Measurements and calculations of steady and oscillatory pressures on a low aspect ratio model at subsonic and transonic speeds

Mabey, D. G.; Welsh, B. L.

doi:10.1016/s0889-9746(87)90307-0

Cited by 7 publications

(4 citation statements)

References 5 publications

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“…For the first test case, pressure-coefficient distribution along the chord at semi-span sections of η equal to 14, 42, and 65% from the root is shown in Figs. 3. Having considered that the present solution is inviscid and viscous effects are not included, the agreement observed between the present results and the experimental data (Mabey et al 1984) is very good. At position η = 65% near the tip of the wing, however, some deviation from the experiment is seen near the trailing edge.…”

Section: Cfl I Cfl Er I Ersupporting

confidence: 82%

“…4. Again at each semi-span section pressure-coefficient distribution along the chord is compared with results of the experiment (Mabey et al 1984). The present results are in excellent agreement with experimental data.…”

Section: Cfl I Cfl Er I Ersupporting

confidence: 79%

“…In this section, supersonic flow is simulated around Royal Aircraft Establishment (RAE) tailplane (Mabey et al 1984) and a rectangular wing. The RAE tailplane has panel aspect ratio of 1.2, taper ratio of 0.27, leading edge sweep angle of 50.2° and approximately an airfoil of NACA 64A010.2.…”

Section: Numerical Resultsmentioning

confidence: 99%

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A New Algorithm for Shock Sensor Calculation at Supersonic Speeds on a 3D-Unstructured Grid

Tahmasbi

Karimian²

2016

J.Aerosp. Technol. Manag.

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In this paper, 3-Dsupersonic flow around two types of wings is solved using a new algorithm for shock sensor calculation. A dual-time-stepping implicit method with 2nd-order accuracy is used for time integration of the equations. In each real time step, the non-linear system of equations is solved by iterating in pseudo-time, using a multi-step integration method. A cell-center finite volume scheme is applied to discretize the solution domain. Governing equations are discretized using 2nd-order central scheme of Jameson. Undesirable oscillations are prevented using artificial dissipation terms containing 2nd and 4th-order derivative terms. The second-order derivative term is proportional to shock sensor, which is a function of pressure gradient in general and is devised to capture shock waves correctly. Appropriate calculation of shock sensor is very important especially for the solution of 3-D supersonic flow on unstructured grids. In this study, a simple efficient algorithm is proposed for shock sensor calculation to stabilize solution in supersonic 3-D flows on unstructured grids. The new algorithm, implemented at an in-house code, is evaluated by comparison of its results with wind tunnel test data and upwindtype differencing scheme of Roe for a tailplane model tested at Royal Aircraft Establishment. The results show that supersonic flow with shock waves has been accurately captured.

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Section: Cfl I Cfl Er I Ersupporting

confidence: 82%

Section: Cfl I Cfl Er I Ersupporting

confidence: 79%

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A New Algorithm for Shock Sensor Calculation at Supersonic Speeds on a 3D-Unstructured Grid

Tahmasbi

Karimian²

2016

J.Aerosp. Technol. Manag.

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“…Only upper surface pressures were measured for this model. 9 Subsonic and transonic calculations for this model using XTRAN3S and a coarse grid were presented in Ref. 19 Calculated steady results from CAP-TSD for a 0 « 0 and M = 1.20 and 1.71 are compared with measured pressures in Figs.…”

Section: Results For Rae Tailplanementioning

confidence: 99%

Calculation of steady and unsteady pressures on wings at supersonic speeds with a transonic small-disturbance code

Bennett

Bland

Batina

et al. 1991

Journal of Aircraft

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A transonic unsteady aerodynamic and aeroelasticity code has been developed for application to realistic aircraft configurations. The code is called CAP-TSD (computational aeroelasticity program-transonic smalldisturbance) and uses a time-accurate approximate factorization algorithm for solution of the unsteady transonic small-disturbance equation that is efficient for solution of steady and unsteady transonic flow problems including supersonic freestream flows. This code can treat complete aircraft geometries with multiple lifting surfaces and bodies. Applications to wings in supersonic freestream flow are presented. Comparisons with selected exact solutions from linear theory are presented, showing generally favorable results. Calculations for both steady and oscillatory cases for the F-5 wing and the Royal Aerospace Establishment tailplane models are compared with experimental data and also show good overall agreement. Selected steady calculations are also compared with the results from a steady flow Euler code. Nomenclature c = airfoil chord c/ a = section lift-curve slope c r = wing reference chord Cjj = pressure coefficient C p = unsteady pressure coefficient normalized by oscillation amplitude / = function defining position of lifting surface k = reduced frequency, = coc r /2£7 M = freestream Mach number t = time, nondimensionalized by U/c r U = freestream velocity a = angle of attack 7 = ratio of specific heats A? = nondimensional time step i? = fractional semispan coordinate 6 = phase angle between unsteady lift and pitch angle 0 = disturbance velocity potential o> = angular frequency Subscripts 0 = mean value 1 = dynamic value

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Effects of heat transfer on aerodynamics and possible implications for wind tunnel tests

Mabey¹

1990

Progress in Aerospace Sciences

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Measurements and calculations of steady and oscillatory pressures on a low aspect ratio model at subsonic and transonic speeds

Cited by 7 publications

References 5 publications

A New Algorithm for Shock Sensor Calculation at Supersonic Speeds on a 3D-Unstructured Grid

A New Algorithm for Shock Sensor Calculation at Supersonic Speeds on a 3D-Unstructured Grid

Calculation of steady and unsteady pressures on wings at supersonic speeds with a transonic small-disturbance code

Effects of heat transfer on aerodynamics and possible implications for wind tunnel tests

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