Boundary Layer Transition Measured by DIT on the PSP Rotor in Forward Flight

2021 ◽  
Author(s):  
A.D. Gardner ◽  
A. Weiss ◽  
J.T. Heineck ◽  
A.D. Overmeyer ◽  
H.R. Spooner ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Infrared Camera ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Forward Flight ◽  
Layer Transition ◽  
Pitching Airfoil ◽  
Bladed Rotor ◽  
Helicopter Main Rotor ◽  
Long Wave Infrared

A well-defined reference set of data for computational fluid dynamics and comprehensive code validation for a scaled helicopter main rotor with boundary layer transition in forward flight is presented. The boundary layer transition was measured using differential infrared thermography (DIT) on the top (suction) side of the NASA/Army “PSP rotor” in the NASA Langley 14-by-22-Foot Subsonic Tunnel. The tests used a FLIR X8500 SLS long-wave infrared camera to observe the three-bladed rotor. The boundary layer transition was detected for forward flight at an advance ratio of 0.3 (115 kt). The measured boundary layer transition positions are consistent with previous measurements and predicted boundary layer transition locations. A method for the analysis of DIT images for a rotor in forward flight is shown and validated based on computational analysis of a pitching airfoil with varying inflow, showing both qualitative and quantitative similarity to experimental data.

Helicopter Rotor Boundary Layer Transition Measurement in Forward Flight Using an Infrared Camera

2020 ◽  
Vol 65 (1) ◽  
pp. 2-14 ◽  
Author(s):  
A. D. Gardner ◽  
C. C. Wolf ◽  
J. T. Heineck ◽  
M. Barnett ◽  
M. Raffel
Keyword(s):  
Boundary Layer ◽  
Infrared Camera ◽  
Ground Effect ◽  
Boundary Layer Transition ◽  
Helicopter Rotor ◽  
Forward Flight ◽  
Layer Transition ◽  
Pitching Airfoil ◽  
Transition Position ◽  
The Difference

Boundary layer transition measurement was demonstrated using differential infrared thermography (DIT) on the top side of a helicopter rotor in forward flight, which detects the difference in the convective heat transfer at the boundary layer transition position. The tests used a FLIR X8500xc SLS long wave infrared camera to observe the DLR EC135 test helicopter rotor. The boundary layer transition was detected for hover out of ground effect (150 ft) and for forward flight at 80 kt (1700 ft). The measured boundary layer transition positions are consistent with previous measurements of the EC135 hovering in ground effect, and with predicted boundary layer transition positions. A method for the analysis of DIT images for a rotor in forward flight is shown, based on computational analysis of a pitching airfoil with varying inflow.

Numerical Study of Wall Influence on Boundary-Layer Transition for Two-Dimensional NACA 16-012 and 4412 Hydrofoil Sections

1990 ◽  
Vol 34 (01) ◽  
pp. 38-47
Author(s):  
R. Latorre ◽  
R. Baubeau
Keyword(s):  
Boundary Layer ◽  
Test Section ◽  
Numerical Study ◽  
Suction Side ◽  
Boundary Layer Separation ◽  
Boundary Layer Transition ◽  
Layer Transition ◽  
Turbulent Separation ◽  
Side Pressure ◽  
Effective Angle

One of the difficulties in hydrofoil model tests is the relatively low Reynolds number of the test piece and the presence of the test section walls. This paper presents the results of systematic calculations of the potential flow field of NA 4412 and NACA 16-012 hydrofoil in a test section with wall-to-chord ratios h/c -1.0. The corresponding boundary-layer calculations using the CERT calculation scheme are presented to show the influence of the nearby walls on shifting the location of the boundary-layer laminar-turbulent separation as well as turbulent separation. By introducing an effective angle of attack, it is possible to obtain close agreement in the calculated and measured suction side pressure distortion as well as the locations of the boundary-layer separation and transition.

Effect of Sweep on a Pitching Finite Wing

2019 ◽  
Vol 64 (3) ◽  
pp. 1-13 ◽  
Author(s):  
A. D. Gardner ◽  
C. B. Merz ◽  
C. C. Wolf
Keyword(s):  
Boundary Layer ◽  
Single Case ◽  
Suction Side ◽  
Maximum Load ◽  
Boundary Layer Transition ◽  
Dynamic Stall ◽  
Sweep Angle ◽  
Tunnel Wall ◽  
Layer Transition ◽  
Finite Wing

An investigation was performed into the effect of positive and negative sweep angle on the boundary layer transition and dynamic stall behavior of a finite wing. The finite wing had a 6:1 aspect ratio, modern (SPP8) tip shape, and positive twist, moving the maximum load on the wing away from the wind tunnel wall. Experiments were performed with sweep Λ = ±30° and Λ = 0° for static polars and sinusoidal pitching. The positively twisted wing shows a S-shaped boundary layer transition on the pressure side similar to that previously seen for helicopter rotor blades in hover. The transition positions on the suction side of the wing are comparable for the same local angle of attack at all values of the sweep at each of the three pressure sections, and for dynamic pitching motions a hysteresis around the static transition positions is seen. Sweeping the wing led to later stall and higher maximum lift for both static polars and dynamic stall, except for a single case. The negative aerodynamic damping is worse for the swept wing than for the unswept wing, except where the delay of stall led to the flow remaining attached.

Effect of Surface Roughness on Conjugate Heat Transfer of a Turbine Vane

2016 ◽  
Author(s):  
Hongyang Li ◽  
Yun Zheng
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Surface Roughness ◽  
Suction Side ◽  
Transition Process ◽  
Considerable Effect ◽  
Boundary Layer Transition ◽  
Layer Transition ◽  
Turbine Vane ◽  
Effect Of Surface

For the purpose of researching the effect of surface roughness on boundary layer transition and heat transfer of turbine blade, a roughness modification approach for γ-Reθ transition model was proposed based on an in-house CFD code. Taking surface roughness effect into consideration, No. 5411 working condition of Mark II turbine vane was simulated and the results were analyzed in detail. Main conclusions are as follows: Surface roughness has little effect on heat transfer of laminar boundary layer, while has considerable effect on turbulent boundary layer. Compared with smooth surface, equivalent sand roughness of 100μm increases the temperature for about 28.4K on suction side, reaching an increase of 5%. Under low roughness degree, effect of shock wave dominants on boundary layer transition process on suction side, while above the critical degree, effect of surface roughness could abruptly change the transition point.

Numerical Simulation of the Boundary Layer Transition in Turbomachinery Flows

2001 ◽  
Author(s):  
Hans Thermann ◽  
Michael Müller ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Separated Flow ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Test Cases ◽  
Compressor Cascade ◽  
Layer Transition ◽  
Transonic Compressor ◽  
Transition Models ◽  
Transition Modeling

The objective of the presented work is to investigate models which simulate boundary layer transition in turbomachinery flows. This study focuses on separated-flow transition. Computations with different algebraic transition models are performed three-dimensionally using an implicit Navier-Stokes flow solver. Two different test cases have been chosen for this investigation: First, a linear transonic compressor cascade, and second an annular subsonic compressor cascade. Both test cases show three-dimensional flow structures with large separations at the side-walls. Additionally, laminar separation bubbles can be observed on the suction and pressure side of the blades of the annular subsonic cascade whereas a shock-induced separation can be found on the suction side of the blades of the linear transonic cascade. Computational results are compared with experiments and the effect of transition modeling is analyzed. It is shown that the prediction of the boundary layer development can be substantially improved compared to fully turbulent computations when algebraic transition models are applied.

Shock-boundary layer interaction and transition modelling in turbomachinery flows

1997 ◽  
Vol 211 (3) ◽  
pp. 225-234 ◽  
Author(s):  
V Michelassi
Keyword(s):  
Boundary Layer ◽  
Turbine Blade ◽  
Turbulence Model ◽  
Flow Separation ◽  
Good Description ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Layer Transition ◽  
Boundary Layer Development ◽  
Layer Development

The transonic turbulent compressible flow in channels and turbine linear cascades is computed by using a Navier-Stokes solver. Turbulence effects are simulated by means of the k-ω turbulence model. A realiability constraint is introduced to improve the turbulence model performances and stability in the presence of stagnation points. In both the flow over the bump and the turbine blade, the shock induces a flow separation that affects the boundary layer development. In both cases the proposed model succeeds in predicting the flow separation. For the flow over the turbine blade a simple transition model based on integral parameters is introduced to mimic the effect of the boundary layer transition across the shock wave on the suction side. Relaminarization is also properly predicted on the pressure side, thereby allowing a good description of the boundary layer development and shock pattern.

A CFD Study of Wake-Induced Transition on a Compressor-Like Flat Plate

2003 ◽  
Author(s):  
D. Keith Walters ◽  
James H. Leylek
Keyword(s):  
Boundary Layer ◽  
Flat Plate ◽  
Axial Compressor ◽  
Suction Side ◽  
General Purpose ◽  
Boundary Layer Transition ◽  
Dramatic Improvement ◽  
Suction Surface ◽  
Layer Transition ◽  
Freestream Velocity

Recent experimental work has documented the importance of wake passing on the behavior of transitional boundary layers on the suction surface of axial compressor blades. This paper documents computational fluid dynamics (CFD) simulations using a commercially-available general-purpose CFD solver, performed on a representative case with unsteady transitional behavior. The study implements a new, advanced version of a three-equation eddy-viscosity model previously developed and documented by the authors, which is capable of resolving boundary-layer transition. It is applied to the test cases of steady and unsteady boundary-layer transition on a 2-D flat plate geometry with a freestream velocity distribution representative of the suction side of a compressor airfoil. The CFD results are analyzed and compared to a similar experimental test case from the open literature. Results with the new model show a dramatic improvement over more typical RANS-based modeling approaches, and highlight the importance of resolving transition in both steady and unsteady compressor aero simulations.

Effects of Large Scale High Freestream Turbulence, and Exit Reynolds Number on Turbine Vane Heat Transfer in a Transonic Cascade

2007 ◽  
Author(s):  
S. Nasir ◽  
J. S. Carullo ◽  
W. F. Ng ◽  
K. A. Thole ◽  
H. Wu ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Reynolds Number ◽  
Large Scale ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Freestream Turbulence ◽  
Suction Surface ◽  
Layer Transition ◽  
Mach Numbers

This paper experimentally and numerically investigates the effect of large scale high freestream turbulence intensity and exit Reynolds number on the surface heat transfer distribution of a turbine vane in a 2-D linear cascade at realistic engine Mach numbers. A passive turbulence grid was used to generate a freestream turbulence level of 16% and integral length scale normalized by the vane pitch of 0.23 at the cascade inlet. The baseline turbulence level and integral length scale normalized by the vane pitch at the cascade inlet were measured to be 2% and 0.05, respectively. Surface heat transfer measurements were made at the midspan of the vane using thin film gauges. Experiments were performed at exit Mach numbers of 0.55, 0.75 and 1.01 which represent flow conditions below, near, and above nominal conditions. The exit Mach numbers tested correspond to exit Reynolds numbers of 9 × 105, 1.05 × 106, and 1.5 × 106, based on true chord. The experimental results showed that the large scale high freestream turbulence augmented the heat transfer on both the pressure and suction sides of the vane as compared to the low freestream turbulence case and promoted slightly earlier boundary layer transition on the suction surface for exit Mach 0.55 and 0.75. At nominal conditions, exit Mach 0.75, average heat transfer augmentations of 52% and 25% were observed on the pressure and suction side of the vane, respectively. An increased Reynolds number was found to induce earlier boundary layer transition on the vane suction surface and to increase heat transfer levels on the suction and pressure surfaces. On the suction side, the boundary layer transition length was also found to be affected by increase changes in Reynolds number. The experimental results also compared well with analytical correlations and CFD predictions.

Investigation of the Effectiveness of Various Types of Boundary Layer Transition Elements of Low Reynolds Number Turbine Bladings

2004 ◽  
Author(s):  
Claus H. Sieverding ◽  
Carlo Bagnera ◽  
A. C. Boege ◽  
Juan A. Cordero Anto`n ◽  
Vincent Lue`re
Keyword(s):  
Boundary Layer ◽  
Guide Vane ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Boundary Layer Separation ◽  
Boundary Layer Transition ◽  
Transition Elements ◽  
Layer Transition ◽  
Optimum Position ◽  
Low Reynolds

The paper describes an experimental investigation of the use of different types of boundary layer transition elements for the control of boundary layer separation at low Reynolds numbers. The tests are carried out in a low speed cascade tunnel for Reynolds numbers between 30000 and 200000. For convenience the author used an existing HP turbine guide vane with ∼63 degree turning. To obtain representative adverse pressure gradients as those existing on the rear suction side of highly loaded LP blades the tests are run at a pitch-to-chord ratio of 1. The transition elements include tripwires, single and double rows of spherical roughness elements, balloon type transition elements and a metal sheet actuated by shape memory alloy springs. The optimum position and height of the transition elements are obtained with systematic tests with the trip wire. All other elements are placed at the same position and have approximately the same height. As expected, the transition elements are very beneficial at low Re numbers but deteriorate the performance at high Re numbers. The advantages and drawbacks of the various configurations are discussed and suggestions for real turbine applications are made.

Computational Fluid Dynamics Study of Wake-Induced Transition on a Compressor-Like Flat Plate

2005 ◽  
Vol 127 (1) ◽  
pp. 52-63 ◽  
Author(s):  
D. Keith Walters ◽  
James H. Leylek
Keyword(s):  
Fluid Dynamics ◽  
Boundary Layer ◽  
Computational Fluid Dynamics ◽  
Flat Plate ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Dramatic Improvement ◽  
Suction Surface ◽  
Layer Transition ◽  
Freestream Velocity

Recent experimental work has documented the importance of wake passing on the behavior of transitional boundary layers on the suction surface of axial compressor blades. This paper documents computational fluid dynamics (CFD) simulations using a commercially available general-purpose CFD solver, performed on a representative case with unsteady transitional behavior. The study implements an advanced version of a three-equation eddy-viscosity model previously developed and documented by the authors, which is capable of resolving boundary layer transition. It is applied to the test cases of steady and unsteady boundary layer transition on a two-dimensional flat plate geometry with a freestream velocity distribution representative of the suction side of a compressor airfoil. The CFD results are analyzed and compared to a similar experimental test case from the open literature. Results with the model show a dramatic improvement over more typical Reynolds-averaged Navier–Stokes (RANS)-based modeling approaches, and highlight the importance of resolving transition in both steady and unsteady compressor aerosimulations.

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