INVISCID FLOW FIELD PAST A POINTED CONE AT AN ANGLE OF ATTACK. PART 1 - ANALYSIS

1965 ◽  
Author(s):  
G. Moretti
Keyword(s):  
Flow Field ◽  
Angle Of Attack ◽  
Inviscid Flow ◽  
Pointed Cone

The Elliptic Solution of the Secondary Flow Problem

1983 ◽  
Vol 105 (3) ◽  
pp. 530-535 ◽  
Author(s):  
S. Abdallah ◽  
A. Hamed
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Three Dimensional ◽  
Inviscid Flow ◽  
Elliptic Solution ◽  
Dimensional Flow ◽  
Flow Vorticity ◽  
Dependent Variables ◽  
Two Dimensional Flow ◽  
Uniqueness Of The Solution

This paper presents the elliptic solution of the inviscid incompressible secondary flow in curved passages. The three-dimensional flow field is synthesized between 3 sets of orthogonal nonstream surfaces. The two-dimensional flow field on each set of surfaces is considered to be resulting from a source/sink distribution. The distribution and strength of these sources are dependent on the variation in the flow properties normal to the surfaces. The dependent variables in this formulation are the velocity components, the total pressure, and the main flow vorticity component. The governing equations in terms of these dependent variables are solved on each family of surfaces using the streamlike function formulation. A new mechanism is implemented to exchange information between the solutions on the three family surfaces, resulting into a unique solution. In addition, the boundary conditions for the resulting systems of equations are carefully chosen to insure the existence and uniqueness of the solution. The numerical results obtained for the rotational inviscid flow in a curved duct are discussed and compared with the available experimental data.

Accelerated flow past a symmetric aerofoil: experiments and computations

2007 ◽  
Vol 591 ◽  
pp. 255-288 ◽  
Author(s):  
T. K. SENGUPTA ◽  
T. T. LIM ◽  
SHARANAPPA V. SAJJAN ◽  
S. GANESH ◽  
J. SORIA
Keyword(s):  
Reynolds Number ◽  
Flow Field ◽  
Angle Of Attack ◽  
Stream Velocity ◽  
Published Data ◽  
Navier Stokes Equation ◽  
Moment Coefficient ◽  
Flow Past ◽  
Naca 0015 ◽  
Accelerated Flow

Accelerated flow past a NACA 0015 aerofoil is investigated experimentally and computationally for Reynolds number Re = 7968 at an angle of attack α = 30°. Experiments are conducted in a specially designed piston-driven water tunnel capable of producing free-stream velocity with different ramp-type accelerations, and the DPIV technique is used to measure the resulting flow field past the aerofoil. Computations are also performed for other published data on flow past an NACA 0015 aerofoil in the range 5200 ≤ Re ≤ 35000, at different angles of attack. One of the motivations is to see if the salient features of the flow captured experimentally can be reproduced numerically. These computations to solve the incompressible Navier–Stokes equation are performed using high-accuracy compact schemes. Load and moment coefficient variations with time are obtained by solving the Poisson equation for the total pressure in the flow field. Results have also been analysed using the proper orthogonal decomposition technique to understand better the evolving vorticity field and its dependence on Reynolds number and angle of attack. An energy-based stability analysis is performed to understand unsteady flow separation.

The Role of Vortices and Unsteady Effects During the Hovering Flight of Dragonflies

1979 ◽  
Vol 83 (1) ◽  
pp. 59-77 ◽  
Author(s):  
STUART B. SAVAGE ◽  
BARRY G. NEWMAN ◽  
DENIS T.-M. WONG
Keyword(s):  
Steady State ◽  
Flow Field ◽  
Flow Visualization ◽  
Inviscid Flow ◽  
Physical Explanation ◽  
Hovering Flight ◽  
Wing Motion ◽  
Unsteady Effects ◽  
Animal Flight

Weis-Fogh and Norberg concluded that steady-state aerodynamics is incapable of explaining how the dragonfly supports its weight during hovering. Norberg also concluded that the wing kinematics of Aeschna juncea L., as determined photographically, are incompatible with those proposed by Weis-Fogh for his Flip mechanism. The present paper has proposed an alternative lift-generating mechanism, various aspects of which are novel from the standpoint of animal flight. Flow visualization tests performed in water established the flow field during a complete cycle of the idealized wing motion. Using this information and unsteady inviscid flow theory the forces were analysed. A plausible balance of horizontal forces and more than sufficient lift were obtained. A physical explanation of the theory is provided for those who do not wish to study the mathematical details.

Two-Dimensional Experimental Investigation on the Effects of a Fowler Flap Gap and Overlap Size on the Wake Flow Field

2016 ◽  
Vol 139 (2) ◽  
Author(s):  
David Demel ◽  
Mohsen Ferchichi ◽  
William D. E. Allan ◽  
Marouen Dghim
Keyword(s):  
Experimental Investigation ◽  
Flow Field ◽  
Angle Of Attack ◽  
Wake Flow ◽  
Two Dimensional ◽  
Edge Region ◽  
Suction Surface ◽  
Trailing Edge ◽  
Piv Measurements ◽  
Image Velocimetry

This work details an experimental investigation on the effects of the variation of flap gap and overlap sizes on the flow field in the wake of a wing-section equipped with a trailing edge Fowler flap. The airfoil was based on the NACA 0014-1.10 40/1.051 profile, and the flap was deployed with 40 deg deflection angle. Two-dimensional (2D) particle image velocimetry (PIV) measurements of the flow field in the vicinity of the main wing trailing edge and the flap region were performed for the optimal flap gap and overlap, as well as for flap gap and overlap increases of 2% and 4% chord beyond optimal, at angles of attack of 0 deg, 10 deg, and 12 deg. For all the configurations investigated, the flow over the flap was found to be fully stalled. At zero angle of attack, increasing the flap gap size was found to have minor effects on the flow field but increased flap overlap resulted in misalignment between the main wing boundary layer (BL) flow and the slot flow that forced the flow in the trailing edge region of the main wing to separate. When the angle of attack was increased to near stall conditions (at angle of attack of 12 deg), increasing the flap gap was found to energize and improve the flow in the trailing edge region of the main wing, whereas increased flap overlap further promoted flow separation on the main wing suction surface possibly steering the wing into stall.

Investigation of the effects of angle of attack and tail deflection angle on the controlling tail flow field

2020 ◽  
Vol 3 (4) ◽  
pp. 309-326
Author(s):  
Ali Mokhtari ◽  
Azadeh Shahrian ◽  
Pooya Javadpoor Langroodi ◽  
Maryam Ghodrat
Keyword(s):  
Flow Field ◽  
Deflection Angle ◽  
Angle Of Attack

scholarly journals Numerical investigation of the low-frequency flow oscillation over a NACA-0012 aerofoil at the inception of stall

2019 ◽  
Vol 11 ◽  
pp. 175682931983368 ◽  
Author(s):  
Yasir A ElAwad ◽  
Eltayeb M ElJack
Keyword(s):  
Experimental Data ◽  
Flow Field ◽  
Separation Bubble ◽  
Lift Coefficient ◽  
Angle Of Attack ◽  
Low Frequency ◽  
Laminar Separation Bubble ◽  
Flow Oscillation ◽  
Laminar Separation ◽  
Naca 0012

High-fidelity large eddy simulation is carried out for the flow field around a NACA-0012 aerofoil at Reynolds number of [Formula: see text], Mach number of 0.4, and various angles of attack around the onset of stall. The laminar separation bubble is formed on the suction surface of the aerofoil and is constituted by the reattached shear layer. At these conditions, the laminar separation bubble is unstable and switches between a short bubble and an open bubble. The instability of the laminar separation bubble triggers a low-frequency flow oscillation. The aerodynamic coefficients oscillate accordingly at a low frequency. The lift and the drag coefficients compare very well to recent high-accuracy experimental data, and the lift leads the drag by a phase shift of [Formula: see text]. The mean lift coefficient peaks at the angle of attack of [Formula: see text], in total agreement with the experimental data. The spectra of the lift coefficient does not show a significant low-frequency peak at angles of attack lower than or equal the stall angle of attack ([Formula: see text]). At higher angles of attack, the spectra show two low-frequency peaks and the low-frequency flow oscillation is fully developed at the angle of attack of [Formula: see text]. The behaviour of the flow-field and changes in the turbulent kinetic energy over one low-frequency flow oscillation cycle are described qualitatively.

Investigation of aerodynamic coefficients at Mach 6 over conical, hemispherical and flat-face spiked body

2017 ◽  
Vol 121 (1245) ◽  
pp. 1711-1732 ◽  
Author(s):  
R. Kalimuthu ◽  
R. C. Mehta ◽  
E. Rathakrishnan
Keyword(s):  
Flow Field ◽  
Blunt Body ◽  
Angle Of Attack ◽  
Separated Flow ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Geometrical Parameters ◽  
Aerodynamic Coefficients ◽  
Aerodynamic Forces ◽  
Flow Field Characteristics

ABSTRACTA forward spike attached to a blunt body significantly alters its flow field characteristics and influences aerodynamic characteristics at hypersonic flow due to formation of separated flow and re-circulation region around the spiked body. An experimental investigation was performed to measure aerodynamic forces for spikes blunt bodies with a conical, hemispherical and flat-face spike at Mach 6 and at an angle-of-attack range from 0° to 8° and length-to-diameterL/Dratio of spike varies from 0.5 to 2.0, whereLis the length of the spike andDis diameter of blunt body. The shape of the leading edge of the spiked blunt body reveals different types of flow field features in the formation of a shock wave, shear layer, flow separation, re-circulation region and re-attachment shock. They are analysed with the help of schlieren pictures. The shock distance ahead of the hemisphere and the flat-face spike is compared with the analytical solution and is showing satisfactory agreement with the schlieren pictures. The influence of geometrical parameters of the spike, the shape of the spike tip and angle-of-attack on the aerodynamic coefficients are investigated by measuring aerodynamic forces in a hypersonic wind tunnel. It is found that a maximum reduction of drag of about 77% was found for hemisphere spike ofL/D= 2.0 at zero angle-of-attack. Consideration for compensation of increased pitching moment is required to stabilise the aerodynamic forces.

Analysis of the Three-Dimensional Flow in a Turbine Scroll

1980 ◽  
Vol 102 (3) ◽  
pp. 297-301 ◽  
Author(s):  
A. Hamed ◽  
E. Baskharone
Keyword(s):  
Finite Element Method ◽  
Finite Element ◽  
Flow Field ◽  
Three Dimensional ◽  
Inviscid Flow ◽  
Geometrical Shape ◽  
The Finite Element Method ◽  
Three Dimensional Flow ◽  
Cross Sectional ◽  
General Flow

The present analysis describes the three dimensional compressible inviscid flow in the scroll and the vaneless nozzle of a radial inflow turbine. The solution to this flow field, which is further complicated by the geometrical shape of the boundaries, is obtained using the finite element method. Symmetric and nonsymmetric scroll cross sectional geometries are investigated to determine their effect on the general flow field and on the exit flow condiitons.

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