Comment on "Angle of Attack and Leading Edge Effects on the Flow about a Flat Plate at Mach Number 18"

AIAA Journal ◽  
1968 ◽  
Vol 6 (12) ◽  
pp. 2459-2459
Author(s):  
JOSEPH H. KEMP
Keyword(s):  
Mach Number ◽  
Flat Plate ◽  
Edge Effects ◽  
Angle Of Attack ◽  
Leading Edge

An experimental investigation of resonant interaction of a rectangular jet with a flat plate

2015 ◽  
Vol 779 ◽  
pp. 751-775 ◽  
Author(s):  
K. B. M. Q. Zaman ◽  
A. F. Fagan ◽  
J. E. Bridges ◽  
C. A. Brown
Keyword(s):  
Mach Number ◽  
Flat Plate ◽  
Cross Section ◽  
Leading Edge ◽  
Resonant Interaction ◽  
Trailing Edge ◽  
Resonant Condition ◽  
Direction Parallel ◽  
Rectangular Jet ◽  
Axis Switching

The interaction between an 8:1 aspect ratio rectangular jet and a flat plate, placed parallel to the jet, is addressed in this study. At high subsonic conditions and for certain relative locations of the plate, a resonance takes place with accompanying audible tones. Even when the tone is not audible the sound pressure level spectra are often marked by conspicuous peaks. The frequencies of these peaks, as functions of the plate’s length, its location relative to the jet as well as jet Mach number, are studied in an effort to understand the flow mechanism. It is demonstrated that the tones are not due to a simple feedback between the nozzle exit and the plate’s trailing edge; the leading edge also comes into play in determining the frequency. With parametric variation, it is found that there is an order in the most energetic spectral peaks; their frequencies cluster in distinct bands. The lowest frequency band is explained by an acoustic feedback involving diffraction at the plate’s leading edge. Under the resonant condition, a periodic flapping motion of the jet column is seen when viewed in a direction parallel to the plate. Phase-averaged Mach number data on a cross-stream plane near the plate’s trailing edge illustrate that the jet cross-section goes through large contortions within the period of the tone. Farther downstream a clear ‘axis switching’ takes place for the time-averaged cross-section of the jet that does not occur otherwise for a non-resonant condition.

scholarly journals Geometric Effects of Thermal Barrier Coating Damage on Turbine Blade Temperatures

2019 ◽  
Author(s):  
Shane Colón ◽  
Mark Ricklick ◽  
Doug Nagy ◽  
Amy Lafleur
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Mach Number ◽  
Flat Plate ◽  
Turbine Blades ◽  
Leading Edge ◽  
Thermal Barrier ◽  
Cold Side ◽  
Hot Gas ◽  
Edge Geometry

Abstract Thermal barrier coatings (TBC) found on turbine blades are a key element in the performance and reliability of modern gas turbines. TBC reduces the heat transfer into turbine blades by introducing an additional surface thermal resistance; consequently allowing for higher gas temperatures. During the service life of the blades, the TBC surface may be damaged due to manufacturing imperfections, handling damage, service spalling, or service impact damage, producing chips in the coating. While an increase in aerofoil temperature is expected, it is unknown to what degree the blade will be affected and what parameters of the chip shape affect this result. During routine inspections, the severity of the chipping will often fall to the discretion of the inspecting engineer. Without a quantitative understanding of the flow and heat transfer around these chips, there is potential for premature removal or possible blade failure if left to operate. The goal of this preliminary study is to identify the major driving parameters that lead to the increase in metal temperature when TBC is damaged, such that more quantitative estimates of blade life and refurbishing needs can be made. A two-dimensional computational Conjugate Heat Transfer model was developed; fully resolving the hot gas path and TBC, bond-coat, and super alloy solids. Representative convective conditions were applied to the cold side to emulate the characteristics of a cooled turbine blade. The hot gas path properties included an inlet temperature of 1600 K with varying Mach numbers of 0.30, 0.59, and 0.80 and Reynolds number of 5.1×105, 7.0×105, and 9.0×105 as referenced from the leading edge of the model. The cold side was given a coolant temperature of 750 K and a heat transfer coefficient of 1500 W/m2*K. The assigned thermal conductivities of the TBC, bond-coat, and metal alloys were 0.7 W/m*K, 7.0 W/m*K, and 11.0 W/m*K, respectively, and layer thicknesses of 0.50 mm, 0.25 mm, and 1.50 mm, respectively. A flat plate model without the presence of the chip was first evaluated to provide a basis of validation by comparison to existing correlations. Comparing heat transfer coefficients, the flat plate model matched within uncertainty to the Chilton-Colburn analogy. In addition, flat plate results captured the boundary layer thickness when compared with Prandtl’s 1/7th power-law. A chip was then introduced into the model, varying the chip width and the edge geometry. The most sensitive driving parameters were identified to be the chip width and Mach number. In cases where the chip width reached 16 times the TBC thickness, temperatures increased by almost 30% when compared to the undamaged equivalents. Additionally, increasing the Mach number of the incoming flow also increased metal temperatures. While the Reynolds number based on the leading edge of the model was deemed negligible, the Reynolds number based on the chip width was found to have a noticeable impact on the blade temperature. In conclusion, this study found that chip edge geometry was a negligible factor, while the Mach number, chip width, and Reynolds number based on the chip width had a significant effect on the total metal temperature.

Effect of airfoil leading edge waviness on flow structures and noise

2016 ◽  
Vol 26 (6) ◽  
pp. 1821-1842 ◽  
Author(s):  
Man Zhang ◽  
Abdelkader Frendi
Keyword(s):  
Mach Number ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Space Time ◽  
Flow Structures ◽  
Mean Flow ◽  
Content Type ◽  
Acoustic Signature ◽  
Depth Analysis ◽  
Time Correlations

Purpose – The tubercles at the leading edge of Humpback Whale flippers have been shown to increase aerodynamic efficiency. The purpose of this paper is to compute the flow structures and noise signature of a NACA0012 airfoil with and without leading edge waviness, and located in the wake of a cylinder using the hybrid RANS-LES method. Design/methodology/approach – The mean flow Mach number is 0.2 and the angle of attack used is 2°. After benchmarking the method using existing experimental results, unsteady computations were then carried-out on both airfoil geometries and for a 2° angle of attack. Findings – Results from these computations confirmed the aerodynamic benefits of the leading edge waviness. Moreover, the wavy leading edge airfoil was found to be at least 4 dB quieter than its non-wavy counterpart. In-depth analysis of the computational results revealed that the wavy leading edge airfoil breaks up the large coherent structures which are then convected at higher speeds down the trough region of the waviness in agreement with previous experimental observations. This result is supported by both the two-point and space-time correlations of the wall pressure. Research limitations/implications – The limitations of the current findings reside in the fact that both the Reynolds number and the flow Mach number are low, therefore not applicable to aircrafts. In order to extend the study to practical aircrafts one needs huge grids and large computational resources. Practical implications – The results obtained here could have a huge implications on the design of future aircrafts and spacecrafts. More specifically, the biggest benefit from such redesign is the reduction of acoustic signature as well as increased efficiency in fuel consumption. Social implications – Reducing acoustic signature from aircrafts has been a major research thrust for NASA and Federal Aviation Administration. The social impact of such reduction would be improved quality of life in airport communities. For military aircrafts, this could results in reduced detectability and hence saving lives. Originality/value – Humpback Whales have been studied by various researchers to understand the effects of leading edge “tubercles” on flow structures. What is new in this study is the numerical confirmation of the effects of the tubercles on the flow structures and the resulting noise radiations. It is shown through the use of two-point correlations and space-time correlations that the flow structures in the trough area are indeed vortex tubes.

Some Boundary Layer Measurements on a Slender Wing at Supersonic Speeds

1963 ◽  
Vol 67 (629) ◽  
pp. 291-295
Author(s):  
R. T. Griffiths
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Aspect Ratio ◽  
Flat Plate ◽  
Layer Thickness ◽  
Incompressible Flow ◽  
Boundary Layer Thickness ◽  
Static Pressure ◽  
Leading Edge ◽  
Slender Wing

SummaryBoundary layer measurements have been made at four positions on a slender gothic wing of aspect ratio 0·75. Test's were made over a range of incidence at M=1·42 and 1·82. With transition fixed by roughness near the leading edge the boundary layer thickness varied little with small positive or negative incidence but was reduced at larger incidences, this being most marked at positive incidence for positions nearest the leading edge due to the influence of the wing vortex. With the exception of positions in the vicinity of the vortex, a good estimate of the boundary layer thickness was given by the theory for incompressible flow over a flat plate and an excellent estimate of the variation of local static pressure and Mach number with incidence was given by not-so-slender wing theory.

Nonlinear lift increase at high angles of attack for double swept waverider

2020 ◽  
Vol 34 (14n16) ◽  
pp. 2040124
Author(s):  
Chuan-Zhen Liu ◽  
Peng Bai
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Lower Surface ◽  
Aerodynamic Performance ◽  
Sweep Angle ◽  
Vital Interest ◽  
High Angles Of Attack ◽  
Nonlinear Increase

The nonlinear increase of the lift of the double swept waverider at high angles of attack is of vital interest. The aerodynamic performance of the double swept waverider is calculated and compared with that of single swept waveriders. Results suggest that the lift nonlinearity of the double swept waverider is stronger than that of equal-planform-area single swept one, and the nonlinearity increases as Mach number increases. Some scholars have proposed the “vortex lift” to explain the nonlinear lift increase, but it is questionable as the main lift of the waverider comes from the lower surface rather than the upper surface. This paper proposes another explanation that the nonlinear lift increase is related to the attachment of shock wave, influenced by the leading-edge sweep angle. The shock wave is more inclined to attach under the lower surface with smaller swept than that of larger swept as angle of attack increases. When the shock wave attaches, the pressure increase via angle of attack is nonlinear, leading to the nonlinearity of lift increase.

scholarly journals Flow Past a Flat Plate With a Vortex/Sink Combination

1996 ◽  
Vol 63 (2) ◽  
pp. 543-550 ◽  
Author(s):  
N. J. Mourtos ◽  
M. Brooks
Keyword(s):  
Flat Plate ◽  
Current Model ◽  
Angle Of Attack ◽  
Separated Flow ◽  
Leading Edge ◽  
Classical Solutions ◽  
Two Dimensional ◽  
Trailing Edge ◽  
Kutta Condition ◽  
Attached Flow

This paper presents a potential flow model for the leading edge vortex over a two-dimensional flat plate at an angle of attack. The paper is an extension of a model by Saffman and Sheffield (1977). A sink has been added in this model in an effort to satisfy the Kutta condition at both the leading edge and the trailing edge of the plate. The introduction of the sink was inspired by the fact that most steady vortices in nature appear in combination with a flow feature which can be interpreted as a sink at their cores when the flow is analyzed in a two-dimensional observation plane. As in the Saffman and Sheffield model, the presence of a vortex results in increased lift; however, in the current model a unique vortex/sink position is found at each angle of attack. A comparison has also been made between the lift and the drag of this model and the corresponding results for two classical solutions of flow over a flat plate: (a) the fully attached flow with the Kutta condition satisfied at the trailing edge only and (b) the Helmholtz solution of fully separated flow.

Unsteady Lift for Impulsively Started Transonic/Supersonic Flow

2015 ◽  
Author(s):  
Chen-Yuan Bai ◽  
Juan Li ◽  
Zi-Niu Wu
Keyword(s):  
Mach Number ◽  
Small Angle ◽  
Angle Of Attack ◽  
Phase Variation ◽  
Leading Edge ◽  
High Angle ◽  
Leading Edge Vortex ◽  
Three Phase ◽  
Edge Vortex ◽  
Starting Flow

The unsteady lift for incompressible starting flow of a flat plate at high angle of attack involves a repeatable three-phase variation: (a) initial lift drop, (b) a Wagner type lift increase enhanced by leading edge vortex and (c) a lift drop due to a lift-decreasing trailing edge vortex spiral induced by the leading edge vortex convected to the trailing edge. For compressible starting flow at small angle of attack, it is well known that the lift experiences an initial drop due to piston effect and then a Wagner type lift increase enhanced by compressibility. The third phase has not been reported in the past. In this paper we consider subsonic, transonic and supersonic starting flow at high angle of attack. Numerical computation using computational fluid dynamics is used to compute the flow and lift behavior is explained using existing theories. It is found that, when the angle of attack is 20 degrees, we still observe the three-phase lift variation for Mach number below 0.8. The second conclusion is that the lift during the Wagner type increase phase is a decreasing function of the Mach number, in contrast to what we know from piston and indicial function method for small angle of attack. Another important conclusion is, when the Mach number is high enough, say above 0.9, only two-phase variation is observed: (a) initial lift drop and (b) Wagner type lift increase. For supersonic starting flow the Wagner type lift increase is replaced by a linear increase.

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