speed Axial Compressor Rotor

2020 ◽  
Author(s):  
Guang Wang ◽  
Wuli Chu
Keyword(s):  
Flow Control ◽  
High Speed ◽  
Active Flow Control ◽  
Axial Compressor ◽  
Synthetic Jet ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Compressor Rotor ◽  
The Stability ◽  
Active Flow

Abstract In order to weaken the negative effect of tip leakage flow and improve the tip flow condition, this paper introduces synthetic jet into the flow control field of axial compressor, and proposes a method of active flow control by arranging synthetic jet at the tip. A high-speed axial compressor rotor of the author’s research group is taken as the numerical simulation object. On the basis of keeping geometric parameters of the synthetic jet actuator unchanged, this paper studies the influence of applying tip synthetic jet on aerodynamic performance of the compressor rotor at three axial positions of −10%Ca, 0%Ca and 21.35%Ca respectively. The results show that when tip synthetic jet is in the above three positions, comprehensive stability margin of the compressor rotor increases by 2.62%, 3.77% and 12.46% respectively, and efficiency near stall point increases by 0.22%, 0.25 and 0.47% respectively. This shows that when tip synthetic jet is far away from blade, the aerodynamic performance improvement of the compressor rotor is limited, and when tip synthetic jet is just above the leading edge, the effect of expanding stability is the best and the efficiency is the most improved. The mechanism of tip synthetic jet can increase the stability of the compressor rotor is that when the actuator is in the blowing stage, it can blow the low-speed air flow of blade top to downstream, and when the actuator is in the suction stage, it can suck the low-speed air flow of blade top into slot, so as to alleviate the top blockage and realize the stability expansion. The mechanism of tip synthetic jet can improve the efficiency of compressor rotor is that the blowing and suction of actuator weaken the intensity of tip leakage flow, reduce the size of vortex core and also reduce the flow loss of the compressor rotor correspondingly.

scholarly journals On Scaling Method to Investigate High-Speed Over-Tip-Leakage Flow at Low-Speed Condition

2017 ◽  
Author(s):  
Hongmei Jiang ◽  
Li He ◽  
Qiang Zhang ◽  
Lipo Wang
Keyword(s):  
High Speed ◽  
Flow Distribution ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Scaling Method ◽  
Low Speed ◽  
Aerodynamic Loss ◽  
Tip Leakage ◽  
Speed Condition

Modern High Pressure Turbine (HPT) blades operate at high speed conditions. The Over-Tip-Leakage (OTL) flow, which plays a major role in the overall loss generation for HPT, can be high-subsonic or even transonic. In practice from the consideration of problem simplification and cost reduction, the OTL flow has been studied extensively in low speed experiments. It has been assumed a redesigned low speed blade profile with a matched blade loading should be sufficient to scale the high speed OTL flow down to the low speed condition. In this paper, the validity of this conventional scaling approach is computationally examined. The CFD methodology was firstly validated by experimental data conducted in both high and low speed conditions. Detailed analyses on the OTL flows at high and low speed conditions indicate that, only matching the loading distribution with a redesigned blade cannot ensure the match of the aerodynamic performance at the low speed condition with that at the high-speed condition. Specifically, the discrepancy in the peak tip leakage mass flux can be as high as 22.2%, and the total pressure loss at the low speed condition is 10.7% higher than the high speed case. An improved scaling method is proposed hereof. As an additional dimension variable, the tip clearance can also be “scaled” down from the high speed to low speed case to match the cross-tip pressure gradient between pressure and suction surfaces. The similarity in terms of the overall aerodynamic loss and local leakage flow distribution can be improved by adjusting the tip clearance, either uniformly or locally. The limitations of this proposed method are also addressed in this paper.

Tip Leakage Losses in Subsonic and Transonic Blade-Rows

2011 ◽  
Author(s):  
Andrew P. S. Wheeler ◽  
Theodosios Korakianitis ◽  
Shashimal Banneheke
Keyword(s):  
Mach Number ◽  
Three Dimensional ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Blade Loading ◽  
Tip Leakage ◽  
Leakage Flows ◽  
Mach Numbers ◽  
Exit Mach Number ◽  
Tip Leakage Flows

In this paper the effect of blade-exit Mach number on unshrouded turbine tip-leakage flows is investigated. Previously published experimental data of a high-pressure turbine blade are used to validate a CFD code, which is then used to study the tip-leakage flow at blade-exit Mach numbers from 0.6 to 1.4. Three-dimensional calculations are performed of a flat-tip and a cavity-tip blade. Two-dimensional calculations are also performed to show the effect of various squealer-tip geometries on an idealized tip-flow. The results show that as the blade-exit Mach number is increased the tip leakage flow becomes choked. Therefore the tip-leakage flow becomes independent of the pressure difference across the tip and hence the blade-loading. Thus the effect of the tip-leakage flow on overall blade loss reduces at blade-exit Mach numbers greater than 1.0. The results suggest that for transonic blade-rows it should be possible to raise blade loading within the tip region without increasing tip-leakage loss.

scholarly journals SHOCKWAVE STUDY ON THE WINGS NACA 0012, NACA 64 - 206, AND NASA SC (2) - 0706 WITH Λ = 15O AT 0.85 MACH NUMBER

2021 ◽  
Vol 2 (1) ◽  
pp. 025-032
Author(s):  
Dewi Puspitasari ◽  
Kasyful Warist Kiat
Keyword(s):  
Mach Number ◽  
High Speed ◽  
Cfd Simulation ◽  
Lower Surface ◽  
Aerodynamic Performance ◽  
Ansys Fluent ◽  
Aircraft Wings ◽  
Contour Plots ◽  
The Difference ◽  
Naca 0012

Airfoil is used as a basic form on aircraft wings. Airfoil on the wing of the aircraft is used to produce lift that will lift the fuselage into the air. Lifting force results from the difference in pressure between the upper surface and the lower surface of an aircraft wings. In high speed flights shockwave will occur at certain parts of the wing which will adversely affect the aerodynamic performance of the wing. Wing aerodynamic performance at high speeds can be improved in various ways, one of which is by giving a angle to the wing span called a swept angle. This study will use 3D CFD simulation methods using Ansys Fluent. The airfoil used are NACA 0012, NACA 64-206, and NASA SC (2) -0706 with a chord length of 1 m, AR = 5, and λ = 1 with backward swept angle Λ = 15 °. Free stream flow is air flowing with Mach Number = 0,85 at sea level and steady conditions. Based on the simulation results, shock occurs on the upper and lower surfaces for NACA 0012 with Cl = 0 due to symmetric airfoil, whereas shock occurs only on the upper surface for NACA 64-206 and NASA SC (2) - 0706 with a Cl / Cd value of 18.55 ( NACA 64-206) and 20.78 (NASA SC (2) - 0706). This simulation also provides a visual representation of Mach Number contour plots in the middle stretch (Midspan) of the wing and Cl and Cd data.

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