Experimental results for surface pressures on cone-derived waveriders in the Mach-number range 3 to 5

1982 ◽  
Author(s):  
M. JISCHKE ◽  
M. RASMUSSEN ◽  
D. DANIEL
Keyword(s):  
Mach Number ◽  
Experimental Results ◽  
Number Range

Aerodynamic Design and Testing of Three Low Solidity Steam Turbine Nozzle Cascades

2004 ◽  
Author(s):  
Bo Song ◽  
Wing F. Ng ◽  
Joseph A. Cotroneo ◽  
Douglas C. Hofer ◽  
Gunnar Siden
Keyword(s):  
Mach Number ◽  
Steam Turbine ◽  
Boundary Layer Transition ◽  
Experimental Results ◽  
Blade Surface ◽  
Layer Transition ◽  
Number Range ◽  
Baseline Design ◽  
Wide Range ◽  
Mach Numbers

Three sets of low solidity steam turbine nozzle cascades were designed and tested. The objective was to reduce cost through a reduction in parts count while maintaining or improving performance. The primary application is for steam turbine high pressure sections where Mach numbers are subsonic and high levels of unguided turning can be tolerated. The baseline Design A has a ratio of pitch to axial chord of 1.2. This is the pitch diameter section of a 50% reaction stage that has been verified by multistage testing on steam to have a high level of efficiency. Designs B and C have ratios of pitch to axial chord of 1.5 and 1.8 respectively. All three designs satisfy the same inlet and exit vector diagrams. Analytical surface Mach number distributions and boundary layer transition predictions are presented. Extensive cascade test measurements were carried out for a broad incidence range from −60 to +35 degrees. At each incidence, four outlet Mach numbers were tested, ranging from 0.2 to 0.8, with the corresponding Reynolds number variation from 1.8×105 to 9.0×105. Experimental results of loss coefficient and blade surface Mach number are presented and compared for the three cascades. The experimental results have demonstrated low losses over the tested Mach number range for a wide range of incidence from −45 to 15 degrees. Designs B and C have lower profile losses than Design A. The associated flow physics is interpreted using the results of wake profile, blade surface Mach number distribution and blade surface oil flow visualization, with the emphasis placed on the loss mechanisms for different flow conditions and the loss reduction mechanism with lower solidity. The effect of the higher profile loading of the lower solidity designs on increased end wall losses induced by increased secondary flow, especially on low aspect ratio designs, is the subject of ongoing studies.

Aerodynamic Design and Testing of Three Low Solidity Steam Turbine Nozzle Cascades

2004 ◽  
Vol 129 (1) ◽  
pp. 62-71 ◽  
Author(s):  
Bo Song ◽  
Wing F. Ng ◽  
Joseph A. Cotroneo ◽  
Douglas C. Hofer ◽  
Gunnar Siden
Keyword(s):  
Mach Number ◽  
Steam Turbine ◽  
Boundary Layer Transition ◽  
Experimental Results ◽  
Base Line ◽  
Blade Surface ◽  
Number Range ◽  
Wide Range ◽  
Mach Numbers ◽  
Line Design

Three sets of low solidity steam turbine nozzle cascades were designed and tested. The objective was to reduce cost through a reduction in parts count while maintaining or improving performance. The primary application is for steam turbine high pressure sections where Mach numbers are subsonic and high levels of unguided turning can be tolerated. The base line design A has a ratio of pitch to axial chord of 1.2. This is the pitch diameter section of a 50% reaction stage that has been verified by multistage testing on steam to have a high level of efficiency. Designs B and C have ratios of pitch to axial chord of 1.5 and 1.8, respectively. All three designs satisfy the same inlet and exit vector diagrams. Analytical surface Mach number distributions and boundary layer transition predictions are presented. Extensive cascade test measurements were carried out for a broad incidence range from −60to+35deg. At each incidence, four outlet Mach numbers were tested, ranging from 0.2 to 0.8, with the corresponding Reynolds number variation from 1.8×105 to 9.0×105. Experimental results of loss coefficient and blade surface Mach number are presented and compared for the three cascades. The experimental results have demonstrated low losses over the tested Mach number range for a wide range of incidence from −45to15deg. Designs B and C have lower profile losses than design A. The associated flow physics is interpreted using the results of wake profile, blade surface Mach number distribution, and blade surface oil flow visualization, with the emphasis placed on the loss mechanisms for different flow conditions and the loss reduction mechanism with lower solidity. The effect of the higher profile loading of the lower solidity designs on increased end wall losses induced by increased secondary flow, especially on low aspect ratio designs, is the subject of ongoing studies.

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.

Performance and Boundary Layer Development of a High Turning Compressor Cascade at Sub- and Supercritical Flow Conditions

2012 ◽  
Author(s):  
Christoph Bode ◽  
Dragan Kožulović ◽  
Udo Stark ◽  
Heinz Hoheisel
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Density Ratio ◽  
Boundary Layer Separation ◽  
Experimental Results ◽  
Numerical Code ◽  
Compressor Cascade ◽  
Boundary Layer Development ◽  
Boundary Layer Interaction ◽  
Inlet Angle

Based on current numerical investigations, the present paper reports on new Q2D midspan-calculations and results for the well known high turning (Δβ = 50°) supercritical (Ma1 = 0.85) compressor cascade V2. A Q2D treatment of the problem was chosen in order to avoid the difficult modelling of the porous endwalls in a corresponding 3D approach. All simulations were done with the RANS solver TRACE of the DLR Cologne in combination with modified versions of the Wilcox turbulence model and Langtry/Menter transition model. Existing experimental Q2D midspan-results for the V2 compressor cascade were used to demonstrate the improved ability of the numerical code to determine performance characteristics, blade pressure and Mach number distributions as well as boundary layer parameter and velocity distributions. The loss characteristics show minimum loss regions when plotted against inlet angle or axial velocity density ratio. Within these regions, increasing with decreasing Mach number, the experimental results were adequately predicted. Outside these regions it turned out difficult to reproduce the experimental results due to increasing boundary layer separation. Furthermore, the prediction quality was very good for subsonic conditions (Ma1 = 0.60) and still reasonable for supercritical conditions (Ma1 = 0.85), where shock/boundary layer interaction made the prediction more difficult.

Fabrication and performance evaluation of two multi-layer passive micromixers

Sensor Review ◽  
2018 ◽  
Vol 38 (3) ◽  
pp. 321-325 ◽  
Author(s):  
Xueye Chen ◽  
Jienan Shen ◽  
Zengliang Hu
Keyword(s):  
Reynolds Number ◽  
Performance Evaluation ◽  
Design Methodology ◽  
Experimental Results ◽  
Microfluidic Chips ◽  
Mixing Index ◽  
Number Range ◽  
Content Type ◽  
Chaotic Convection ◽  
And Performance

PurposeThe purpose of this study is to provide a micromixer for achieving effective mixing of two liquids. The mixing of two liquids is difficult to achieve in microfluidic chips because they cannot form turbulence at small dimensions and velocities.Design/methodology/approachIn this paper, four kinds of passive micromixers based on splitting–recombination and chaotic convection are compared. First, a better E-shape mixing unit based on the previous F-shape mixing unit has been designed. Then, the E-shape mixing units are further combined to form three micromixers (i.e. E-mixer, SESM and FESM).FindingsFinally, the mixing experimental results show that the mixing indexes of E-mixer, SESM and FESM are more than those of F-mixer when the Reynolds number range is from 0.5 to 100. And at Re = 15, the lowest mixing index of E-mixer is 71%, which is the highest of the four micromixers.Originality/valueAt Re = 80, the highest mixing index of F-mixer and E-mixer is 92 and 94 per cent, respectively, and then it begins to decrease. But the mixing index of SESM and FESM remains close to 100 per cent.

Equilibrium Thermodynamic Properties Behind Strong Shocks in Helium

1968 ◽  
Vol 72 (686) ◽  
pp. 155-159
Author(s):  
M. Lalor ◽  
H. Daneshyar
Keyword(s):  
Mach Number ◽  
Partition Function ◽  
Shock Waves ◽  
Thermodynamic Properties ◽  
Strong Shock ◽  
Equilibrium Thermodynamic ◽  
Ionized Gas ◽  
Number Range ◽  
Shock Mach Number ◽  
Strong Shock Waves

Summary Tables of equilibrium thermodynamic properties of the ionized gas formed behind strong shock waves in Helium are presented, in the Mach number range 10 to 30, for initial pressures of 0-1, 0-5, 1, 5, 10, 50, 100 torr. The effect of the inclusion of the full partition function series is demonstrated in the Mach number range 20 to 30. A numerical solution has been developed such that the only experimental quantities required for its use are the shock Mach number and the pre-shock conditions.

Experimental Results of Flow Nozzle Based on PTC 6 for High Reynolds Number

2014 ◽  
Author(s):  
Noriyuki Furuichi ◽  
Kar-Hooi Cheong ◽  
Yoshiya Terao ◽  
Shinichi Nakao ◽  
Keiji Fujita ◽  
...  
Keyword(s):  
Reynolds Number ◽  
High Reynolds Number ◽  
Discharge Coefficient ◽  
Experimental Results ◽  
Flow Conditions ◽  
Number Range ◽  
Discharge Coefficients ◽  
Reynolds Number Range ◽  
Lower Reynolds Number ◽  
High Reynolds

Discharge coefficients for three flow nozzles based on ASME PTC 6 are measured under many flow conditions at AIST, NMIJ and PTB. The uncertainty of the measurements is from 0.04% to 0.1% and the Reynolds number range is from 1.3×105 to 1.4×107. The discharge coefficients obtained by these experiments is not exactly consistent to one given by PTC 6 for all examined Reynolds number range. The discharge coefficient is influenced by the size of tap diameter even if at the lower Reynolds number region. Experimental results for the tap of 5 mm and 6 mm diameter do not satisfy the requirements based on the validation procedures and the criteria given by PTC 6. The limit of the size of tap diameter determined in PTC 6 is inconsistent with the validation check procedures of the calibration result. An enhanced methodology including the term of the tap diameter is recommended. Otherwise, it is recommended that the calibration test should be performed at as high Reynolds number as possible and the size of tap diameter is desirable to be as small as possible to obtain the discharge coefficient with high accuracy.

scholarly journals Shock-induced separated flows on the lee surface of delta wings

1987 ◽  
Vol 91 (903) ◽  
pp. 128-141 ◽  
Author(s):  
S. N. Seshadri ◽  
K. Y. Narayan
Keyword(s):  
Pressure Rise ◽  
Leading Edge ◽  
Experimental Results ◽  
Separated Flows ◽  
Delta Wings ◽  
Number Range ◽  
Pressure Distributions ◽  
Boundary Layer Interaction ◽  
Oil Flow ◽  
Attached Flow

Experiments were conducted to study shock-induced separated flows on the lee surface of delta wings with sharp leading edge at supersonic speeds. Two sets of delta wings of different thickness (10° and 25° normal angle), each with leading edge sweep angles varying from 45° to 70°, were tested. The measurements, carried out in a Mach number range from 1.4 to 3.0, included oil flow visualisations (on both sets of wings) and static pressure distributions (on the thicker wings only). Using the test results, some features of shock-induced separated flows, including in particular the boundary between this type of flow and fully attached flow, have been determined. The experimental results indicate that this boundary does not seem to show any significant dependence on wing thickness within the limit of thicknesses tested. It is shown that this boundary can be predicted for thin delta wings using a well known criterion for incipient separation in a glancing shock wave boundary layer interaction, namely that a pressure rise of 1.5 is required across the shock. Comparison of the predicted boundary with experimental results (from oil flow visualisations) shows good agreement.

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