A viscous-shock-layer analysis of the Martian aerothermal environment

Gupta, R. N.; Moss, James N.; Sutton, Kenneth; Lee, K. P.

doi:10.2514/6.1991-1345

Cited by 18 publications

(3 citation statements)

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“…Studies of entry into Venus by Sutton 27 and Fujita et al 28,29,30 have shown about a 20% reduction in the radiative heating by ablation. For Mars entry, with an m-dot of 0.075, Gupta et al 31 predicted a radiative heating reduction of 50% with the introduction of ablation.…”

Section: Introductionmentioning

confidence: 99%

Influence of Ablation on Radiative Heating for Earth Entry

Johnston

Gnoffo

Sutton

2009

Journal of Spacecraft and Rockets

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Using the coupled ablation and radiation capability recently included in the LAURA flowfield solver, this paper investigates the influence of ablation on the shock-layer radiative heating for Earth entry. The extension of the HARA radiation model, which provides the radiation predictions in LAURA, to treat a gas consisting of the elements C, H, O, and N is discussed. It is shown that the absorption coefficient of air is increased with the introduction of the C and H elements. A simplified shock layer model is studied to show the impact of temperature, as well as the abundance of C and H, on the net absorption or emission from an ablation contaminated boundary layer. It is found that the ablation species reduce the radiative flux in the vacuum ultraviolet, through increased absorption, for all temperatures. However, in the infrared region of the spectrum, the ablation species increase the radiative flux, through strong emission, for temperatures above 3,000 K. Thus, depending on the temperature and abundance of ablation species, the contaminated boundary layer may either provide a net increase or decrease in the radiative flux reaching the wall. To assess the validity of the coupled ablation and radiation LAURA analysis, a previously analyzed Mars-return case (15.24 km/s), which contains significant ablation and radiation coupling, is studied. Exceptional agreement with previous viscous shock-layer results is obtained. A 40% decrease in the radiative flux is predicted for ablation rates equal to 20% of the free-stream mass flux. The Apollo 4 peak-heating case (10.24 km/s) is also studied. For ablation rates up to 3.4% of the free-stream mass flux, the radiative heating is reduced by up to 19%, while the convective heating is reduced by up to 87%. Good agreement with the Apollo 4 radiometer data is obtained by considering absorption in the radiometer cavity. For both the Mars return and the Apollo 4 cases, coupled radiation alone is found to reduce the radiative heating by 30 -60% and the convective heating by less than 5%. = thickness of the constant-property layers defined in Fig. 3, with i = 1 or 2 (cm) h = absorption coefficient (cm -1 ) = blowing reduction parameter = trasmissivity defined in Eq. (2) Subscripts h = indicates a spectral dependence in terms of eV Superscripts abl = refers to ablation species, meaning species containing any C or H atoms air = refers to air species, meaning species containing no C or H atoms Abbreviations eV = electron volts; the frequency in eV, labeled h , is equal to 1.24x10 -4 / c IR = infrared; refers here to the spectral region below 6 eV NIST = National Institute of Standards and Technology OP = Opacity Project VUV = vacuum ultraviolet; refers to the spectral region above 6 eV

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Section: Introductionmentioning

confidence: 99%

Influence of Ablation on Radiative Heating for Earth Entry

Johnston

Gnoffo

Sutton

2009

Journal of Spacecraft and Rockets

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“…That portion of energy which is radiated away from the vehicle and escapes the shock l a yer causes the phenomenon called radiative cooling. So far, ows with signi cant radiative cooling have been simulated only by xedshock s c hemes including the viscous-shock-layer VSL scheme 1,4 or inviscid equilibrium schemes 5 . This approximation simpli es the calculation of the coupled, radiating ow eld but may not be valid for ow situations characteristic of aerobraking reentry into the Earth's atmosphere upon return from the Moon or Mars.…”

Section: Introductionmentioning

confidence: 99%

Coupled radiation effects in thermochemical nonequilibrium shock-capturing flowfield calculations

Lin

Robert

Peter

1994

Journal of Thermophysics and Heat Transfer

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Lunar and Mars return conditions are examined using the LAURA ow eld code and the LORAN radiation code to assess the e ect of radiative coupling on axisymmetric thermochemical nonequilibrium ows. Coupling of the two codes is achieved iteratively. Special treatment required to couple radiation in a shock-capturing method is discussed. Results indicate that while coupling e ects are generally the same as occur in equilibrium ows, under certain conditions radiation can modify the chemical kinetics of a nonequilibrium ow and thus alter relaxation processes. Coupling e ects are found to be small for all cases considered, except for a ve meter diameter aerobrake returning from Mars at 13.6 km sec.

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“…The stagnation point pressure is 0.064 atm. Maximum stagnation point pressure occurs at t= 91.5 s in the undershoot trajectory with a value of 0.094 atm.An independent c heck on the stagnation point maximum heating was performed by using a Mars atmosphere version of the Viscous Shock Layer VSL code 16 . This method accounts for thermochemical nonequilibrium conditions in the shock layer and includes a 16 species kinetics model of which the 8 species model used in LAURA is a subset.…”

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confidence: 99%

Aerothermal Heating Predictions for Mars Microprobe

Mitcheltree

DiFulvio

Horvath

et al. 1999

Journal of Spacecraft and Rockets

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A combination of computational predictions and experimental measurements of the aerothermal heating expected on the two Mars Microprobes during their entry to Mars are presented. The maximum, non-ablating, heating rate at the vehicle's stagnation point at = 0 0 is predicted for an undershoot trajectory to be 194 W=cm 2 with associated stagnation point pressure of 0.064 atm. Maximum stagnation point pressure occurs later during the undershoot trajectory and is 0.094 atm. From computations at seven overshoot-trajectory points, the maximum heat load expected at the stagnation point is near 8800 J=cm 2 . Heat rates and heat loads on the vehicle's afterbody are much lower than the forebody. At zero degree angle-of-attack, heating over much of the hemispherical afterbody is predicted to be less than 2 percent of the stagnation point v alue. Good qualitative agreement is demonstrated for forebody and afterbody heating between CFD calculations at Mars entry conditions and experimental thermographic phosphor measurements from the Langley 20-Inch Mach 6 Air Tunnel. A n o v el approach which incorporates six degree-of-freedom trajectory simulations to perform a statistical estimate of the e ect of angle-of-attack, and other o -nominal conditions, on heating is included. Nomenclature B = Ballistic coe cient, kg=m 2 C h = heat transfer coe cient M = Mach n umber P = pressure, atm q = heat rate, W=cm 2 R n = nose radius, m s = surface distance from geometric stagnation point, m t = independent v ariable time, s V = v elocity, m=s x; z = independent spatial dimensions, m = angle-of-attack, deg = side-slip angle, deg = density, kg=m 3 = standard deviation IntroductionWhen the Mars Surveyor 98 Lander is launched in January of 1999, it will transport not only its own lander to Mars, but two small soil penetrators. These Aerospace Engineer, Aerothermodynamics Branch, Aeroand Gas-Dynamics Division, NASA Langley Research Center, Senior member AIAA.y Aerospace Engineer, Aerothermodynamics Branch, Aeroand Gas-Dynamics Division, NASA Langley Research Center, Senior member AIAA.z Aerospace Engineer, Aerothermodynamics Branch, Aeroand Gas-Dynamics Division, NASA Langley Research Center, Senior member AIAA.x Aerospace Engineer, Vehicle Analysis Branch, Space Systems and Concepts Division, NASA Langley Research Center, Member AIAA.Copyright c 1998 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17, U.S. Code. The U.S. Government has a r o y alty-free license to exercise all rights under the copyright claimed herein for governmental purposes. All other rights are reserved by the copyright o wner.two Mars Microprobes 1 are the second of the Deep Space missions from NASA's New Millennium Program O ce. Upon arrival at Mars, the penetrators will be released from the cruise stage and begin a free fall to the surface. This paper focuses on predicting the convective heating which the aeroshells will encounter during the hypersonic portion of that Mars entry. Knowledge of the...

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A viscous-shock-layer analysis of the Martian aerothermal environment

Cited by 18 publications

References 4 publications

Influence of Ablation on Radiative Heating for Earth Entry

Influence of Ablation on Radiative Heating for Earth Entry

Coupled radiation effects in thermochemical nonequilibrium shock-capturing flowfield calculations

Aerothermal Heating Predictions for Mars Microprobe

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