Hypersonic Boundary-Layer Transition

Stetson, Kenneth F.

doi:10.1007/978-1-4612-0379-7_7

Cited by 44 publications

(12 citation statements)

References 23 publications

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“…For the HEG data, the cone is slightly blunted with a nose radius of R n = 2.5 mm. According to the swallowing length correlations reported by Stetson (1980), the measurement station is at about twice the swallowing length, but the local boundary layer thickness may still be influenced by the nose bluntness (Stetson et al 1984). This may in part explain why the measured frequencies are slightly larger than the sharp-cone predictions.…”

Section: Stability Trendsmentioning

confidence: 96%

Stability of highly cooled hypervelocity boundary layers

2015

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The influence of high levels of wall cooling on the stability of hypervelocity boundary layers is investigated. Such conditions are relevant to experiments in high-enthalpy impulse facilities, where the wall temperature is much smaller than the free-stream temperature, as well as to some real flight scenarios. Some effects of wall cooling are well known, for instance, the stabilization of the first mode and destabilization of the second mode. In this paper, several new instability phenomena are investigated that arise only for high Mach numbers and high levels of wall cooling. In particular, certain unstable modes can travel supersonically with respect to the free stream, which changes the nature of the dispersion curve and leads to instability over a much wider band of frequencies. The cause of this phenomenon, the range of parameters for which it occurs and its implications for boundary layer stability are examined. Additionally, growth rates are systematically reported for a wide range of conditions relevant to high-enthalpy impulse facilities, and the stability trends in terms of Mach number and wall temperature are mapped out. Thermal non-equilibrium is included in the analysis and its influence on the stability characteristics of flows in impulse facilities is assessed.

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Section: Stability Trendsmentioning

confidence: 96%

Stability of highly cooled hypervelocity boundary layers

2015

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“…Usually only the dominant parameters are included and the others neglected. This means that when a correlation is applied to a new situation, it is assumed that all the neglected effects are unchanged, which is seldom the case (Stetson, 1987). Therefore if no experimental transition data exists for a desired application, an estimate of the transition Reynolds number is essentially guesswork (Anderson, 2006).…”

Section: Factors Influencing Transitionmentioning

confidence: 99%

“…Numerical modelling cannot be used to model disturbances such as boundary layer trips as they induce by-pass transition mechanisms which make the development of analytical models extremely difficult (Stetson, 1992).…”

Section: Importance To Hypersonicsmentioning

confidence: 99%

“…A lot of research has been completed to model the growth of these disturbances providing a good understanding of the natural transition process. However, disturbances such as isolated roughness are known to introduce bypass mechanisms making modelling extremely difficult (Stetson, 1992). Therefore there is currently "no general mechanism-based theory for determining the conditions under which roughness can cause transition (Schneider, 2007)."…”

Section: Effects Of Roughness On Transitionmentioning

confidence: 99%

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Experimental investigation of a three dimensional scramjet engine at hypervelocity conditions

Wise¹

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Air-breathing propulsion has the ability to provide more economical access-to-space than current rocket based systems. Access-to-space requires hypersonic flight within the atmosphere, the air-breathing cycle best suited to this is the scramjet. Next generation launch systems require a scramjet stage to operate into the hypervelocity regime (> 3 km/s). Performance data at these conditions is scarce due to the inability of ground test facilities to replicate the high total pressures and temperatures associated with flight. Also, flight experiments are subject to material limitations and are generally prohibitively expensive. This thesis details experiments which add to the currently limited data-set of scramjet performance at hypervelocity conditions.The Mach 12 rectangular-to-elliptical shape-transitioning (RESTM12) engine is currently under investigation at The University of Queensland as a potential candidate for access-to-space. The RESTM12 engine is designed to accelerate from Mach 6-12 as a part of a rocket-scramjet-rocket access-to-space system. Previous testing of the RESTM12 engine has shown good performance at off-design conditions and now as a result of this investigation, robust performance at its design condition.The primary aim of this thesis was to:"investigate experimentally whether robust supersonic combustion can be generated in a three-dimensional scramjet flowpath at a Mach 12 flight condition in an impulse facility."In order to achieve this two experimental campaigns where undertaken in the T4 Stalker Tube. Both experiments used the facilities Mach 10 nozzle to generate a test flow equivalent to Mach 11.8 flight at 38.3 km altitude. Initial experiments investigating hypervelocity boundary layer transition showed that natural transition would not occur in a length equivalent to the forebody and inlet of the engine. The ingestion of a laminar boundary layer into the inlet of the RESTM12 engine would increase the likelihood of developing a separation in the inlet that would not allow it to operate as intended. To reduce the susceptibility of the inlet to separation a boundary layer trip was required.i Based on the Hyper-X program's boundary layer trip development experiments (Berry et al., 2001a), several combinations of trip geometry, height and location were tested on a flat plate model resulting in a successful trip configuration. Now confident that a trip configuration capable of providing the RESTM12 engine with a turbulent boundary layer was designed, the engine was tested at its design condition.A half-scale model of the RESTM12 scramjet engine was tested in a semi-free-jet configuration utilising both inlet and step injection of gaseous hydrogen fuel. Once it was established that the inlet was operating correctly and sufficient test time was available the experiments began with fuel injection. Initial tests showed that robust combustion occurred when inlet injection was employed. The most successful fuelling configuration was a combined injection scheme where 31% of the fuel was ...

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“…The prediction of natural turbulent transition in the hypersonic regime is an ongoing effort 28,29 . The influence of boundary layer instabilities on turbulence is not well known and is a function of many parameters Cavity 6 American Institute of Aeronautics and Astronautics that are unique to a given configuration and flow conditions.…”

Section: Smooth Body Transitionmentioning

confidence: 99%

Aeroheating Environments for a Mars Smart Lander

Edquist

Liechty

Hollis

et al. 2006

Journal of Spacecraft and Rockets

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A proposed Mars Smart Lander is designed to reach the surface via lifting-body atmospheric entry (α = 16 deg) to within 10 km of the target site. CFD predictions of the forebody aeroheating environments are given for a direct entry from a 2005 launch. The solutions were obtained using an 8-species gas in thermal and chemical nonequilibrium with a radiative-equilibrium wall temperature boundary condition. Select wind tunnel data are presented from tests at NASA Langley Research Center. Turbulence effects are included to account for both smooth body transition and turbulence due to heatshield penetrations. Natural transition is based on a momentum-thickness Reynolds number value of 200. The effects of heatshield penetrations on turbulence are estimated from wind tunnel tests of various cavity sizes and locations. Both natural transition and heatshield penetrations are predicted to cause turbulence prior to the nominal trajectory peak heating time. Laminar and turbulent CFD predictions along the trajectory are used to estimate heat rates and loads. The predicted peak turbulent heat rate of 63 W/cm 2 on the heatshield leeward flank is 70% higher than the laminar peak. The maximum integrated heat load for a fully turbulent heat pulse is 38% higher than the laminar load on the heatshield nose. The predicted aeroheating environments with uncertainty factors will be used to design a thermal protection system. NOMENCLATUREmomentum thickness ∞ freestream --------------* Vehicle Analysis Branch, Senior Member, AIAA. † Aerothermodynamics Branch.

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Hypersonic Boundary-Layer Transition

Cited by 44 publications

References 23 publications

Stability of highly cooled hypervelocity boundary layers

Stability of highly cooled hypervelocity boundary layers

Experimental investigation of a three dimensional scramjet engine at hypervelocity conditions

Aeroheating Environments for a Mars Smart Lander

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