Near Field Pressure Measurement around Free Flight 69 Degree Swept Back Delta Wing Model

The experimental results of high level fluctuation excitation by external Mach’s wave in the boundary layer of delta wing model with blunt leading edges at Mach numbers M = 2, 2.5, 4 are presented. The exitation areas and mass flow pulsation levels in the conditions of subsonic, sonic and supersonic leading edges have been defined. It was found that the maximum value of the pulsations is 12–15 % and varies only slightly from the flow conditions around of the delta wing

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Near-field tip vortex behind a swept wing model

Experiments in Fluids ◽

10.1007/s00348-005-0056-y ◽

2005 ◽

Vol 40 (1) ◽

pp. 141-155 ◽

Cited By ~ 23

Author(s):

P. Gerontakos ◽

T. Lee

Keyword(s):

Near Field ◽

Tip Vortex ◽

Swept Wing ◽

Wing Model

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Experimental investigation of pressure distribution characteristics of a flying-wing model using PSP

International Journal of Modern Physics B ◽

10.1142/s0217979220400883 ◽

2020 ◽

Vol 34 (14n16) ◽

pp. 2040088

Author(s):

Hongbiao Wang ◽

Baoshan Zhu ◽

Jian Xiong

Keyword(s):

Pressure Distribution ◽

Surface Pressure ◽

Pressure Measurement ◽

Pressure Coefficient ◽

Leading Edge ◽

Distribution Characteristics ◽

Characteristic Analysis ◽

Camera System ◽

Wing Model ◽

The Impact

To investigate the static pressure distribution characteristics of a flying-wing model, an advanced binary pressure sensitive paint (PSP) technique is introduced. It has low-temperature sensitivity and can compensate the errors induced by temperature. The pressure measurement test was performed in 0.6 m trisonic wind tunnel at angles of attack ranging from 0[Formula: see text] to 12[Formula: see text] in supersonic condition, adopting a low-aspect-ratio flying wing model. The binary PSP is sprayed on the upper surface of the model while pressure taps are installed on the upper surface of the right wing. Luminescent images of two probes are acquired with a color charge-coupled-device camera system and processed with calibration results. During the test, the surface pressure is measured by PSP and transducer, respectively. The results obtained show that the binary paint is of advantage to the surface pressure measurement and flow characteristic analysis. The high-resolution pressure spectra at different angle of attack clearly reveal the impact of leading edge vortex on the upper surface pressure distributions. The pressure measured by PSP also agrees well with the pressure tap results. The root mean square error of pressure coefficient is 0.01 at [Formula: see text], [Formula: see text].

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Surface Pressure Measurement on Supersonic Free-Flight Projectiles Using Unsteady PSP Techniques

2018 Aerodynamic Measurement Technology and Ground Testing Conference ◽

10.2514/6.2018-3315 ◽

2018 ◽

Author(s):

Daiju Numata ◽

Kiyonobu Ohtani

Keyword(s):

Surface Pressure ◽

Pressure Measurement ◽

Free Flight

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Computational and Analytical Investigation of Aerodynamic Derivatives of Similitude Delta Wing Model at Hypersonic Speeds

International Journal of Technology ◽

10.14716/ijtech.v8i3.6319 ◽

2017 ◽

Vol 8 (3) ◽

pp. 366

Author(s):

Musavir Bashir ◽

S. A. Khan ◽

Qummare Azam ◽

Ayub Ahmed Janvekar

Keyword(s):

Delta Wing ◽

Analytical Investigation ◽

Wing Model ◽

Aerodynamic Derivatives ◽

Hypersonic Speeds ◽

Derivatives Of

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Effects of the Propeller Advance Ratio on Delta Wing UAV Leading Edge Vortex

International Journal of Automotive and Mechanical Engineering ◽

10.15282/ijame.16.3.2019.10.0522 ◽

2019 ◽

Vol 16 (3) ◽

pp. 6958-6970 ◽

Cited By ~ 1

Author(s):

K. A. Kasim ◽

P. Segard ◽

S. Mat ◽

S. Mansor ◽

M. N. Dahalan ◽

...

Keyword(s):

Swirling Flow ◽

Delta Wing ◽

Active Flow Control ◽

Pressure Coefficient ◽

Leading Edge ◽

Adverse Pressure Gradient ◽

Leading Edge Vortex ◽

Wing Model ◽

Edge Vortex ◽

Advance Ratio

Delta wing is a triangular-shaped platform that can be applied into the unmanned aerial vehicle (UAV) or drone applications. However, the flow above the delta wing is governed by complex leading-edge vortex structures which result in complicated aerodynamics behaviour. At higher angles of attack, the vortex burst can take place when the swirling flow is unable to sustain the adverse pressure gradient. More studies are needed to understand these vortex phenomena. This paper addresses an experimental study of active flow control called propeller on a generic 55° swept angle sharp-edged delta wing model. In this experiment, a propeller was placed at two different locations. The first location was at the apex of the wing while the second position was at the rear of the wing. The experiments were conducted in a 1.5 × 2.0 m2 closed-loop wind tunnel facility at Universiti Teknologi Malaysia. The freestream velocities were set at 20 m/s and 25 m/s. The research consisted of an intensive surface pressure measurement above the wing surface to investigate the effects of rotating propeller towards the leading-edge vortex. The experiments were divided into four configurations. The clean wing configuration was performed without the propeller and followed by pusher-propeller configuration using 10-inch 9-inch propellers. The final configuration was the tractor-propeller with a 10-inch propeller. The results emphasise the influences of the propeller size and its location corresponding to vortex properties above the delta-winged UAV model. The findings had indicated that the vortex peak is increased when the propeller is installed for both pusher and tractor configurations. The results also indicate that the pressure coefficient is increased when the propeller advance ratio increases.

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On the Vortex Breakdown Phenomenon in High Angle of Attack Flows Over Delta Wing Geometries

Volume 1: Advances in Aerospace Technology ◽

10.1115/imece2014-39354 ◽

2014 ◽

Author(s):

Eric D. Robertson ◽

Varun Chitta ◽

D. Keith Walters ◽

Shanti Bhushan

Keyword(s):

Vortex Breakdown ◽

Delta Wing ◽

Angle Of Attack ◽

Leading Edge ◽

Navier Stokes ◽

Detached Eddy Simulation ◽

High Angle ◽

High Angle Of Attack ◽

Eddy Simulation ◽

Wing Model

Using computational methods, an investigation was performed on the physical mechanisms leading to vortex breakdown in high angle of attack flows over delta wing geometries. For this purpose, the Second International Vortex Flow Experiment (VFE-2) 65° sweep delta wing model was studied at a root chord Reynolds number (Recr) of 6 × 106 at various angles of attack. The open-source computational fluid dynamics (CFD) solver OpenFOAM was used in parallel with the commercial CFD solver ANSYS® FLUENT. For breadth, a variety of classic closure models were applied, including unsteady Reynolds-averaged Navier-Stokes (URANS) and detached eddy simulation (DES). Results for all cases are analyzed and flow features are identified and discussed. The results show the inception of a pair of leading edge vortices originating at the apex for all models used, and a region of steady vortical structures downstream in the URANS results. However, DES results show regions of massively separated helical flow which manifests after vortex breakdown. Analysis of turbulence quantities in the breakdown region gives further insight into the mechanisms leading to such phenomena.

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The Analysis about External Compressible Flow around an Aircraft Wing

Advanced Materials Research ◽

10.4028/www.scientific.net/amr.1044-1045.654 ◽

2014 ◽

Vol 1044-1045 ◽

pp. 654-658

Author(s):

Wei Long ◽

Zai Shuai Ling ◽

Zhen Dang

Keyword(s):

Flow Field ◽

Velocity Vector ◽

Delta Wing ◽

Angle Of Attack ◽

Flow Simulation ◽

Aircraft Wing ◽

Wing Model ◽

Fluent Simulation ◽

Lift And Drag ◽

Variation Tendency

The Steady flow simulation to selected the delta wing model for different angles of attack in the Maher number.The law of flow field changes with the angle of attack is gotten.Through the FLUENT simulation,The variation tendency of coefficient of lift and drag in the different angle of attack is gotten.Further reveals the change rule of Maher number, pressure, velocity and other parameters in the different angle of attack.With increasing angle of attack, Maher number distribution is sparse of the same position increases and the greater numerical.the distribution of velocity vector is sparse of the same position increases and the greater numerical.the pressure distribution is sparse of the same position increases and the greater numerical.

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