Skin-friction measurements in three-dimensional, supersonic shock-wave/boundary-layer interaction

Wideman, Jeffrey K.; Brown, James L.; Miles, John B.; Özcan, O.

doi:10.2514/3.12503

Cited by 17 publications

(17 citation statements)

References 9 publications

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“…Taken from Wideman, et al 13 , Equation 13 reduces to the basic hydrodynamic LISF equation which for this case is…”

Section: Semi-direct Methodsmentioning

confidence: 99%

Development of fiber optic aerodynamic sensors for high Reynolds number supersonic flows

Pulliam¹,

Schetz²

2001

39th Aerospace Sciences Meeting and Exhibit

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Aerospace Engineering (ABSTRACT)The purpose of the project was to examine fiber optic sensors for the measurement of pressure, skin friction, temperature, and heat flux in high Reynolds number, supersonic flow. Using a standard fiber optic signal conditioning unit (specifically a broadband interferometric system using spectra), the work centered around determining under what conditions these sensors will work effectively and quantifying the total system limitations.An interferometric-based, fiber optic skin friction sensor was developed for the measurement of wall shear stress in complex, supersonic flows. This sensor type was tested successfully in laminar, incompressible flow, and supersonic flow up to Mach 1.92, Mach 2.4 and 3.0 flow, in which the sensor operated with varying success. A micromachined, fiber optic pressure sensor was also tested in these supersonic conditions, also with varying success. The accurate operation of these sensors was found to be tied to the flow conditions and the fiber optic, signal processing system.A correlation was found to exist between the energy of the flow, either through its dynamic pressure or through external disturbances such as shocks or separation, and the noise in the signals, expressed by the variance of the gap estimate, for the pressure and skin friction sensors in these flows. The energy of the flow couples with the mechanical properties of the sensor reducing the fringe contrast of the signal used by the optical signal processing system to determine a gap estimate. As the energy of the flow is increased and the sensor is excited, the fringe contrast is reduced. A practical limit of a normalized fringe contrast of 0.10 was found for producing accurate gap estimates in real flows. A consequence is that there is a limit to the dynamic pressure of the flow for the sensors to operate accurately, which is demonstrated by the varying success of the supersonic wind tunnel tests. This correlation is sensor specific, meaning that sensors can be designed to operate successfully in any flow. Also, the signal processing system, which forms the other end of the total system, could be improved to allow accurate measurements with the current sensors.iii

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“…Taken from Wideman, et al 13 , Equation 13 reduces to the basic hydrodynamic LISF equation which for this case is…”

Section: Semi-direct Methodsmentioning

confidence: 99%

Development of fiber optic aerodynamic sensors for high Reynolds number supersonic flows

Pulliam¹,

Schetz²

2001

39th Aerospace Sciences Meeting and Exhibit

View full text Add to dashboard Cite

show abstract

“…[13] Further separating the report of Widemann et al [11] from the present work is the fixed flare angle of 20 degrees used in that study. The present work will utilise a varying wing angle experiment to obtain measurements of flow topology with respect to angle, as opposed to the fixed topology utilised by Wideman et al The 20 degree fixed angle of the flare represents the higher end of sweep angles used in the present work.…”

Section: Heat Transfer In Wing-root Interactions 22mentioning

confidence: 76%

“…Existing literature in experimental studies of this kind of interaction are confined to supersonic and "near hypersonic" conditions, with one experimental study observing interactions at Mach 3 [11], the highest velocity for this kind of investigation to the Author's knowledge. Hung et.…”

Section: Heat Transfer In Wing-root Interactions 22mentioning

confidence: 99%

“…They noted that flow topology bifurcates at sufficiently high angles of sweep (68° for the studied conditions), with a noted reduction in upstream influence of the flow field and weakening of the horseshoe vortex formed about the wedge. The Wideman, et al [11] study was carried out at the NASA Ames Research Center, utilising Laser Interferometric Skin-Friction (LISF) [12] to quantify skin friction. It confirmed the general topology of this flow involved the interaction of a cylinder/offset flare combination with an incoming turbulent boundary layer, by visualising the oil patterns.…”

Section: Heat Transfer In Wing-root Interactions 22mentioning

confidence: 99%

“…Combined with the relatively lower Mach number of 3 compared with a Mach number of 8 proposed for the present work, the Widemann et. al [11] paper will serve only as a guide for the expected topology as well as possibly providing additional insight into how higher Mach numbers affect these types of flow topologies.…”

Section: Heat Transfer In Wing-root Interactions 22mentioning

confidence: 99%

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Experimental Study of Hypersonic Wing/Fin Root Heating at Mach 8

Schwarz¹

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The junction between wings or fins and the body of supersonic flight vehicles generate complex vortical structures and shock impingements. In hypersonic flight this interaction can lead to high localised heating rates. While the general topology of this flow has been established, confidence in the heating rates predicted by computational simulation is not high.This study utilises a dense set of heat transfer measurements on a model consisting of a flat plate and a cylinder with an adjustable sweep angle to help better understand and quantify flow features and heating rates on a swept protruding wing in hypersonic flight. Tests were conducted in the T4 shock tunnel facility at the University of Queensland Centre for Hypersonics at conditions simulating Mach 8 flight at 30km altitude. The resulting data is intended to help better understand flow topologies over wing-root junctions at hypersonic speeds, where relatively little testing has been conducted.A series of tests were conducted at both "low" and "high" pressure conditions, representing approximate atmospheric conditions at Mach 8 (total enthalpy ca. 3 MJ/kg) and a higher pressure case respectively, the latter being tested with the intention of inducing turbulence and measuring the effects of higher enthalpy and total pressure (a flow trip was later designed to ensure transition of the upstream boundary layer). The present study succeeded in correlating heating rates to empirical equations in the flat plate region upstream of the interaction, where the flow could be modelled as a simple flat plate of known Reynolds number.Correlation to theory, where applicable, was good. The use of a trip successfully caused flow to consistently transition from laminar to turbulent. Schlieren images were used to successfully match measured heating peaks with observed shock impingements.The present work was able to demonstrate the effectiveness of a trip in triggering turbulence.It demonstrated a correlation between flat plate heating and sweep angle, as well as providing a profile of heating rates along the centreline of a wing protruding from a wing root. Schlieren images added further insight by demonstrating positive the relationship between recirculation region size and peak heating rates.Future work should focus on obtaining more reliable data for the scenarios studied herein, particularly on the wing, where peak heating is expected to significantly exceed heating rates observed on the flat plate.ii

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