Synchronized Velocity and Pressure Measurements of Supersonic Flow over a Finite Span Cavity with Leading Edge Slot Blowing

2016 ◽  
Author(s):  
Benjamin George ◽  
Lawrence S. Ukeiley ◽  
Louis N. Cattafesta ◽  
Kunihiko Taira
Keyword(s):  
Supersonic Flow ◽  
Leading Edge ◽  
Pressure Measurements ◽  
Finite Span

Pressure loading on curved leading edge wings in supersonic flow

1986 ◽  
Vol 23 (7) ◽  
pp. 574-581 ◽  
Author(s):  
Milton E. Vaughn ◽  
John E. Burkhalter
Keyword(s):  
Supersonic Flow ◽  
Leading Edge ◽  
Pressure Loading

Surface pressures on a wing moving with supersonic speed

1974 ◽  
Vol 341 (1625) ◽  
pp. 177-193 ◽  
Keyword(s):  
Supersonic Flow ◽  
Delta Wing ◽  
Leading Edge ◽  
Ideal Gas ◽  
Numerical Results ◽  
Coordinate Surface ◽  
Supersonic Speed ◽  
Extrapolation Procedure ◽  
Specific Heats ◽  
Limiting Values

Estimates for pressures on the surface of a given delta wing at zero incidence in a steady uniform stream of air are obtained by numerically integrating two semi-characteristic forms of equations which govern the inviscid supersonic flow of an ideal gas with constant specific heats. In one form of the equations coordinate surfaces are fixed in space so that the surface of the wing, which has round sonic leading edges, is a coordinate surface. In the other, two families of coordinates are chosen to be stream-surfaces. For each form of the equations, a finite difference method has been used to compute the supersonic flow around the wing. Convergence of the numerical results, as the mesh is refined, is slow near the leading edge of the wing and an extrapolation procedure is used to predict limiting values for the pressures on the surface of the wing at two stations where theoretical and experimental results have been given earlier by another worker. At one station differences between the results given here and the results given earlier are significant. The two methods used here produce consistent values for the pressures on the surface of the wing and, on the basis of this numerical evidence together with other cited numerical results, it is concluded that the pressures given here are close to the true theoretical values.

Tilt and Interfacial Fluid Pressure Measurements of a Disk Sliding on a Polymeric Pad

2005 ◽  
Vol 127 (1) ◽  
pp. 198-205 ◽  
Author(s):  
Sum Huan Ng ◽  
Len Borucki ◽  
C. Fred Higgs ◽  
Inho Yoon ◽  
Andre´s Osorno ◽  
...  
Keyword(s):  
Negative Pressure ◽  
Reynolds Equation ◽  
Fluid Pressure ◽  
Leading Edge ◽  
Dominant Factor ◽  
Pressure Measurements ◽  
Capacitive Sensing ◽  
One Dimensional ◽  
Pressure Mapping ◽  
Pressure Region

Previous experimental work has shown that negative fluid pressure does develop at the disk/pad interface during chemical mechanical polishing. However, these studies dealt with one-dimensional measurement and modeling. To better understand the problem, two-dimensional pressure mapping is carried out. In addition, the orientation of the disk is measured with a capacitive sensing technique. Results reveal a large negative pressure region at the disk/pad interface that is skewed toward the leading edge of the disk. The disk is also found to be leaning down toward the leading edge and toward the center of the pad. A mixed-lubrication model based on the Reynolds equation and taking into account the disk orientation angles has been developed. Modeling and experimental results show similar trends, indicating the tilting of the disk as a dominant factor in causing the negative pressure phenomenon.

Fan Stability With Leading Edge Damage: Blind Prediction and Validation

2021 ◽  
Author(s):  
E. J. Gunn ◽  
T. Brandvik ◽  
M. J. Wilson ◽  
R. Maxwell
Keyword(s):  
Supersonic Flow ◽  
High Speed ◽  
Leading Edge ◽  
Rotating Stall ◽  
Stall Inception ◽  
Blind Prediction ◽  
Edge Damage ◽  
The Impact ◽  
Operating Points ◽  
Relative Flow

Abstract This paper considers the impact of a damaged leading edge on the stall margin and stall inception mechanisms of a transonic, low pressure ratio fan. The damage takes the form of a squared-off leading edge over the upper half of the blade. Full-annulus, unsteady CFD simulations are used to explain the stall inception mechanisms for the fan at low- and high-speed operating points. A combination of steady and unsteady simulations show that the fan is predicted to be sensitive to leading edge damage at low speed, but insensitive at high speed. This blind prediction aligns well with rig test data. The difference in response is shown to be caused by the change between subsonic and supersonic flow regimes at the leading edge. Where the inlet relative flow is subsonic, rotating stall is initiated by growth and propagation of a subsonic leading edge flow separation. This separation is shown to be triggered at higher mass flow rates when the leading edge is damaged, reducing the stable flow range. Where the inlet relative flow is supersonic, the flow undergoes a supersonic expansion around the leading edge, creating a supersonic flow patch terminated by a shock on the suction surface. Rotating stall is triggered by growth of this separation, which is insensitive to leading edge shape. This creates a marked difference in sensitivity to damage at low- and high-speed operating points.

Effect of flat wall length on decay and shock structure of a supersonic square wall jet

2021 ◽  
pp. 095441002110580
Author(s):  
Venkata Satya Manikanta Tammabathula ◽  
Venkata Sai Krishna Ghanta ◽  
Tharaka Narendra Sridhar Bandla
Keyword(s):  
Supersonic Jet ◽  
Leading Edge ◽  
Shock Structure ◽  
Wall Jet ◽  
Pressure Measurements ◽  
Maximum Enhancement ◽  
Flat Wall ◽  
Decay Behaviour ◽  
Core Length ◽  
Supersonic Core Length

Experiments were conducted to find the effect of wall length on the decay behaviour and shock structure of a supersonic wall jet issuing from c-d nozzle of the square-shaped exit. A straight flat wall of width same as the side length of the square was attached to the lip of the nozzle such that the leading edge of the wall and the side of the square aligned properly which allowed the supersonic jet to graze past the flat wall. Experiments were conducted with five different wall lengths, that is, [Formula: see text] = 0.5, 1, 2, 4 and 8. Wall pressure measurements were made from leading edge to the trailing edge of the wall along its centreline. Schlieren flow visualization of the jet flow over the wall for the different wall lengths revealed the shock pattern and the effect of the wall length on the shock structure. The shock structure and jet deflection were significantly affected due to the presence of the wall. There was an upward jet deflection for [Formula: see text] up to [Formula: see text] whereas a downward jet deflection was observed for [Formula: see text]. Noticeable changes in the shock structure were observed for the wall lengths up to 2 D h. The wall length also significantly affected the jet decay characteristics and supersonic core length. Maximum enhancement in jet decay and maximum reduction in supersonic core length resulted when the wall length was [Formula: see text]. However, when the wall length was increased to [Formula: see text], there was a significant reduction in jet decay and a recovery of [Formula: see text]. Presence of wall always resulted a reduction in Lsc irrespective of wall length. The wall effect was to induce a more precipitous pressure drop closer to the nozzle exit, and a more gradual drop farther from it for [Formula: see text] > [Formula: see text].

On the flow field of a rapidly oscillating airfoil in a supersonic flow

1974 ◽  
Vol 62 (4) ◽  
pp. 811-827 ◽  
Author(s):  
M. Kurosaka
Keyword(s):  
Supersonic Flow ◽  
Flow Field ◽  
Leading Edge ◽  
Flow Region ◽  
Relative Magnitude ◽  
Frequency Parameter ◽  
Unsteady Supersonic Flow ◽  
Long Time ◽  
Intimate Knowledge ◽  
Repeated Use

This paper examines the features of the flow field off the surface of an oscillating flat-plate airfoil immersed in a two-dimensional supersonic flow Although the exact linearized solution for a supersonic unsteady airfoil has been known for a long time, its expression in the form of an integral is not convenient for a physical interpretation. In the present paper, the quintessential features of the flow field are extracted from the exact solution by obtaining an asymptotic expansion in descending powers of a frequency parameter through the repeated use of the stationary-phase and steepest descent methods. It is found that the flow field consists of two dominant and competing signals: one is the acoustic ray or that component arising from Lighthill's ‘convecting slab’ and the other is the leading-edge disturbance propagating as a convecting wavelet. The flow field is found to be divided into several identifiable regions defined by the relative magnitude of the signals, and the asymptotic expansions appropriate for each flow region are derived along with their parametric restrictions. Such intimate knowledge of the flow field in unsteady, supersonic flow is of interest for interference aerodynamics and related acoustic problems.

scholarly journals Interpreting Aerodynamics of a Transonic Impeller from Static Pressure Measurements

2018 ◽  
Vol 2018 ◽  
pp. 1-9
Author(s):  
Fangyuan Lou ◽  
John Charles Fabian ◽  
Nicole Leanne Key
Keyword(s):  
Flow Field ◽  
High Speed ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Pressure Measurements ◽  
High Loss ◽  
Supersonic Inlet ◽  
Mach Numbers ◽  
Inlet Conditions

This paper investigates the aerodynamics of a transonic impeller using static pressure measurements. The impeller is a high-speed, high-pressure-ratio wheel used in small gas turbine engines. The experiment was conducted on the single stage centrifugal compressor facility in the compressor research laboratory at Purdue University. Data were acquired from choke to near-surge at four different corrected speeds (Nc) from 80% to 100% design speed, which covers both subsonic and supersonic inlet conditions. Details of the impeller flow field are discussed using data acquired from both steady and time-resolved static pressure measurements along the impeller shroud. The flow field is compared at different loading conditions, from subsonic to supersonic inlet conditions. The impeller performance was strongly dependent on the inducer, where the majority of relative diffusion occurs. The inducer diffuses flow more efficiently for inlet tip relative Mach numbers close to unity, and the performance diminishes at other Mach numbers. Shock waves emerging upstream of the impeller leading edge were observed from 90% to 100% corrected speed, and they move towards the impeller trailing edge as the inlet tip relative Mach number increases. There is no shock wave present in the inducer at 80% corrected speed. However, a high-loss region near the inducer throat was observed at 80% corrected speed resulting in a lower impeller efficiency at subsonic inlet conditions.

Some Experiments on the Reattachment of a Laminar Boundary Layer Separating From a Rearward Facing Step on a Flat Plate Aerofoil

1960 ◽  
Vol 64 (599) ◽  
pp. 668-672 ◽  
Author(s):  
T. W. F. Moore
Keyword(s):  
Boundary Layer ◽  
Liquid Film ◽  
Flat Plate ◽  
Laminar Boundary Layer ◽  
Reverse Flow ◽  
Leading Edge ◽  
Reynolds Numbers ◽  
Pressure Measurements ◽  
Flow Vortex

Summary:The results of experiments on the reattachment of a laminar boundary layer, separating from a rearward facing step in a flat plate aerofoil, are correlated with the properties of the short leading edge bubble which forms on thin aerofoils near the stall.The experiments, comprising pressure measurements, Pitot explorations, liquid film and smoke studies, indicate that for all Reynolds numbers above the value given by the Owen-KIanfer criterion the reattachment is turbulent behind a stationary air reverse flow vortex bubble. It is also found that the reattachment is laminar for Reynolds numbers below the critical, which further supports Crabtree's interpretation of the Owen-KIanfer criterion in terms of the condition for the growth of turbulent bursts.

Supersonic Flow Past Wing-Body Combinations

1953 ◽  
Vol 4 (3) ◽  
pp. 287-314 ◽  
Author(s):  
W. Chester
Keyword(s):  
Supersonic Flow ◽  
Aspect Ratio ◽  
Leading Edge ◽  
The Body ◽  
Infinite Cylinder ◽  
Main Stream ◽  
Exact Equation ◽  
Body Of Revolution ◽  
Flow Past ◽  
General Conditions

SummaryThe supersonic flow past a combination of a thin wing and a slender body of revolution is discussed by means of the linearised equation of motion. The exact equation is first established so that the linearised solution can be fed back and the order of the error terms calculated. The theory holds under quite general conditions which should be realised in practice.The wing-body combination considered consists of a wing symmetrically situated on a pointed body of revolution and satisfying the following fairly general conditions. The wing leading edge is supersonic at the root, and the body is approximately cylindrical downstream of the leading edge. The body radius is of an order larger than the wing thickness, but is small compared with the chord or span of the wing.It is found that if the wing and body are at the same incidence, and the aspect ratio of the wing is greater than 2 (M2-1)-½, where M is the main stream Mach number, the lift is equivalent to that of the complete wing when isolated. If the wing only is at incidence then the lift is equivalent to that of the part of the wing lying outside the body.The presence of the body has a more significant effect on the drag. If, for example, the body is an infinite cylinder of radius a, and the wing is rectangular with aspect ratio greater than 2(M2-1)-½, then the drag of the wing is decreased by a factor (1-2a/b), where 2b is the span of the wing.When these conditions do not hold the results are not quite so simple but are by no means complicated.

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