scholarly journals Encounters with Vortices in a Turbine Nozzle Passage

2012 ◽  
Vol 2012 ◽  
pp. 1-10 ◽  
Author(s):  
J. P. Gostelow ◽  
A. Mahallati ◽  
W. E. Carscallen ◽  
A. Rona
Keyword(s):  
Convex Surface ◽  
Leading Edge ◽  
Surface Flow ◽  
Secondary Flows ◽  
Energy Separation ◽  
Streamwise Vorticity ◽  
Suction Surface ◽  
Time Resolved ◽  
Wide Range ◽  
Horseshoe Vortices

Experiments were conducted on the flow through a transonic turbine cascade. Secondary flows and a wide range of vortex types were encountered, including horseshoe vortices, shock-induced passage vortices, and streamwise vortices on the suction surface. In the separation region on the suction surface, a large rollup of passage vorticity occurred. The blunt leading edge gave rise to strong horseshoe vortices and secondary flows. The suction surface had a strong convex curvature over the forward portion and was quite flat further downstream. Surface flow visualization was performed and this convex surface displayed coherent streamwise vorticity. At subsonic speeds, strong von Kármán vortex shedding resulted in a substantial base pressure deficit. For these conditions, time-resolved measurements were made of the Eckert-Weise energy separation in the blade wake. At transonic speeds, exotic shedding modes were observed. These phenomena all occurred in experiments on the flow around one particular turbine nozzle vane in a linear cascade.

The Role of Streamwise Vorticity in Flows Over Turbomachine Blade Suction Surfaces

2011 ◽  
Author(s):  
J. P. Gostelow ◽  
W. A. McMullan ◽  
G. J. Walker ◽  
A. Mahallati
Keyword(s):  
Leading Edge ◽  
Surface Flow ◽  
Streamwise Vorticity ◽  
Suction Surface ◽  
Laminar Separation ◽  
Eddy Simulation ◽  
Low Reynolds Number Flows ◽  
Surface Time ◽  
New Findings ◽  
Streamwise Streaks

Streamwise streaks and vortices are frequently encountered in low Reynolds number flows over blading. Observations have shown that, in addition to flows over concave pressure surfaces, convex suction surfaces are also influenced by streamwise vortices. These observations are based on surface flow visualization studies and computational work with highly resolved Large Eddy Simulation. Fine scale organized streaks exist in the laminar regions of turbine and compressor blading and are predictable. For a turbine blade with a blunt leading edge, at Reynolds numbers typical of aircraft cruise conditions, the streamwise vorticity may persist, on a time-average basis, to influence the entire suction surface. Time resolution is required to capture the flow complexity that is fundamental for an understanding of the physical behavior of the laminar boundary layer and its separation and transition. Progress has been made in modeling and predicting transition and laminar separation and the new findings of interesting vortical behavior need to be incorporated. In the leading edge region spanwise vorticity may promote early transition and bubble closure; further downstream streamwise vorticity may become established. The physics of this streamwise vorticity imposes severe requirements on the temporal and spatial resolution of both experimental and computational methods. A narrow spanwise computational strip does not allow the streamwise vorticity to settle into an organized pattern; if it is to become organized, an adequate spanwise domain is required.

The Application of Non-Axisymmetric Profiled End-Walls for Axial Flow Turbines in the Embedded Stage Environment

2013 ◽  
Author(s):  
Andreas Lintz ◽  
Liping Xu ◽  
Marios Karakasis
Keyword(s):  
Axial Flow ◽  
Leading Edge ◽  
Surface Flow ◽  
Secondary Flows ◽  
Experimental Results ◽  
Front Part ◽  
Flow Conditions ◽  
Suction Surface ◽  
Inlet Conditions ◽  
End Wall

In this paper, an assessment of the effectiveness of non-axisymmetric profiled end-walls in the embedded stage environment at varying inlet conditions is presented. Both numerical and experimental results were obtained in a three-stage model turbine which offers flow conditions representative of embedded blade rows in a typical high pressure steam turbine. The end-wall profile design was carried out using automatic optimization in conjunction with 3D RANS CFD. The design target is to reduce the end-wall losses by reducing the loading in the front part of the passage, which resulted in a single trough close to the blade suction surface in the leading edge region. 5-hole probe traverses and surface flow visualization show that the intensity of the secondary flows is reduced by about 10%, but overall loss is only reduced slightly. Experimental results have been obtained for the cylindrical end-wall and three different trough depths. With increasing depth, transitional effects at the end-walls might come into play, increasing the total pressure loss in the boundary layer region. The effects of the end-wall design is similar at positive and negative incidence, despite the reduced loading in the front part of the passage at negative incidence. At very high negative incidence angles, such as those occurring at the stator tip with rotor shroud leakage flows, the mechanism of secondary flow generation changes, so that a design under nominal inlet flow conditions shows no effect on the exit flow field.

Measured Endwall Flow and Passage Heat Transfer in a Linear Blade Passage With Endwall and Leading Edge Modifications

2007 ◽  
Author(s):  
Gazi I. Mahmood ◽  
Sumanta Acharya
Keyword(s):  
Static Pressure ◽  
Leading Edge ◽  
Suction Side ◽  
Axial Direction ◽  
Secondary Flows ◽  
Blade Surface ◽  
Streamwise Vorticity ◽  
Suction Surface ◽  
Throat Region ◽  
Nusselt Numbers

The role of contouring via leading edge fillets or 3-dimensional end wall contouring is examined and compared. The contour endwall profile varies along both the pitch-direction and the axial direction and represents full-passage contouring. The fillets are employed at the corner of blade leading edge and endwall and extend about one-third of the blade axial chord along the blade. Both the endwall contouring and leading edge fillet are employed at the passage bottom endwall. Measurements are obtained for static pressure, pitchwise velocity, flow yaw angle, streamwise vorticity, and Nusselt number on the endwall. The Reynolds number based on the axial chord and passage inlet velocity for the measurements is 2.39×105. The blade surface static pressure coefficients near the endwall indicate no effects of the leading fillet on the blade surface. However, the static pressure difference from the blade-pressure to the blade-suction surface near the endwall shows that the pressure differences generally decrease with the contour endwall and leading edge fillet. This results in less turning of the endwall region flow from the pressure side to suction side in the passage as indicated by reductions in the flow yaw angles and pitchwise velocity with the structural modifications. The streamwise vorticity of the passage vortex is also affected indicating weaker secondary flows in the endwall region for the contour endwall and leading edge fillet cases. The endwall Nusselt numbers are smaller especially upstream of the throat region for the contour endwall and fillet cases.

Measurements of Secondary Losses in a Turbine Cascade With the Implementation of Nonaxisymmetric Endwall Contouring

2009 ◽  
Vol 132 (1) ◽  
Author(s):  
D. C. Knezevici ◽  
S. A. Sjolander ◽  
T. J. Praisner ◽  
E. Allen-Bradley ◽  
E. A. Grover
Keyword(s):  
Kinetic Energy ◽  
Axial Flow ◽  
Surface Flow ◽  
Secondary Flows ◽  
Test Facility ◽  
Suction Surface ◽  
Blade Passage ◽  
Low Pressure Turbine ◽  
Endwall Contouring ◽  
Pressure And Velocity Measurements

An approach to endwall contouring has been developed with the goal of reducing secondary losses in highly loaded axial flow turbines. The present paper describes an experimental assessment of the performance of the contouring approach implemented in a low-speed linear cascade test facility. The study examines the secondary flows of a cascade composed of Pratt & Whitney PAKB airfoils. This airfoil has been used extensively in low-pressure turbine research, and the present work adds intrapassage pressure and velocity measurements to the existing database. The cascade was tested at design incidence and at an inlet Reynolds number of 126,000 based on inlet midspan velocity and axial chord. Quantitative results include seven-hole pneumatic probe pressure measurements downstream of the cascade to assess blade row losses and detailed seven-hole probe measurements within the blade passage to track the progression of flow structures. Qualitative results take the form of oil surface flow visualization on the endwall and blade suction surface. The application of endwall contouring resulted in lower secondary losses and a reduction in secondary kinetic energy associated with pitchwise flow near the endwall and spanwise flow up the suction surface within the blade passage. The mechanism of loss reduction is discussed in regard to the reduction in secondary kinetic energy.

Experimental Studies of Leading Edge Contouring Influence on Secondary Losses in Transonic Turbines

2012 ◽  
Author(s):  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten Fransson ◽  
Boris I. Mamaev ◽  
Mats Annerfeldt
Keyword(s):  
Gas Turbine ◽  
Guide Vane ◽  
Experimental Studies ◽  
Three Dimensional ◽  
Leading Edge ◽  
Secondary Flows ◽  
Noticeable Effect ◽  
Flow Angle ◽  
Wide Range ◽  
Contour Plots

An experimental study of the hub leading edge contouring using fillets is performed in an annular sector cascade to observe the influence of secondary flows and aerodynamic losses. The investigated vane is a three dimensional gas turbine guide vane (geometrically similar) with a mid-span aspect ratio of 0.46. The measurements are carried out on the leading edge fillet and baseline cases using pneumatic probes. Significant precautions have been taken to increase the accuracy of the measurements. The investigations are performed for a wide range of operating exit Mach numbers from 0.5 to 0.9 at a design inlet flow angle of 90°. Data presented include the loading, fields of total pressures, exit flow angles, radial flow angles, as well as profile and secondary losses. The vane has a small profile loss of approximately 2.5% and secondary loss of about 1.1%. Contour plots of vorticity distributions and velocity vectors indicate there is a small influence of the vortex-structure in endwall regions when the leading edge fillet is used. Compared to the baseline case the loss for the filleted case is lower up to 13% of span and higher from 13% to 20% of the span for a reference condition with Mach no. of 0.9. For the filleted case, there is a small increase of turning up to 15% of the span and then a small decrease up to 35% of the span. Hence, there are no significant influences on the losses and turning for the filleted case. Results lead to the conclusion that one cannot expect a noticeable effect of leading edge contouring on the aerodynamic efficiency for the investigated 1st stage vane of a modern gas turbine.

scholarly journals Suppression of Mixed-Flow Pump Instability and Surge by the Active Alteration of Impeller Secondary Flows

1993 ◽  
Author(s):  
Akira Goto
Keyword(s):  
Flow Characteristic ◽  
Secondary Flows ◽  
Wake Flow ◽  
Streamwise Vorticity ◽  
Jet Injection ◽  
Positive Slope ◽  
Suction Surface ◽  
Mixed Flow ◽  
Pump System ◽  
Flow Pump

An active method for enhancing pump stability, featuring water jet injection at impeller inlet, was applied to a mixed-flow pump. The stall margin, between the design point and the positive slope region of the head-flow characteristic, was most effectively enlarged by injecting the jet in the counter-rotating direction of the impeller. The counter-rotating streamwise vorticity along the casing, generated by the velocity discontinuity due to the jet injection, altered the secondary flow pattern in the impeller by opposing the passage vortex and assisting the tip leakage vortex motion. The location of the wake flow was displaced away from the casing-suction surface corner of the impeller, thus avoiding the onset of the extensive corner separation, the cause of positive slope region of the head-flow characteristic. This method was also confirmed to be effective for stabilizing a pump system already in a state of surge.

Surface Flow Patterns as Visualised by Dust Deposits on the Blades of a Fan

1963 ◽  
Vol 67 (633) ◽  
pp. 589-594 ◽  
Author(s):  
E. T. Hignett ◽  
M. M. Gibson
Keyword(s):  
Boundary Layer ◽  
Wind Tunnel ◽  
Leading Edge ◽  
Surface Flow ◽  
Scale Model ◽  
Flow Visualisation ◽  
Pressure Gradients ◽  
Suction Surface ◽  
Layer Flow ◽  
Dust Deposits

Investigations by one of the authors in connection with the design of a fan for a blower type of wind tunnel showed that regular and repeatable dust patterns occurred on the blades of a one-quarter scale model fan of 18 inches diameter. Dust was deposited on the fan blades along the leading-edge and on the suction surface over an area thought to be the turbulent region of the boundary layer. The introduction of isolated protuberances on the dust free area of a blade gave rise to turbulence wedges in which dust was also deposited and this was interpreted as confirmation of the coincidence of the dust deposits with regions of turbulent boundary-layer flow. These deposits showed the existence of a considerable extent of laminar flow on the suction surface of each blade close to the root, a region where high lift coefficients would be expected with associated adverse pressure gradients. Two-dimensional wind tunnel experiments were made to confirm the interpretation of the observed dust patterns by comparison with the smoke filament and volatile liquid methods of flow visualisation and these are reported in Reference 2.

Compressor Leading Edge Spikes: A New Performance Criterion

2010 ◽  
Vol 133 (2) ◽  
Author(s):  
Martin N. Goodhand ◽  
Robert J. Miller
Keyword(s):  
Leading Edge ◽  
Transition Process ◽  
Performance Criterion ◽  
Original Design ◽  
Suction Surface ◽  
Compressor Blades ◽  
Time Resolved ◽  
Surface Pressure Distribution ◽  
Small Spike ◽  
Profile Loss

Compressor blades often have a small “spike” in the surface pressure distribution at the leading edge. This may result from blade erosion, manufacture defects, or compromises made in the original design process. This paper investigates the effect of these spikes on profile loss, and presents a criterion to ensure they are not detrimental to compressor performance. In the first part of the paper, two geometries of leading edge are tested. One has a small spike, typical of those found on modern compressors; the other has no spike, characteristic of an idealized leading edge. Testing was undertaken on the stator of a single-stage low speed compressor. The time resolved boundary layer was measured using a hot-wire microtraversing system. It is shown that the presence of the small spike changes the time resolved transition process on the suction surface, but that this results in no net increase in loss. In the second part of the paper, spike height is systematically changed using a range of leading edge geometries. It is shown that below a critical spike height, the profile loss is constant. If the critical spike height is exceeded, the leading edge separates and profile loss rises by 30%. Finally, a criterion is developed, based on the total diffusion across the spike. Three different leading edge design philosophies are investigated. It is shown that if the spike diffusion factor is kept below 0.1 over the blade’s incidence range, performance is unaffected by leading edge geometry.

Aerodynamic and Heat Transfer Characterization of a Nozzle Vane Cascade With and Without Platform Cooling

2015 ◽  
Author(s):  
G. Barigozzi ◽  
A. Perdichizzi ◽  
M. Henze ◽  
J. Krueckels
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Cooling System ◽  
Leading Edge ◽  
Surface Flow ◽  
Secondary Flows ◽  
Intensity Level ◽  
Film Cooling Effectiveness ◽  
Nozzle Vane

In the present paper, aerodynamic and thermal performance of a linear nozzle vane cascade is fully assessed. Tests have been carried out with and without platform cooling, with coolant ejected through a slot located upstream of the leading edge. Cooling air is also ejected through a row of cylindrical holes located upstream of the slot, simulating a combustor cooling system. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at variable cooling injection conditions. Aero-thermal characterization of vane platform was obtained through 5-hole probe measurements, oil & dye surface flow visualizations, measurements of end wall adiabatic film cooling effectiveness and heat transfer coefficient. The platform cooling scheme operated at nominal injection rate was shown to effectively reduce the heat load over most of the platform surface, with only a small increase in secondary flows loss. Combustor holes injection resulted beneficial in controlling momentum of coolant approaching the cascade, thus limiting the secondary flows growth and resulting in an increase of the coolant film length inside of the passage.

Performance Metrics and Experimental Testing of Erosion-Resistant Compressor Blade Coatings

2014 ◽  
Author(s):  
Sean G. Leithead ◽  
William D. E. Allan ◽  
Linruo Zhao
Keyword(s):  
Titanium Nitride ◽  
Standardized Testing ◽  
Gas Turbine ◽  
Performance Metrics ◽  
Experimental Testing ◽  
Leading Edge ◽  
Compressor Blade ◽  
Secondary Flows ◽  
Suction Surface ◽  
Durability Test

A durability test rig for erosion-resistant gas turbine engine compressor blade coatings was designed and commissioned. Bare and coated 17-4PH steel modified NACA 6505-profile blades were spun at an average speed of 10 860 rpm and exposed to garnet sand-laden air for 5 hours at an average sand concentration of 2.5gm3of air and a blade leading edge (LE) Mach number of 0.50. The rig was designed to represent a first stage axial compressor. Two 16μm-thick coatings were tested: Titanium Nitride (TiN) and Chromium-Aluminum-Titanium Nitride (CrAlTiN), both applied using an Arc Physical Vapour Deposition technique. A composite scale, defined as the LAZ score, was devised to compare the durability performance of bare and coated blades based on mass-loss and blade dimension changes. The bare blades’ LAZ score was set as a benchmark of 1.00, with the TiN-coated and CrAlTiN-coated blades obtaining respective scores of 0.69 and 0.41. A lower score identified a more erosion-resistant coating. Major locations of blade wear included: trailing edge (TE), LE and rear suction surface (SS). TE thickness was reduced, the LE became blunt, and the rear SS was scrubbed by overtip and recirculation zone vortices. The erosion effects of secondary flows were found to be significant. Erosion damage due to reflected particles was absent due to a low blade solidity of 0.7. The rig is best suited for durability evaluation of erosion-resistant coatings after being proven worthy of consideration for gas turbine engines through ASTM standardized testing.

Sign in / Sign up

Export Citation Format

Share Document