Experimental and computational surface and flow-field results for anall-body hypersonic aircraft

Lockman, W. K.; Lawrence, Scott L.; Cleary, J. W.

doi:10.2514/6.1990-3067

Cited by 4 publications

(3 citation statements)

References 6 publications

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“…Cross flow effects slightly increase aerodynamic heating on the flare. A computation using the Reynolds analogy (43) consistently underpredicts the data but not to the degree shown in Figure 9 for a Mach number 10.6 flow. Figure 11 shows comparisons between computed and measured Stanton number (non-dimensionalized by freestream density and velocity) along the delta-wing configuration 43 described for pressure coefficient results of Figure 5 (same wind tunnel conditions as well) and Mach number 7.4.…”

Section: Surface Heat Transfermentioning

confidence: 72%

“…Crossflow effects on the pressure coefficient are not as significant as those shown in Figure 3 (for smaller M ∞ ). Figure 5 shows the pressure coefficient distribution along a delta-wing body with a double-wedge airfoil, 3.83-degree leading edge angle and 7.63-degree trailing edge angle, tested in a wind tunnel at Mach number 7.4 and Reynolds number, based on total body length, of 25 million (Re based on body thickness of 2.23 million, D = 8.16 cm) 43 . The model was tested at zero yaw and 5.0°.…”

Section: Resultsmentioning

confidence: 99%

“…These discrepancies may be due to wing edge effects that are ignored by the code. A computation using the Reynolds analogy (43) consistently underpredicts the data. Reynolds analogy computations shown in Figures 9, 10 and 11 demonstrate that this approximation becomes less accurate for higher Mach numbers.…”

Section: Surface Heat Transfermentioning

confidence: 91%

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Aerothermodynamic Analysis for Axisymmetric Projectiles at Supersonic/Hypersonic Speeds

Nusca

1993

Engineering Computations

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An aerothermodynamic design code for axisymmetric projectiles has been developed using a viscous-inviscid interaction scheme. Separate solution procedures for the inviscid and the viscous (boundary layer) fluid dynamic equations are coupled by an iterative solution procedure. Non-equilibrium, equilibrium and perfect gas boundary layer equations are included. The non-equilibrium gas boundary layer equations assume a binary mixture (two species; atoms and molecules) of chemically reacting perfect gases. Conservation equations for each species include finite reaction rates applicable to high temperature air. The equilibrium gas boundary layer equations assume infinite rate reactions, while the perfect gas equations assume no chemical reactions. Projectile near-wall and surface flow profiles (velocity, pressure, density, temperature and heat transfer) representing converged solutions to both the inviscid and viscous equations can be obtained in less than two minutes on minicomputers. A technique for computing local reverse flow regions is included. Computations for yawed projectiles are accomplished using a coordinate system transformation technique that is valid for small angle-of-attack. Computed surface pressure, heat transfer rates and aerodynamic forces and moments for 1.25 ≤ Mach No. ≤ 10.5 are compared to wind tunnel and free flight measurements on flat plate, blunt-cone, and projectile geometries such as a cone-cylinder-flare.

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Section: Surface Heat Transfermentioning

confidence: 72%

Section: Resultsmentioning

confidence: 99%

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Aerothermodynamic Analysis for Axisymmetric Projectiles at Supersonic/Hypersonic Speeds

Nusca

1993

Engineering Computations

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Solution-adaptive techniques in computational heat transfer and fluid flow

Acharya

1994

Computational Mechanics

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Correlation of HIFiRE-5a Flight Data with Computed Pressure and Heat Transfer

Jewell

Miller

Kimmel

2017

Journal of Spacecraft and Rockets

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The HIFiRE-5 test article was an elliptic cone with a 2.5-mm nose radius and 2:1 aspect ratio and a 7-degree minor-axis half-angle. The vehicle was flown in April 2012. The upper stage of the sounding rocket failed to ignite, resulting in a peak Mach number of about 3 instead of the target of 7. Flight heat flux and pressure data (reduced from almost 300 thermocouples and 50 pressure transducers) have been compared to α-and β-dependent CFD results for pressure distribution, as well as laminar and turbulent heat-transfer results. Computations were performed at three time points in the ascent trajectory. At each time point, five values each of angle of attack and yaw, ranging from-5.0° to 5.0°, were computed. CFD pressures, normalized with p ∞ , were interpolated to the flight Mach numbers at specified times throughout the ascent and descent trajectories. At each flight time, α and β were estimated from measured pressure by determining the α-β combination that minimized the RMS difference between the measured and computed pressures. The vehicle attitude, as determined from measured pressure, was compared to the vehicle attitude derived from Inertial Measurement Unit (IMU) results for α and β from the flight. The two methods showed excellent agreement for the entirety of the ascent and reentry portions of the trajectory. A similar normalization of the laminar and turbulent heat transfer CFD results with St was compared to flight heat transfer measurements, and transition locations were inferred. Finally, a computational heat conduction analysis was made to verify assumptions inherent in the calculation of heat flux from temperature.

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Experimental and computational surface and flow-field results for anall-body hypersonic aircraft

Cited by 4 publications

References 6 publications

Aerothermodynamic Analysis for Axisymmetric Projectiles at Supersonic/Hypersonic Speeds

Aerothermodynamic Analysis for Axisymmetric Projectiles at Supersonic/Hypersonic Speeds

Solution-adaptive techniques in computational heat transfer and fluid flow

Correlation of HIFiRE-5a Flight Data with Computed Pressure and Heat Transfer

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