scholarly journals Modeling of Cavitating Flow through Waterjet Propulsors

2012 ◽  
Vol 2012 ◽  
pp. 1-13 ◽  
Author(s):  
Jules W. Lindau ◽  
Christopher Pena ◽  
Warren J. Baker ◽  
James J. Dreyer ◽  
William L. Moody ◽  
...  
Keyword(s):  
Fluid Dynamics ◽  
Total Pressure ◽  
Axial Flow ◽  
Flow Coefficient ◽  
Experimental Measurements ◽  
Computational Results ◽  
Suction Surface ◽  
Blade Passage ◽  
Wide Range ◽  
Flow Through

A computational-fluid-dynamics-based modeling effort to capture flow through an axial flow waterjet propulsor is presented. The effort covered the waterjet flow over a wide range of flow coefficients and into cavitation-driven breakdown. The computations are presented in cavitation at two values of flow coefficient through a series of decreasing operating inlet total pressure. The computational results are compared to experimental measurements. Suction-surface and tip-gap cavitation patterns are presented and compared to experimental photographs. Presented computational solutions are blade-passage steady and periodic. The computational results apply a powering iteration methodology to facilitate coupling of rotor, stator, and inflow and outflow ducting.

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.

Performance Study of a Mixed Flow Impeller Covering an Unsteady Flow Field

1992 ◽  
Vol 206 (2) ◽  
pp. 83-93 ◽  
Author(s):  
S Sarkar
Keyword(s):  
Axial Flow ◽  
Rotating Stall ◽  
Operating Conditions ◽  
Performance Study ◽  
Mixed Flow ◽  
Pump Impeller ◽  
Wide Range ◽  
Reversal Flow ◽  
Flow Through ◽  
Flow Pump

The results presented here are part of a detailed programme measuring the aerodynamics of a high specific speed mixed flow pump impeller over a wide range of operating conditions, including its behaviour in the unsteady stalled regime. The aim is to elucidate the physics of the flow through such an impeller. The noticeable features are the formation of part-span rotating stall cells having no periodicity and organized structure at reduced flow and also the shifting positions of reversal flow pockets as the flowrate changes. Measurements of loss and its variation with span-wise positions and flowrates enable the variation of local efficiency to be determined. The overall flow picture is similar to that expected in an axial flow impeller, though the present impeller displays a narrow stall hysteresis loop almost right through its operating range.

A New Physics Based Unsteady Transition Model Using the Universal Intermittency Function

2019 ◽  
Author(s):  
Ali Nikparto ◽  
Meinhard T. Schobeiri
Keyword(s):  
Heat Transfer ◽  
Transport Equation ◽  
Turbine Blades ◽  
Turbulence Models ◽  
New Physics ◽  
Intermittent Flow ◽  
Experimental Measurements ◽  
Computational Results ◽  
Suction Surface ◽  
Turbine Engine

Abstract The flow inside a gas turbine engine has unique complexities. One of the important characteristics of such flow field is the existence of periodic unsteady wakes, originating from stator–rotor interaction. The unsteady wakes, with their highly vortical core, impinge on the downstream blade surfaces and cause an intermittent transition of the boundary layer from laminar to turbulent. The relative intermittency value corresponding to the wake vortical core and the calm region outside the wake, irrespective of freestream turbulence intensity and wake frequency, exhibits a universal behavior which is best described by the universal intermittency function of Chakka and Schobeiri [1, 2]. This study aims at introducing a new physics-based universal intermittency function that in conjunction with the current turbulence models accurately predicts the unsteady behavior of an intermittent flow. For that reason, a transport equation for turbulence intermittency was proposed based on this function and was integrated into a RANS based solver with k-ω turbulence model. The model was tested for reliability. Experimental aerodynamics and heat transfer measurements conducted at Turbomachinery Performance and Flow research Lab (TPFL) at Texas A&M University, were used as benchmark tests. For experimental measurements, an unsteady linear cascade facility in TPFL was used to produce the periodic unsteady flow condition. Moving wakes, originating from upstream blades, were simulated in this facility by rods attached to two parallel timing belts in front of the turbine blades. Heat transfer measurements along the suction surface were conducted utilizing a specially manufactured blade with an internal heater core, instrumented with liquid crystal. All Measurements and calculations were conducted at Reynolds number of 264,000. The computational results, obtained from implementing the new enhanced intermittency transport equation into the solver, are compared with (a) experimental measurements and (b) with the computational results from RANS that incorporates Langtry-Menter [3, 4] method.

Visualization Study of Flow in Axial Flow Inducer

1972 ◽  
Vol 94 (4) ◽  
pp. 777-787 ◽  
Author(s):  
B. Lakshminarayana
Keyword(s):  
Radial Velocity ◽  
Secondary Flow ◽  
Three Dimensional ◽  
Axial Flow ◽  
Flow Theory ◽  
Secondary Flows ◽  
Flow Coefficient ◽  
Flow Through

A visualization study of the flow through a three ft dia model of a four bladed inducer, which is operated in air at a flow coefficient of 0.065, is reported in this paper. The flow near the blade surfaces, inside the rotating passages, downstream and upstream of the inducer is visualized by means of smoke, tufts, ammonia filament, and lampblack techniques. Flow is found to be highly three dimensional, with appreciable radial velocity throughout the entire passage. The secondary flows observed near the hub and annulus walls agree with qualitative predictions obtained from the inviscid secondary flow theory. Based on these investigations, methods of modeling the flow are discussed.

The Use of Circumferentially Varying Stagger Guide Vanes in an Axial Flow Pump or Compressor

1990 ◽  
Vol 112 (2) ◽  
pp. 294-297 ◽  
Author(s):  
J. H. Horlock
Keyword(s):  
Total Pressure ◽  
Guide Vane ◽  
Axial Flow ◽  
Guide Vanes ◽  
Actuator Disk ◽  
Outlet Angle ◽  
Flow Through ◽  
Axial Flow Pump ◽  
Flow Pump

An actuator disk analysis is given of the flow through a guide vane and rotor combination. It is shown that changes in total pressure across the rotor are in general related to circumferential variations in guide vane outlet angle. In particular known variations in inlet total pressure may be eliminated by suitable circumferential changes in guide vane stagger.

Effects of Bleed Rate and Endwall Location on the Aerodynamic Behavior of a Circular Hole Bleed Off-Take

2006 ◽  
Vol 129 (4) ◽  
pp. 645-658 ◽  
Author(s):  
B. A. Leishman ◽  
N. A. Cumpsty ◽  
J. D. Denton
Keyword(s):  
Circular Hole ◽  
Unstructured Mesh ◽  
Three Dimensional ◽  
Air Pressure ◽  
Suction Surface ◽  
Primary Flow ◽  
Jet Engine ◽  
Blade Passage ◽  
Pressure Surface ◽  
Flow Through

In a jet engine bleed off-takes on the hub and casing endwalls, part way through the compressor, supply high-pressure air for cooling, sealing, de-icing, and aircraft cabin air applications; bleed also assists compressor operation at part-speed conditions. Two separate issues are of interest: the bleed off-take air pressure and the interaction of the bleed off-take with the primary flow through the blade passage. In this paper, the aerodynamic behavior is presented for a circular-hole bleed off-take at three endwall locations within a stationary cascade blade passage: at midpassage; near the blade pressure surface; and near the blade suction surface. Results from low-speed cascade experiments are complemented by three-dimensional numerical calculations using an unstructured mesh-based solver, in which the blade passage and bleed off-take geometry are fully modeled. The bleed off-take location and the magnitude of bleed rate influence the off-take air pressure and the interaction with the primary passage flow. For optimum design at zero and low bleed rates, off-takes near the blade pressure surface give the highest bleed air pressures and minimum loss in the blade passage. For minimum blade passage loss at higher bleed rates, however, it is necessary to take bleed near the blade suction surface. The paper discusses the causes for this pattern of behavior.

Wake–Boundary Layer Interactions in an Axial Flow Turbine Rotor at Off-Design Conditions

1989 ◽  
Vol 111 (2) ◽  
pp. 181-192 ◽  
Author(s):  
H. P. Hodson ◽  
J. S. Addison
Keyword(s):  
Rotor Blade ◽  
Axial Flow ◽  
Leading Edge ◽  
Surface Flow ◽  
Design Flow ◽  
Flow Coefficient ◽  
Experimental Investigations ◽  
Pressure Side ◽  
Suction Surface ◽  
Edge Separation

A series of experimental investigations has been undertaken in a single-stage low-speed turbine. The measurements involved rotor blade surface flow visualization, surface-mounted hot-film anemometry, and exit pitot traverses. The effects of varying the flow coefficient and Reynolds number upon the performance of the rotor blade at midspan are described. At the design flow coefficient (φ = 0.495), the rotor pressure surface flow may be regarded as laminar, while on the suction surface, laminar flow gives way to unsteady stator wake-induced transition and then to turbulent flow. Over the range of Reynolds numbers investigated (1.8×105–3.3×105), the rotor midspan performance is dominated by the suction surface transition process; suction surface separation is prevented and the rotor midspan loss coefficient remains approximately constant throughout the range. At positive incidence, suction surface leading edge separation and transition are caused by a velocity overspeed. Reattachment occurs as the flow begins to accelerate toward the throat. The loss associated with the separation becomes significant with increasing incidence. At negative incidence, a velocity overspeed causes leading edge separation of the pressure side boundary layers. Reattachment generally occurs without full transition. The suction surface flow is virtually unaffected. Therefore, the rotor midspan profile loss remains unchanged from the zero incidence value until pressure side stall occurs.

A Numerical Investigation of Loss Coefficient Variation in Various Incidence Angles in Tandem Blades Cascade

2014 ◽  
Author(s):  
Arash Soltani Dehkharqani ◽  
Masoud Boroomand ◽  
Hamzeh Eshraghi
Keyword(s):  
Pressure Ratio ◽  
Axial Flow ◽  
Suction Side ◽  
Loss Coefficient ◽  
Flow Coefficient ◽  
Flow Angle ◽  
Multi Stage ◽  
Wide Range ◽  
And Performance ◽  
Flow Compressor

There is a severe tendency to reduce weight and increase power of gas turbine. Such a requirement is fulfilled by higher pressure ratio of compressor stages. Employing tandem blades in multi-stage axial flow compressors is a promising methodology to control separation on suction sides of blades and simultaneously implement higher turning angle to achieve higher pressure ratio. The present study takes into account the high flow deflection capabilities of the tandem blades consisting of NACA-65 airfoil with fixed percent pitch and axial overlap at various flow incidence angles. In this regard, a two-dimensional cascade model of tandem blades is constructed in a numerical environment. The inlet flow angle is varied in a wide range and overall loss coefficient and deviation angles are computed. Moreover, the flow phenomena between the blades and performance of both forward and afterward blades are investigated. At the end, the aerodynamic flow coefficient of tandem blades are also computed with equivalent single blades to evaluate the performance of such blades in both design and off-design domain of operations. The results show that tandem blades are quite capable of providing higher deflection with lower loss in a wide range of operation and the base profile can be successfully used in design of axial flow compressor. In comparison to equivalent single blades, tandem blades have less dissipation because the momentum exerted on suction side of tandem blades confines the size of separation zone near trailing edges of blades.

Effects of Bleed Rate and Endwall Location on the Aerodynamic Behaviour of a Circular Hole Bleed Off-Take

2004 ◽  
Author(s):  
B. A. Leishman ◽  
N. A. Cumpsty ◽  
J. D. Denton
Keyword(s):  
Circular Hole ◽  
Unstructured Mesh ◽  
Three Dimensional ◽  
Air Pressure ◽  
Suction Surface ◽  
Primary Flow ◽  
Jet Engine ◽  
Blade Passage ◽  
Pressure Surface ◽  
Flow Through

In a jet engine bleed off-takes on the hub and casing endwalls, part-way through the compressor, supply high-pressure air for cooling, sealing, de-icing and aircraft cabin air applications; bleed also assists compressor operation at part-speed conditions. Two separate issues are of interest: the bleed off-take air pressure and the interaction of the bleed off-take with the primary flow through the blade passage. In this paper, the aerodynamic behaviour is presented for a circular-hole bleed off-take at three endwall locations within a stationary cascade blade passage: at mid-passage; near the blade pressure-surface; and near the blade suction-surface. Results from low-speed cascade experiments are complemented by three-dimensional numerical calculations using an unstructured mesh-based solver, in which the blade passage and bleed off-take geometry are fully modelled. The bleed off-take location and the magnitude of bleed rate influence the off-take air pressure and the interaction with the primary passage flow. For optimum design at zero and low bleed rates, off-takes near the blade pressure-surface give the highest bleed air pressures and minimum loss in the blade passage. For minimum blade passage loss at higher bleed rates, however, it is necessary to take bleed near the blade suction-surface. The paper discusses the causes for this pattern of behaviour.

An Experimental Investigation of Stator Clocking Effects in a Two-Stage Low-Speed Axial Compressor

2011 ◽  
Author(s):  
J. Sta¨ding ◽  
D. Wulff ◽  
G. Kosyna ◽  
B. Becker ◽  
V. Gu¨mmer
Keyword(s):  
Total Pressure ◽  
Axial Compressor ◽  
Axial Flow ◽  
Leading Edge ◽  
Global Scale ◽  
Flow Coefficient ◽  
Maximum Efficiency ◽  
Pressure Measurements ◽  
Low Speed ◽  
The Impact

The impact of stator clocking on performance and flow of a 2.5-stage axial compressor has been investigated. Stator clocking, the circumferential indexing of adjacent stator rows with equal blade counts, is known as a potential means to modify the flow field in multistage turbomachinery and increase overall efficiencies of both turbines and compressors. These potential effects on turbomachine performance are due to wake-airfoil interactions and primarily depend on the alignment of the downstream stator row with the upstream stator wake path. The present survey describes and discusses the experimental research on stator clocking effects in a low-speed 2.5-stage axial flow compressor, using front loaded CDA blade sections and cantilevered stator rows with identical blade counts. Conventional static pressure tappings were used to locate global peaks in compressor performance for varying Stator 2 clocking positions at different flow coefficients. Results of unsteady total pressure measurements obtained by means of a high-frequency pressure transducer, embedded in the Stator 2 leading edge, give information on Stator 1 wake propagation. Traverse data from pneumatic 5-hole probes show the impact of stator indexing on Stator 2 exit total pressure at different blade spans. Regardless of flow coefficient, the variations of overall compressor efficiency due to Stator 2 clocking are around 0.2% and are exhibiting a near-sinusoidal trend over the clocking angle. It is shown that total pressure measurements at mid-span of Stator 2 leading edge suggest best overall performances for design and low loading conditions, if the Stator 1 wakes pass through mid-passage of Stator 2. At high loading, however, maximum efficiency locates the wake path directly at the leading edge. Due to a considerable span-wise skewness of the upstream stator wake, the aerodynamic clocking position for Stator 2 varies from hub to tip. While it is shown again that this effect weakens the advantages of airfoil indexing on a global scale, stator clocking shows much more potential if only a single blade section is considered.

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