Transition Experiments on Large Bluntness Cones with Distributed Roughness in Hypersonic Flight

Reda, Daniel C.; Wilder, Michael C.; Prabhu, Dinesh K.

doi:10.2514/6.2012-446

Cited by 5 publications

(5 citation statements)

References 31 publications

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“…The grouped Re k;k curves at the time of BLT indicate that roughness may have contributed, but the magnitudes are lower than what would be expected to cause BLT based on recent data. Ballistic range testing on blunt models produced BLT from distributed roughness for Re k;k values between 150 and 300 [47]. It is plausible that larger roughness elements from the PICA tile gap filler could have contributed to BLT at some of the turbulent MISP locations.…”

Section: E Boundary-layer Transition and Roughness Effectsmentioning

confidence: 97%

“…However, no in-flight estimates of the inflight PICA roughness were possible. Figure 21 indicates that a roughness height closer to 2×10 −3 m would have been needed to reach the conservative ballistic criterion (Re k;k > 150) [47].…”

Section: E Boundary-layer Transition and Roughness Effectsmentioning

confidence: 99%

“…LAURA laminar Re k;k at MISP locations on the BET (assumed k 2 × 10 −3 m) compared with ballistic range data[47].…”

mentioning

confidence: 99%

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Mars Science Laboratory Heat Shield Aerothermodynamics: Design and Reconstruction

Edquist

Hollis

Johnston

et al. 2014

Journal of Spacecraft and Rockets

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The Mars Science Laboratory heat shield was designed to withstand a fully turbulent heat pulse using information from ground testing and computational analysis on a preflight design trajectory. Instrumentation on the flight heat shield measured in-depth temperatures to permit reconstruction of the surface heating. The data indicate that boundary-layer transition occurred at five of seven measurement locations before peak heating. Data oscillations at three pressure measurement locations may also indicate transition. This paper presents the heat shield temperature and pressure data, possible explanations for the timing of boundary-layer transition, and a comparison of reconstructed and computational heating on the actual trajectory. A smooth-wall boundary-layer Reynolds number that was used to predict transition is compared with observed transition at various heat shield locations. A single transition Reynolds number criterion does not uniformly explain the timing of boundary-layer transition observed during flight. A roughness-based Reynolds number suggests that transition due to discrete or distributed roughness elements occurred. However, the distributed roughness height from acreage heat shield material would have needed to be larger than expected. The instrumentation confirmed the predicted location of maximum turbulent heat flux near the lee-side shoulder. The reconstructed heat flux at that location is bounded by smooth-wall turbulent convective heating, indicating that a significant augmentation due to surface roughness did not occur. Turbulent heating on the downstream side of the heat shield nose exceeded smooth-wall convective levels, assuming a supercatalytic surface, indicating that roughness may have augmented heating. The stagnation region heating also exceeded calculated convective heating; the cause of elevated heating may be attributed to a combination of shocklayer radiation and a heating augmentation of unknown origin that was also evident in ground test data.m 2 ∕s D = aeroshell diameter, m H r = recovery enthalpy, J∕kg h = altitude, km h w = wall enthalpy, J∕kg k = roughness height, m k s = equivalent sand-grain roughness height, m k = roughness augmentation Reynolds number, ρ w U τ k s ∕μ w M = Mach number m = entry system mass, kg p = pressure, Earth atm; one Earth atm is equal to 101,325 Pa Q = integrated heat load, ∫ q dt, J∕cm 2 q = heat flux, W∕cm 2 q ∞ = freestream dynamic pressure, 1∕2ρ ∞ V 2 ∞ , Pa Re k;k = roughness Reynolds number, ρ k u k k∕μ k Re θ = momentum thickness Reynolds number, ρ e u e θ∕μ e Re ∞ = freestream Reynolds number, ρ ∞ V ∞ D∕μ ∞ St = Stanton number, q w ∕ρ ∞ V ∞ H ∞ − H w T = temperature,°C t = time from atmospheric interface, s U τ = shear velocity, τ w ∕ρ w p , m∕s V = velocity relative to atmosphere, km∕s α = angle of attack, deg Subscripts D = aeroshell diameter e = boundary-layer edge, where H e is equal to 0.99H ∞ i = atmospheric interface lam = laminar trim = aerodynamic trim condition turb = turbulent w = wall w, k = rough wall w, 0 = smooth wall ∞ = fr...

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Section: E Boundary-layer Transition and Roughness Effectsmentioning

confidence: 97%

Section: E Boundary-layer Transition and Roughness Effectsmentioning

confidence: 99%

See 1 more Smart Citation

Mars Science Laboratory Heat Shield Aerothermodynamics: Design and Reconstruction

Edquist

Hollis

Johnston

et al. 2014

Journal of Spacecraft and Rockets

View full text Add to dashboard Cite

show abstract

“…The grouped Re k,k curves at the time of BLT indicates that roughness may have contributed, but the magnitudes are lower than what would be expected to cause BLT based on recent data; ballistic range testing on blunt models produced BLT from distributed roughness for Re k,k values between 150 and 300. 46 It is plausible that larger roughness elements from the PICA tile gap filler could have contributed to BLT at some of the turbulent MISP locations. However, no in-flight measurements of the in-flight PICA roughness were attempted.…”

Section: E Boundary Layer Transition and Roughness Effectsmentioning

confidence: 99%

Mars Science Laboratory Heatshield Aerothermodynamics: Design and Reconstruction

Edquist¹,

Hollis

Johnston³

et al. 2013

44th AIAA Thermophysics Conference

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The Mars Science Laboratory heatshield was designed to withstand a fully-turbulent heat pulse based on test results and computational analysis on a pre-flight design trajectory. Instrumentation on the flight heatshield measured in-depth temperatures in the thermal protection system. The data indicate that boundary layer transition occurred at 5 of 7 thermocouple locations prior to peak heating. Data oscillations at 3 pressure measurement locations may also indicate transition. This paper presents the heatshield temperature and pressure data, possible explanations for the timing of boundary layer transition, and a qualitative comparison of reconstructed and computational heating on the actual trajectory. A smooth-wall boundary layer Reynolds number that is typically used to predict transition is compared to observed transition at various heatshield locations. A uniform transition Reynolds number criterion does not explain the timing of boundary layer transition observed during flight. A roughness-based Reynolds number suggests that transition due to discrete or distributed roughness elements occurred. However, the distributed roughness height from acreage heatshield material would have needed to be larger than pre-flight measurements. The instrumentation confirmed the predicted location of maximum turbulent heat flux near the leeside shoulder. The reconstructed heat flux at that location is bounded by smooth-wall turbulent convective heating, indicating that a significant augmentation due to surface roughness did not occur. Turbulent heating on the downstream side of the heatshield nose exceeded smooth-wall convective levels assuming a super-catalytic surface, indicating that roughness may have augmented heating. The stagnation region heating also exceeded calculated convective heating; the cause of elevated heating may be attributed to a combination of shock layer radiation and a heating augmentation of unknown origin that was also evident in ground test data.Mach number m entry system mass (kg) p pressure (Earth atm, 1 Earth atm = 101,325 P a) Q convective heat load, q dt (J/cm 2 ) q heating rate (W/cm 2 ) q ∞ freestream dynamic pressure,roughness Reynolds number, ρ k u k k/µ k Re θ momentum thickness Reynolds number, ρ e u e θ/µ e St Stanton number,from atmospheric interface (s) U τ shear velocity, τ w /ρ w (m/s) V velocity relative to atmosphere (km/s) α angle of attack (deg) γ inertial flight path angle (deg) emissivity θ momentum thickness (m) ρ density (kg/m 3 ) σ Stefans constant = 5.67 x 10 −8 W/(m 2 − K 4 ) τ shear stress (P a) Acronyms AEDC Arnold Engineering Development Center B-L Baldwin-Lomax BET best estimated trajectory BLT boundary layer transition CEV Crew Exploration Vehicle CFD computational fluid dynamics

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“…However, data obtained in Brazhko et al 4 clearly indicate that reversal can be observed at fixed Re 1,1 (therefore fixed flow disturbances level) and increased R. In this case, reversal still could be triggered by a bypass scenario due to the possible increase of boundary layer sensitivity to bypass with increasing radius. Unfortunately, no flight 23 or ballistic range 24 data on smooth-body reversal are available. This is often due to ablation or too small Re 1,R values realized.…”

Section: Introductionmentioning

confidence: 99%

Laminar–turbulent transition reversal on blunt ogive body of revolution at hypersonic speeds

Vaganov

Noev

Radchenko

et al. 2018

Proceedings of the Institution of Mechanical Engineers, Part G:

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The influence of flow parameters and nose radius on the location of laminar–turbulent transition is investigated. The model is an ogive-conical body of revolution having half angle about 9°. Experiments were conducted in a shock tunnel at Mach number 5. The transition location was diagnosed by the heat transfer rate distribution determined with the aid of luminescent temperature converters. It is shown that transition reversal can occur either (a) in the absence of turbulent wedges or (b) at a constant level of freestream disturbances. Both increasing and decreasing branches of Re∞,t (Re∞,R) dependency were observed at constant nose radius while varying only the unit Reynolds number.

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Transition Experiments on Large Bluntness Cones with Distributed Roughness in Hypersonic Flight

Cited by 5 publications

References 31 publications

Mars Science Laboratory Heat Shield Aerothermodynamics: Design and Reconstruction

Mars Science Laboratory Heat Shield Aerothermodynamics: Design and Reconstruction

Mars Science Laboratory Heatshield Aerothermodynamics: Design and Reconstruction

Laminar–turbulent transition reversal on blunt ogive body of revolution at hypersonic speeds

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