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.

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.

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.

scholarly journals Transonic flutter characteristics of an airfoil with morphing devices

2020 ◽  
pp. 095441002095304
Author(s):  
Shun He ◽  
Shijun Guo ◽  
Wenhao Li
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Aerodynamic Force ◽  
Unsteady Aerodynamics ◽  
Leading Edge ◽  
Trailing Edge ◽  
Airfoil Surface ◽  
Transonic Flutter ◽  
Morphing Technology ◽  
And Shifts

An investigation into transonic flutter characteristic of an airfoil conceived with the morphing leading and trailing edges has been carried out. Computational fluid dynamics (CFD) is used to calculate the unsteady aerodynamic force in transonic flow. An aerodynamic reduced order model (ROM) based on autoregressive model with exogenous input (ARX) is used in the numerical simulation. The flutter solution is determined by eigenvalue analysis at specific Mach number. The approach is validated by comparing the transonic flutter characteristics of the Isogai wing with relevant literatures before applied to a morphing airfoil. The study reveals that by employing the morphing trailing edge, the shock wave forms and shifts to the trailing edge at a lower Mach number, and aerodynamic force stabilization happens earlier. Meanwhile, the minimum flutter speed increases and transonic dip occurs at a lower Mach number. It is also noted that leading edge morphing has negligible effect on the appearance of the shock wave and transonic flutter. The mechanism of improving the transonic flutter characteristics by morphing technology is discussed by correlating shock wave location on airfoil surface, unsteady aerodynamics with flutter solution.

scholarly journals Resonant Interaction of a Rectangular Jet with a Flat-plate

2014 ◽  
Author(s):  
Khairul Q. Zaman ◽  
Amy Fagan ◽  
Michelle Clem ◽  
Clifford A. Brown
Keyword(s):  
Flat Plate ◽  
Resonant Interaction ◽  
Rectangular Jet

scholarly journals Numerical Investigation of a Rectangular Jet Exhausting over a Flat Plate with Periodic Surface Deformations at the Trailing Edge

Aerospace ◽  
2021 ◽  
Vol 8 (11) ◽  
pp. 314
Author(s):  
Colby Niles Horner ◽  
Adrian Sescu ◽  
Mohammed Afsar ◽  
Eric Collins
Keyword(s):  
Flat Plate ◽  
Experimental Studies ◽  
Wing Surface ◽  
Wave Number ◽  
Noise Levels ◽  
Trailing Edge ◽  
Vehicle Performance ◽  
Rectangular Jet ◽  
Sound Levels ◽  
Jet Interactions

Multiple competing factors are forcing aircraft designers to reconsider the underwing engine pod configuration typically seen on most modern commercial aircraft. One notable concern is increasing environmental regulations on noise emitted by aircraft. In an attempt to satisfy these constraints while maintaining or improving vehicle performance, engineers have been experimenting with some innovative aircraft designs which place the engines above the wings or embedded in the fuselage. In one configuration, a blended wing concept vehicle utilizes rectangular jet exhaust ports exiting from above the wing ahead of the trailing edge. While intuitively one would think that this design would reduce the noise levels transmitted to the ground due to the shielding provided by the wing, experimental studies have shown that this design can actually increase noise levels due to interactions of the jet exhaust with the aft wing surface and flat trailing edge. In this work, we take another look at this rectangular exhaust port configuration with some notional modifications to the geometry of the trailing edge to determine if the emitted noise levels due to jet interactions can be reduced with respect to a baseline configuration. We consider various horizontal and vertical offsets of the jet exit with respect to a flat plate standing in for the aft wing surface. We then introduce a series of sinusoidal deformations to the trailing edge of the plate of varying amplitude and wave number. Our results show that the emitted sound levels due to the jet–surface interactions can be significantly altered by the proposed geometry modifications. While sound levels remained fairly consistent over many configurations, there were some that showed both increased and decreased sound levels in specific directions. We present results here for the simulated configurations which showed the greatest decrease in overall sound levels with respect to the baseline. These results provide strong indications that such geometry modifications can potentially be tailored to optimize for further reductions in sound levels.

On the Initial Structure of the Interaction between Wings and Bodies at Supersonic Speeds

1968 ◽  
Vol 19 (2) ◽  
pp. 135-148 ◽  
Author(s):  
K. Stewartson
Keyword(s):  
Supersonic Flow ◽  
Flat Plate ◽  
Cross Section ◽  
Leading Edge ◽  
Potential Curve ◽  
Pressure Curve ◽  
Initial Structure ◽  
Elementary Functions

SummaryUsing the methods of the linearised theory of supersonic flow, the shape of the pressure curve (or the potential curve) on the root chord of a wing, extending symmetrically from a fuselage of circular cylindrical cross-section, and near its leading edge is calculated in the following three cases: (a) the wing has a rounded leading edge and is non-lifting, (b) the wing is a flat plate at incidence with a supersonic leading edge, (c) the same as (b) except that the leading edge is subsonic. In all cases the fuselage is non-lifting and the plan form of the wing is a half delta. The results in (a) and (b) can be expressed in terms of elementary functions; the results in (b) and (c) are tabulated.

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