A survey of leeside flow and heat transfer on delta planform configurations

Dunavant, J. C.; Walberg, Gerald D.; Narayan, KY

doi:10.2514/6.1976-118

Cited by 10 publications

(6 citation statements)

References 2 publications

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“…Figures 6 and 13 of Ref. 6 show such plots for the cone and the orbiter, respectively, and the similarity between these results and the present one on a delta wing is fairly obvious.…”

Section: Heat-transfer Distributionssupporting

confidence: 85%

“…A recent survey paper by Dunavant etal. 6 reviews the literature and highlights problem areas that need further investigation.…”

Section: Introductionmentioning

confidence: 98%

“…The compression field/oblique shock wave is formed as the flow coming in from the leading edge turns to an axial direction in response to the flow coming in from the opposite leading edge. The details of the vortex Presented as Paper 76-118 at the AIAA 14th Aerospace Sciences Meeting, Washington, D.C., Jan. 26-28, 1976; submitted March 10, 1977; revision received Oct. 6,1977 flowfield, however, vary to a considerable extent, depending on the freestream and body geometrical parameters. Based on the details of the flow structure, the leeside flowfield of a delta wing can be classified as one of several types: 1) attached flows without any vortices, 2) attached flow with vortices embedded within the viscous layer, 3) flow attached at the leading edge but separated at the foot of the inboard shock wave, and 4) leading-edge separated flow.…”

Section: Introductionmentioning

confidence: 98%

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Leeside Flowfield and Heat Transfer of a Delta Wing at IM = PO

Narayan

1978

AIAA Journal

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An experimental investigation was carried out to study the flowfield and heat-transfer characteristics of the leeside of an 80-deg swept delta wing at M^ -10, Flow visualization studies were complemented by extensive heat-transfer measurements to obtain the details of the flowfield. Attached vortex-free flows and flows with inboard and leading-edge separation were observed. The dominant contributor to the heating load on the leeside is seen to be the vortices generated due to the boundary-layer separation, either at the leading edge or inboard. The experimental results also show that the angle of attack for incipient and leading-edge separation is strongly dependent on freestream Reynolds number. NomenclatureC Pm = model specific heat M^ = freestream Mach number R^ = freestream unit Reynolds number R^ = stream wise Reynolds number (x •/?«>) R^f = model length Reynolds number (1-R&) N S(tQo = Stanton number based on freestream conditionŝ s/,oo = Stanton number normalized with the corresponding centerline value q = heat-transfer rate t =time T r = recovery temperature T w = wall temperature x -coordinate along centerline from wing apex, positive downstream a -angle of attack of lee surface p m = model density -angle that a conical ray through apex makes with the centerline 0 S = angle that the separation line makes with centerline A = wing leading-edge sweepback angle X = viscous interaction parameter (M 3^ /V/?^/) \ m = model skin thickness

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“…Figures 6 and 13 of Ref. 6 show such plots for the cone and the orbiter, respectively, and the similarity between these results and the present one on a delta wing is fairly obvious.…”

Section: Heat-transfer Distributionssupporting

confidence: 85%

“…A recent survey paper by Dunavant etal. 6 reviews the literature and highlights problem areas that need further investigation.…”

Section: Introductionmentioning

confidence: 98%

Section: Introductionmentioning

confidence: 98%

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Leeside Flowfield and Heat Transfer of a Delta Wing at IM = PO

Narayan

1978

AIAA Journal

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“…We propose that this interaction between the embedded shock wave and the leeside boundary layer is similar to a glancing interaction between an oblique shock wave and boundary layer on a flat plate. The possibility that these two flows might be similar was first suggested by Dunavant et a/ (9) . (Rough estimates of the boundary between attached flow and flow with shock induced separation on a delta wing were also made using this proposal; however a systematic study does not seem to have been made.)…”

Section: Prediction Of Incipient Separation On the Delta Wing Lee Surmentioning

confidence: 85%

Shock-induced separated flows on the lee surface of delta wings

Seshadri¹,

Narayan²

1987

Aeronaut. j.

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Experiments were conducted to study shock-induced separated flows on the lee surface of delta wings with sharp leading edge at supersonic speeds. Two sets of delta wings of different thickness (10° and 25° normal angle), each with leading edge sweep angles varying from 45° to 70°, were tested. The measurements, carried out in a Mach number range from 1.4 to 3.0, included oil flow visualisations (on both sets of wings) and static pressure distributions (on the thicker wings only). Using the test results, some features of shock-induced separated flows, including in particular the boundary between this type of flow and fully attached flow, have been determined. The experimental results indicate that this boundary does not seem to show any significant dependence on wing thickness within the limit of thicknesses tested. It is shown that this boundary can be predicted for thin delta wings using a well known criterion for incipient separation in a glancing shock wave boundary layer interaction, namely that a pressure rise of 1.5 is required across the shock. Comparison of the predicted boundary with experimental results (from oil flow visualisations) shows good agreement.

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“…A considerable amount of work has been done by many investigators [1][2][3][4][5][6] wanting to understand the complex flow phenomena occurring on the lee side of delta wings at hypersonic speeds. These investigators have performed flow visualization, pressure and heat transfer measurements.…”

Section: Introductionmentioning

confidence: 99%

Thin flat balances for testing thin delta wings at hypersonic speeds

Ramesh¹

2001

Meas. Sci. Technol.

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This paper concerns the design and development of two special novel thin internal strain gauge balances for testing thin slab delta wings at hypersonic speeds. Of the two balances, one is a thin three-component balance and the other is a thin six-component balance. Unlike the conventional balances, which are circular in cross section, the balances discussed here are flat and have rectangular cross sections. Here, the design philosophy and the metallic architecture of these balances are discussed in detail. This paper is divided into three parts wherein the first part deals with the thin three-component balance the second part addresses the thin six-component balance and, in the third part, typical results obtained using these two balances are presented. The thin three-component balance has a thickness of 2.5 mm and can be housed inside a wing as thin as 6 mm. Because of its thinness, the axial output exhibited non-linear interaction from the normal force loading. Hence, combined loading is performed. In order to deduce the second-order coefficients, a new method has been proposed and a comparison with the existing method shows that this is better than the previous method. This balance has been used to generate aerodynamic data on thin slab delta wings with and without lee-side balance housing bodies. In this way, it is demonstrated that the lee-side bodies do interfere even when they come into the aerodynamic shadow of the wing. The thin six-component balance has a thickness of 4 mm and can be housed inside an 8 mm thick wing. Although this balance is also thin, it exhibited excellent linearity in all the bridges. In spite of this, it is subjected to several possible combinations of combined loading and the `work back' loads are calculated using the linear coefficients. The work back loads agree very well with the applied loads. Hence, linear coefficients have been used for the purpose of data reduction. The effects of flap deflection on the aerodynamic characteristics of thin slab delta wings have been studied using this balance. All the experiments were conducted at Mach 8.2 and a Reynolds number of 2.13×106.

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A survey of leeside flow and heat transfer on delta planform configurations

Cited by 10 publications

References 2 publications

Leeside Flowfield and Heat Transfer of a Delta Wing at IM = PO

Leeside Flowfield and Heat Transfer of a Delta Wing at IM = PO

Shock-induced separated flows on the lee surface of delta wings

Thin flat balances for testing thin delta wings at hypersonic speeds

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