Measurement of Downwash at a Mach Number of 1.45 Behind Two Wings of Finite Span

1951 ◽  
Vol 55 (481) ◽  
pp. 43-51
Author(s):  
W. F. Hilton
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Rear Surface ◽  
Similar Magnitude ◽  
Trailing Edge ◽  
Finite Span ◽  
Finite Wing

Measurements were made of the downwash effects behind two finite wings 3.1 percent, thick, having square and 20° raked tips respectively. The tests were conducted at a Mach number of 1.45 and a Reynolds number of 1.2 millions by traversing a yawmeter 1.62 chords behind the trailing edge of the finite wings.In general, a maximum downwash of the order of ½° per degree of wing incidence was observed in that portion of the tip Mach cone behind the wing, and a maximum upwash of similar magnitude was observed in that part of the tip Mach cone situated outboard of the wing.Thus it is apparent that these effects are large enough to affect the lift on any surface situated in the tip Mach cone behind a finite wing. In particular, placing the rear surface in the downwash region behind a finite wing, will tend to reduce the overall lift while placing it in the upwash region will tend to magnifiy the variations of lift initiated by the finite wing.

The Effect of Pulsed Blowing on the Boundary Layer of a Highly Loaded Low Pressure Turbine Blade

2013 ◽  
Author(s):  
Marion Mack ◽  
Roland Brachmanski ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Mach Number ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
Trailing Edge ◽  
Low Pressure Turbine ◽  
Wide Range ◽  
Highly Loaded

The performance of the low pressure turbine (LPT) can vary appreciably, because this component operates under a wide range of Reynolds numbers. At higher Reynolds numbers, mid and aft loaded profiles have the advantage that transition of suction side boundary layer happens further downstream than at front loaded profiles, resulting in lower profile loss. At lower Reynolds numbers, aft loading of the blade can mean that if a suction side separation exists, it may remain open up to the trailing edge. This is especially the case when blade lift is increased via increased pitch to chord ratio. There is a trend in research towards exploring the effect of coupling boundary layer control with highly loaded turbine blades, in order to maximize performance over the full relevant Reynolds number range. In an earlier work, pulsed blowing with fluidic oscillators was shown to be effective in reducing the extent of the separated flow region and to significantly decrease the profile losses caused by separation over a wide range of Reynolds numbers. These experiments were carried out in the High-Speed Cascade Wind Tunnel of the German Federal Armed Forces University Munich, Germany, which allows to capture the effects of pulsed blowing at engine relevant conditions. The assumed control mechanism was the triggering of boundary layer transition by excitation of the Tollmien-Schlichting waves. The current work aims to gain further insight into the effects of pulsed blowing. It investigates the effect of a highly efficient configuration of pulsed blowing at a frequency of 9.5 kHz on the boundary layer at a Reynolds number of 70000 and exit Mach number of 0.6. The boundary layer profiles were measured at five positions between peak Mach number and the trailing edge with hot wire anemometry and pneumatic probes. Experiments were conducted with and without actuation under steady as well as periodically unsteady inflow conditions. The results show the development of the boundary layer and its interaction with incoming wakes. It is shown that pulsed blowing accelerates transition over the separation bubble and drastically reduces the boundary layer thickness.

An Experimental Study of Turbine Vane Heat Transfer With Water-Air Cooling

1996 ◽  
Author(s):  
Nirm V. Nirmalan ◽  
John A. Weaver ◽  
Larry D. Hylton
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Mach Number ◽  
Mass Flow ◽  
Coolant Temperature ◽  
Air Cooling ◽  
Gas Temperature ◽  
Temperature Ratio ◽  
Trailing Edge ◽  
Exit Mach Number

This paper presents data showing the improvement in cooling effectiveness of turbine vanes through the application of water-air cooling technology in an industrial/utility engine application. The technique utilizes a finely dispersed water-in-air mixture that impinges on the internal surfaces of turbine airfoils to produce very high cooling rates. An airfoil was designed to contain a standard impingement tube which distributes the water-air mixture over the inner surface of the airfoil. The water flash vaporizes off the airfoil inner wall. The resulting mixture of air-steam-water droplets is then routed through a pin fin array in the trailing edge region of the airfoil where additional water is vaporized. The mixture then exits the airfoil into the gas path through trailing edge slots. Experimental measurements were made in a three-vane, linear, two-dimensional cascade. The principal independent parameters — Mach number, Reynolds number, wall-to-gas temperature ratio, and coolant-to-gas mass flow ratio — were maintained over ranges consistent with typical engine conditions. Five impingement tubes were utilized to study geometry scaling, impingement tube-to-airfoil wall gap spacing, impingement tube hole diameter, and impingement tube hole patterns. The test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, gas-to-coolant temperature ratio, water- and air-coolant-to-gas mass flow ratios, and impingement tube geometry. Heat transfer effectiveness data obtained in this program demonstrated that overall cooling levels typical for air cooled vanes could be achieved with the water-air cooling technique with reductions of cooling air flow of significantly more than 50%.

An Experimental Study of Turbine Vane Heat Transfer With Water–Air Cooling

1998 ◽  
Vol 120 (1) ◽  
pp. 50-60 ◽  
Author(s):  
N. V. Nirmalan ◽  
J. A. Weaver ◽  
L. D. Hylton
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Mach Number ◽  
Mass Flow ◽  
Coolant Temperature ◽  
Air Cooling ◽  
Gas Temperature ◽  
Temperature Ratio ◽  
Trailing Edge ◽  
Exit Mach Number

This paper presents data showing the improvement in cooling effectiveness of turbine vanes through the application of water–air cooling technology in an industrial/utility engine application. The technique utilizes a finely dispersed water-in-air mixture that impinges on the internal surfaces of turbine airfoils to produce very high cooling rates. An airfoil was designed to contain a standard impingement tube, which distributes the water–air mixture over the inner surface of the airfoil. The water flash vaporizes off the airfoil inner wall. The resulting mixture of air–steam–water droplets is then routed through a pin fin array in the trailing edge region of the airfoil where additional water is vaporized. The mixture then exits the airfoil into the gas path through trailing edge slots. Experimental measurements were made in a three-vane, linear, two-dimensional cascade. The principal independent parameters—Mach number, Reynolds number, wall-to-gas temperature ratio, and coolant-to-gas mass flow ratio—were maintained over ranges consistent with typical engine conditions. Five impingement tubes were utilized to study geometry scaling, impingement tube-to-airfoil wall gap spacing, impingement tube hole diameter, and impingement tube hole patterns. The test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, gas-to-coolant temperature ratio, water-and air-coolant-to-gas mass flow ratios, and impingement tube geometry. Heat transfer effectiveness data obtained in this program demonstrated that overall cooling levels typical for air-cooled Vanes could be achieved with the water–air cooling technique with reductions of cooling air flow of significantly more than 50 percent.

Morphology of the forced oscillatory flow past a finite-span wing at low Reynolds number

2007 ◽  
Vol 571 ◽  
pp. 327-357 ◽  
Author(s):  
K. PARKER ◽  
K. D. VON ELLENRIEDER ◽  
J. SORIA
Keyword(s):  
Reynolds Number ◽  
Oscillatory Flow ◽  
Three Dimensional ◽  
Wake Flow ◽  
Half Cycle ◽  
Two Dimensional ◽  
Reynolds Stresses ◽  
Trailing Edge ◽  
Finite Span ◽  
Spanwise Vorticity

A study of the morphology of the vortical skeleton behind a flapping NACA0030 wing with a finite aspect ratio of 3, is undertaken. The motivation for this work originates with the proposal that thrust can be efficiently produced by flapping aerofoils. The test condition corresponds to a Strouhal number of 0.35, Reynolds number, based on aerofoil chord, of 600 and an amplitude of flapping, equal to the chord length of the wing. This test condition corresponds to the optimal thrust-producing case in infinite-span flapping wings. This study investigates the effect of wing three-dimensionality on the structure of the wake-flow. This is accomplished here, by quantitatively describing the spatio-temporal variations in the velocity, vorticity and Reynolds stresses for the finite-span-wing case.Preliminary flow visualizations suggest that the presence of wingtip vortices for the three-dimensional-wing case, create a different vortical structure to the two-dimensional-wing case. In the case of a two-dimensional-wing, the flow is characterized by the interaction of leading- and trailing-edge vorticity, resulting in the formation of a clear reverse Kármán vortex street at the selected test condition. In the case of a three-dimensional-wing, the flow exhibits a high degree of complexity and three-dimensionality, particularly in the midspan region. Using phase-averaged particle image velocimetry measurements of the forced oscillatory flow, a quantitative analysis in the plane of symmetry of the flapping aerofoil was undertaken. Using a triple decomposition of the measured velocities, the morphological characteristics of the spanwise vorticity is found to be phase correlated with the aerofoil kinematics. Reynolds stresses in the direction of oscillation are the dominant dissipative mechanism. The mean velocity profiles resemble a jet, indicative of thrust production. Pairs of strong counter-rotating vortices from the leading- and trailing-edge of the aerofoil are shed into the flow at each half-cycle. The large-scale structure of the flow is characterized by constructive merging of spanwise vorticity. The midspan region is populated by cross-sections of interconnected vortex rings.

scholarly journals Simulation of Gas Flow Over a Micro Cylinder

2009 ◽  
Author(s):  
Yuan Hu ◽  
Quanhua Sun ◽  
Jing Fan
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Drag Coefficient ◽  
Gas Flow ◽  
Rarefied Gas ◽  
Low Reynolds Number ◽  
Pressure Drag ◽  
Low Reynolds ◽  
Rarefied Gas Effect ◽  
Gas Effect

Gas flow over a micro cylinder is simulated using both a compressible Navier-Stokes solver and a hybrid continuum/particle approach. The micro cylinder flow has low Reynolds number because of the small length scale and the low speed, which also indicates that the rarefied gas effect exists in the flow. A cylinder having a diameter of 20 microns is simulated under several flow conditions where the Reynolds number ranges from 2 to 50 and the Mach number varies from 0.1 to 0.8. It is found that the low Reynolds number flow can be compressible even when the Mach number is less than 0.3, and the drag coefficient of the cylinder increases when the Reynolds number decreases. The compressible effect will increase the pressure drag coefficient although the friction coefficient remains nearly unchanged. The rarefied gas effect will reduce both the friction and pressure drag coefficients, and the vortex in the flow may be shrunk or even disappear.

Auto and Cross Correlation Measurements in a Turbine Cascade Using a Digital Correlator

1982 ◽  
Author(s):  
P. J. Bryanston-Cross ◽  
J. J. Camus
Keyword(s):  
Mach Number ◽  
High Speed ◽  
Cross Correlation ◽  
Turbine Blades ◽  
Image Plane ◽  
Strongly Correlated ◽  
Trailing Edge ◽  
Number Range ◽  
Cross Correlations ◽  
Digital Correlator

A simple technique has been developed which samples the dynamic image plane information of a schlieren system using a digital correlator. Measurements have been made in the passages and in the wakes of transonic turbine blades in a linear cascade. The wind tunnel runs continuously and has independently variable Reynolds and Mach number. As expected, strongly correlated vortices were found in the wake and trailing edge region at 50 KHz. Although these are strongly coherent we show that there is only limited cross-correlation from wake to wake over a Mach no. range M = 0.5 to 1.25 and variation of Reynolds number from 3 × 105 to 106. The trailing edge fluctuation cross correlations were extended both upstream and downstream and preliminary measurements indicate that this technique can be used to obtain information on wake velocity. The vortex frequency has also been measured over the same Mach number range for two different cascades. The results have been compared with high speed schlieren photographs.

Investigation of the 3D Unsteady Rotor Pressure Field in a HP Turbine Stage

2002 ◽  
Author(s):  
E. Valenti ◽  
J. Halama ◽  
R. De´nos ◽  
T. Arts
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Pressure Field ◽  
Pressure Ratio ◽  
Turbine Stage ◽  
Time Resolved ◽  
Mach Number Distribution ◽  
Rotor Surface ◽  
Exit Mach Number ◽  
Unsteady Pressure Measurements

This paper presents steady and unsteady pressure measurements at three span locations (15, 50 and 85%) on the rotor surface of a transonic turbine stage. The data are compared with the results of a 3D unsteady Euler stage calculation. The overall agreement between the measurements and the prediction is satisfactory. The effects of pressure ratio and Reynolds number are discussed. The rotor time-averaged Mach number distribution is very sensitive to the pressure ratio of the stage since the incidence of the flow changes as well as the rotor exit Mach number. The time-resolved pressure field is dominated by the vane trailing edge shock waves. The incidence and intensity of the shock strongly varies from hub to tip due to the radial equilibrium of the flow at the vane exit. The decrease of the pressure ratio attenuates significantly the amplitude of the fluctuations. An increase of the pressure ratio has less significant effect since the change in the vane exit Mach number is small. The effect of the Reynolds number is weak for both the time-averaged and the time-resolved rotor static pressure at mid-span, while it causes an increase of the pressure amplitudes at the two other spans.

Computational Study of the Risø-B1-18 Airfoil with a Hinged Flap Providing Variable Trailing Edge Geometry

2005 ◽  
Vol 29 (2) ◽  
pp. 89-113 ◽  
Author(s):  
Niels Troldborg
Keyword(s):  
Reynolds Number ◽  
Unsteady Flow ◽  
Computational Study ◽  
Turbine Blades ◽  
Aerodynamic Characteristics ◽  
Flow Conditions ◽  
Wind Turbine Blades ◽  
Trailing Edge ◽  
Edge Geometry ◽  
Fully Developed Turbulent Flow

A comprehensive computational study, in both steady and unsteady flow conditions, has been carried out to investigate the aerodynamic characteristics of the Risø-B1-18 airfoil equipped with variable trailing edge geometry as produced by a hinged flap. The function of such flaps should be to decrease fatigue-inducing oscillations on the blades. The computations were conducted using a 2D incompressible RANS solver with a k-w turbulence model under the assumption of a fully developed turbulent flow. The investigations were conducted at a Reynolds number of Re = 1.6 · 106. Calculations conducted on the baseline airfoil showed excellent agreement with measurements on the same airfoil with the same specified conditions. Furthermore, a more widespread comparison with an advanced potential theory code is presented. The influence of various key parameters, such as flap shape, flap size and oscillating frequencies, was investigated so that an optimum design can be suggested for application with wind turbine blades. It is concluded that a moderately curved flap with flap chord to airfoil curve ratio between 0.05 and 0.10 would be an optimum choice.

scholarly journals Experimental Investigation of Tandem Blade Cascades With Double-Circular ARC Profiles

1985 ◽  
Author(s):  
Wu Guochuan ◽  
Zhuang Biaonan ◽  
Guo Bingheng
Keyword(s):  
Experimental Investigation ◽  
Reynolds Number ◽  
Mach Number ◽  
Total Pressure ◽  
Incidence Angle ◽  
Axial Direction ◽  
Geometrical Parameters ◽  
Optimal Conditions ◽  
Turning Angle ◽  
Circular Arc

24 double circular are tandem blade cascades of three different chord-ratios were investigated under different displacements in peripheral and axial direction. The inlet Mach number was 0.3. The Reynolds number based on blade chord was 2.7×105. The characteristics of the tandem blade cascades, such as the dependence of turning angle and coefficient of total pressure loss on incidence angle were obtained. The ranges of main geometrical parameters under optimal conditions were recommended.

Viscous tails in Hele-Shaw flow

1971 ◽  
Vol 46 (3) ◽  
pp. 569-576
Author(s):  
C. J. Wood
Keyword(s):  
Reynolds Number ◽  
Trailing Edge ◽  
Kutta Condition ◽  
Quantitative Discrepancy ◽  
Edge Flow ◽  
Hele Shaw Cell

An experiment has been performed, using pulsed dye injection on an aerofoil in a Hele-Shaw cell. The purpose was to observe the form of the trailing-edge flow when the Reynolds number was high enough to permit separation and the initiation of a Kutta condition. The experiment provides a successful confirmation of the existence of a ‘viscous tail’ as predicted by Buckmaster (1970) although there is an unexplained quantitative discrepancy.

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