Transition detection for laminar flow aircraft using microphones beneath the surface of laser drilled suction panels

2000 ◽  
Vol 214 (3) ◽  
pp. 143-155
Author(s):  
R V Barrett
Keyword(s):  
Boundary Layer ◽  
Wind Tunnel ◽  
Laminar Flow ◽  
Reynolds Numbers ◽  
Suction Surface ◽  
Microphone Signal ◽  
High Background ◽  
Low Reynolds Numbers ◽  
High Background Noise ◽  
Suction Flow

The possibility of detecting transition through the very small laser drilled perforations in panels representing the suction surface of a hybrid laminar flow aircraft is examined. The method uses miniature microphones to detect changes to the noise received from the boundary layer. Tests using a flat plate rig in a low-turbulence wind tunnel at Reynolds numbers up to 3.8 million per metre, demonstrate that the boundary layer state can be defined in this manner, most simply through measurement of the root mean square (r.m.s.) of the microphone signal. It is shown that the r.m.s. reaches a peak in the transition zone and that when the boundary layer is fully turbulent the value is still significantly higher than it was before transition. Porosity in the range 0.8-6.4 percent was examined, with nominal hole diameters of 0.06 and 0.10 mm in 0.9 mm thick laser drilled suction surface specimens. Suction flow through the surface was found not adversely to affect the operation of the system. The experiment was limited to low Reynolds numbers because the high background noise in the wind tunnel made detection of the boundary layer element of the signal increasingly difficult to define as speed increased. It is considered that test in flight will be needed to prove fully the validity of the method. A preliminary design of an installation for this purpose is suggested that allows the suction flow to be maintained over the measuring region.

A Low Reynolds Number Experimental Evaluation of Tubercles on a Low-Pressure Turbine Cascade

2019 ◽  
Author(s):  
Stephen A. Pym ◽  
Asad Asghar ◽  
William D. E. Allan ◽  
John P. Clark
Keyword(s):  
Boundary Layer ◽  
Turbine Blade ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
Blade Profile ◽  
Suction Surface ◽  
Low Reynolds Numbers ◽  
Low Pressure Turbine ◽  
Low Reynolds ◽  
Highly Loaded

Abstract Aircraft are operating at increasingly high-altitudes, where decreased air density and engine power settings have led to increasingly low Reynolds numbers in the low-pressure turbine portion of modern-day aeroengines. These operating conditions, in parallel with highly-loaded blade profiles, result in non-reattaching laminar boundary layer separation along the blade suction surface, increasing loss and decreasing engine performance. This work presents an experimental investigation into the potential for integrated leading-edge tubercles to improve blade performance in this operating regime. A turn-table cascade test-section was constructed and commissioned to test a purpose-designed, forward-loaded, low-pressure turbine blade profile at various incidences and Reynolds numbers. Baseline and tubercled blades were tested at axial chord Reynolds numbers at and between 15 000 and 60 000, and angles of incidence ranging from −5° to +10°. Experimental data collection included blade surface pressure measurements, total pressure loss in the blade wakes, hot-wire anemometry, surface hot-film measurements, and surface flow visualization using tufts. Test results showed that the implementation of tubercles did not lead to a performance enhancement. However, useful conclusions were drawn regarding the ability of tubercles to generate stream-wise vortices at ultra-low Reynolds numbers. Additional observations helped to characterize the suction surface boundary layer over the highly-loaded, low-pressure turbine blade profile when at off-design conditions. Recommendations were made for future work.

Predicting Separation and Transitional Flow in Turbine Blades at Low Reynolds Numbers—Part II: The Application to a Highly Separated Turbine Blade Cascade Geometry

2010 ◽  
Vol 133 (3) ◽  
Author(s):  
Darius D. Sanders ◽  
Walter F. O’Brien ◽  
Rolf Sondergaard ◽  
Marc D. Polanka ◽  
Douglas C. Rabe
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Performance Prediction ◽  
Flow Behavior ◽  
Reynolds Numbers ◽  
Transitional Flow ◽  
Suction Surface ◽  
Low Reynolds Numbers ◽  
Low Pressure Turbine ◽  
Low Reynolds

There has been a need for improved prediction methods for low pressure turbine (LPT) blades operating at low Reynolds numbers. This is known to occur when LPT blades are subjugated to high altitude operations causing a decrease in the inlet Reynolds number. Boundary layer separation is more likely to be present within the flowfield of the LPT stages due to increase in the region adverse pressure gradients on the blade suction surface. Accurate CFD predictions are needed in order to improve design methods and performance prediction of LPT stages operating at low Reynolds numbers. CFD models were created for the flow over two low pressure turbine blade designs using a new turbulent transitional flow model, originally developed by Walters and Leylek (2004, “A New Model for Boundary Layer Transition Using a Single Point RANS Approach,” ASME J. Turbomach., 126(1), pp. 193–202). Part I of this study applied Walters and Leylek’s model to a cascade CFD model of a LPT blade airfoil with a light loading level. Flows were simulated over a Reynolds number range of 15,000–100,000 and predicted the laminar-to-turbulent transitional flow behavior adequately. It showed significant improvement in performance prediction compared to conventional RANS turbulence models. Part II of this paper presents the application of the prediction methodology developed in Part I to both two-dimensional and three-dimensional cascade models of a largely separated LPT blade geometry with a high blade loading level. Comparisons were made with available experimental cascade results on the prediction of the inlet Reynolds number effect on surface static pressure distribution, suction surface boundary layer behavior, and the wake total pressure loss coefficient. The kT-kL-ω transitional flow model accuracy was judged sufficient for an understanding of the flow behavior within the flow passage, and can identify when and where a separation event occurs. This model will provide the performance prediction needed for modeling of low Reynolds number effects on more complex geometries.

Wind Tunnel Aeroacoustic Tests of Six Airfoils for Use on Small Wind Turbines*

2004 ◽  
Vol 126 (4) ◽  
pp. 974-985 ◽  
Author(s):  
Paul Migliore ◽  
Stefan Oerlemans
Keyword(s):  
Boundary Layer ◽  
Wind Tunnel ◽  
Wind Turbines ◽  
Leading Edge ◽  
Reynolds Numbers ◽  
Noise Sources ◽  
Low Reynolds Numbers ◽  
Trailing Edge Noise ◽  
Small Wind Turbines ◽  
Inflow Turbulence

Aeroacoustic tests of seven airfoils were performed in an open jet anechoic wind tunnel. Six of the airfoils are candidates for use on small wind turbines operating at low Reynolds numbers. One airfoil was tested for comparison to benchmark data. Tests were conducted with and without boundary layer tripping. In some cases, a turbulence grid was placed upstream in the test section to investigate inflow turbulence noise. An array of 48 microphones was used to locate noise sources and separate airfoil noise from extraneous tunnel noise. Trailing-edge noise was dominant for all airfoils in clean tunnel flow. With the boundary layer untripped, several airfoils exhibited pure tones that disappeared after proper tripping was applied. In the presence of inflow turbulence, leading-edge noise was dominant for all airfoils.

Prediction of Transitional and Separated Boundary Layers in a Compressor Cascade

2014 ◽  
Author(s):  
Manu Kamin ◽  
Joseph Mathew
Keyword(s):  
Boundary Layer ◽  
Separation Bubble ◽  
Three Dimensional ◽  
Reynolds Numbers ◽  
Reynolds Stress Model ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Low Reynolds Numbers ◽  
Pressure Distributions ◽  
Flow Features

Numerical simulations were performed of experiments from a cascade of stator blades at three low Reynolds numbers representative of flight conditions. Solutions were assessed by comparing blade surface pressures, velocity and turbulence intensity along blade normals at several stations along the suction surface and in the wake. At Re = 210,000 and 380,000 the laminar boundary layer over the suction surface separates and reattaches with significant turbulence fluctuations. A new 3-equation transition model, the k-kL-ω model, was used to simulate this flow. Predicted locations of the separation bubble, and profiles of velocity and turbulence fluctuations on blade-normal lines at various stations along the blade were found to be quite close to measurements. Suction surface pressure distributions were not as close at the lower Re. The solution with the standard k-ω SST model showed significant differences in all quantities. At Re = 640,000 transition occurs earlier and it is a turbulent boundary layer that separates near the trailing edge. The solution with the Reynolds stress model was found to be quite close to the experiment in the separated region also, unlike the k-ω SST solution. Three-dimensional computations were performed at Re = 380,000 and 640,000. In both cases there were no significant differences between the midspan solution from 3D computations and the 2D solutions. However, the 3D solutions exhibited flow features observed in the experiments — the nearly 2D structure of the flow over most of the span at 380,000 and the spanwise growth of corner vortices from the endwall at 640,000.

Laminar flow through a slowly rotating straight pipe

1985 ◽  
Vol 150 ◽  
pp. 1-21 ◽  
Author(s):  
Kamyar Mansour
Keyword(s):  
Boundary Layer ◽  
Laminar Flow ◽  
Reynolds Numbers ◽  
Similarity Parameter ◽  
Low Reynolds Numbers ◽  
Boundary Layer Analysis ◽  
Layer Analysis ◽  
Square Root Singularity ◽  
Flow Through ◽  
Steady Laminar

We consider fully developed steady laminar flow through a pipe that is rotating slowly about a line perpendicular to its own axis. The solution is expanded for low Reynolds numbers in powers of a single combined similarity parameter and the series extended to 34 terms by computer. Analysis shows that convergence is limited by a square-root singularity on the negative real axis of the similarity parameter. An Euler transformation and extraction of the leading, secondary and tertiary singularities at infinity render the series accurate for all values of the similarity parameter. The major conclusion of this investigation is that the friction ratio in a slowly rotating pipe grows asymptotically as the ⅛ power of the similarity parameter and not as the ¼ power as previously deduced from boundary-layer analysis. This discrepancy between the present computer-extended method and boundary-layer analysis has also occurred in the similar problem of flow through a loosely coiled pipe (Van Dyke 1978).

Essential Ingredients for the Computation of Steady and Unsteady Blade Boundary Layers

1991 ◽  
Vol 113 (4) ◽  
pp. 608-616 ◽  
Author(s):  
H. M. Jang ◽  
J. A. Ekaterinaris ◽  
M. F. Platzer ◽  
T. Cebeci
Keyword(s):  
Boundary Layer ◽  
Boundary Layers ◽  
Stokes Equations ◽  
Inviscid Flow ◽  
Reynolds Numbers ◽  
Navier Stokes ◽  
Transitional Region ◽  
Low Reynolds Numbers ◽  
Flow Equations ◽  
Low Reynolds

Two methods are described for calculating pressure distributions and boundary layers on blades subjected to low Reynolds numbers and ramp-type motion. The first is based on an interactive scheme in which the inviscid flow is computed by a panel method and the boundary layer flow by an inverse method that makes use of the Hilbert integral to couple the solutions of the inviscid and viscous flow equations. The second method is based on the solution of the compressible Navier–Stokes equations with an embedded grid technique that permits accurate calculation of boundary layer flows. Studies for the Eppler-387 and NACA-0012 airfoils indicate that both methods can be used to calculate the behavior of unsteady blade boundary layers at low Reynolds numbers provided that the location of transition is computed with the en method and the transitional region is modeled properly.

An Experimental Investigation of the Effect of Freestream Turbulence on the Wake of a Separated Low-Pressure Turbine Blade at Low Reynolds Numbers

1999 ◽  
Vol 122 (2) ◽  
pp. 431-433 ◽  
Author(s):  
C. G. Murawski ◽  
K. Vafai
Keyword(s):  
Reynolds Number ◽  
Turbine Blade ◽  
Turbulence Intensity ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
Freestream Turbulence ◽  
Suction Surface ◽  
Low Reynolds Numbers ◽  
Low Pressure Turbine ◽  
Flow Reynolds Number

An experimental study was conducted in a two-dimensional linear cascade, focusing on the suction surface of a low pressure turbine blade. Flow Reynolds numbers, based on exit velocity and suction length, have been varied from 50,000 to 300,000. The freestream turbulence intensity was varied from 1.1 to 8.1 percent. Separation was observed at all test Reynolds numbers. Increasing the flow Reynolds number, without changing freestream turbulence, resulted in a rearward movement of the onset of separation and shrinkage of the separation zone. Increasing the freestream turbulence intensity, without changing Reynolds number, resulted in shrinkage of the separation region on the suction surface. The influences on the blade’s wake from altering freestream turbulence and Reynolds number are also documented. It is shown that width of the wake and velocity defect rise with a decrease in either turbulence level or chord Reynolds number. [S0098-2202(00)00202-9]

The influence of the flow separation bubble and transition location on the profile drag of three 4-digit NACA aerofoil profiles

2021 ◽  
pp. 0309524X2110550
Author(s):  
Moutaz Elgammi ◽  
Tonio Sant ◽  
Atiyah Abdulmajid Ateeah
Keyword(s):  
Wind Tunnel ◽  
Flow Separation ◽  
Separation Bubble ◽  
Reynolds Numbers ◽  
Force Balance ◽  
Navier Stokes ◽  
Separation Mechanism ◽  
Numerical Work ◽  
Low Reynolds Numbers ◽  
Low Reynolds

Modeling of the flow over aerofoil profiles at low Reynolds numbers is difficult due to the complex physics associated with the laminar flow separation mechanism. Two major problems arise in the estimation of profile drag: (1) the drag force at low Reynolds numbers is extremely small to be measured in a wind tunnel by force balance techniques, (2) the profile drag is usually calculated by pressure integration, hence the skin friction component of drag is excluded. In the present work, three different 4-digit NACA aerofoils are investigated. Measurements are conducted in an open-ended subsonic wind tunnel, while numerical work is performed by time Reynolds-averaged Navier Stokes (RANS) coupled with the laminar-kinetic-energy ( K-kl-w) turbulence model. The influence of the flow separation bubbles and transition locations on the profile drag is discussed and addressed. This paper gives important insights into importance of measurements at low Reynolds numbers for better aerodynamic loads predictions.

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