Stardust backshell and back interface plate design verification tests in the NASA Ames Arc Jet facilities

Johnson, Catherine C.; Tran, Huy K.; Smith, Marnell; Dill, H.H.

doi:10.2514/6.1997-2483

Cited by 10 publications

(3 citation statements)

References 1 publication

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“…The forebody heat shield of the SRC was manufactured as a single piece, the material of which was made of phenolic impregnated carbon ablator (PICA); 19 the after body is covered with silicone elastomeric charring ablator (SLA-561V). 20 In this study, the heat shield material selected in the simulation is TACOT (Theoretical Ablative Composite for Open Testing). 21 TACOT is a fictitious material that based on PICA, and it has very similar performance compared to PICA.…”

Section: Iia Heat Shield Materialsmentioning

confidence: 99%

Numerical Investigation of Geometric Effects of Stardust Return Capsule Heat Shield

Weng

Martin

2015

53rd AIAA Aerospace Sciences Meeting

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Sample geometry is very influential in small charring ablative articles where 1D assumption might not be accurate. In heat shield design, 1D is often assumed since the nose radius is much larger than the thickness of charring. Whether the 1D assumption is valid for the heat shield is unknown. Therefore, the geometric effects of Stardust sample return capsule heat shield are numerically studied using a material response program. The developed computer program models material charring, conductive heat transfer, surface energy balance, pyrolysis gas transport and orthotropic material properties in 3D Cartesian coordinates. Simulation results show that the centerline temperatures predicted by 3D model are quite close to 1D model at the surface, but not the case inside the material. The pyrolysis surface gas blowing behaviors are quite similar but differences are observed at later time. Orthotropic model predicted a very different heat shield response to both the isotropic model and the 1D model.

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Section: Iia Heat Shield Materialsmentioning

confidence: 99%

Numerical Investigation of Geometric Effects of Stardust Return Capsule Heat Shield

Weng

Martin

2015

53rd AIAA Aerospace Sciences Meeting

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“…These penetrations of the heatshield require extensive analysis and testing to validate their design. 2 Removable doors are usually required to allow access to the payload after integration of the aeroshell. It is desirable to integrate energetic hardware such as Radioisotope Thermoelectric Generators (RTG's) at the last possible moment, even as late as when the payload and spacecraft are stacked on the launch vehicle.…”

Section: Introductionmentioning

confidence: 99%

Inflatable Re-entry Vehicle Experiment (IRVE) Design Overview

Hughes

Dillman

Starr

et al. 2005

18th AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar

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Inflatable aeroshells offer several advantages over traditional rigid aeroshells for atmospheric entry. Inflatables offer increased payload volume fraction of the launch vehicle shroud and the possibility to deliver more payload mass to the surface for equivalent trajectory constraints. An inflatable's diameter is not constrained by the launch vehicle shroud. The resultant larger drag area can provide deceleration equivalent to a rigid system at higher atmospheric altitudes, thus offering access to higher landing sites. When stowed for launch and cruise, inflatable aeroshells allow access to the payload after the vehicle is integrated for launch and offer direct access to vehicle structure for structural attachment with the launch vehicle. They also offer an opportunity to eliminate system duplication between the cruise stage and entry vehicle. There are however several potential technical challenges for inflatable aeroshells. First and foremost is the fact that they are flexible structures. That flexibility could lead to unpredictable drag performance or an aerostructural dynamic instability. In addition, durability of large inflatable structures may limit their application. They are susceptible to puncture, a potentially catastrophic insult, from many possible sources. Finally, aerothermal heating during planetary entry poses a significant challenge to a thin membrane. NASA Langley Research Center and NASA's Wallop's Flight Facility are jointly developing inflatable aeroshell technology for use on future NASA missions. The technology will be demonstrated in the Inflatable Re-entry Vehicle Experiment (IRVE). This paper will detail the development of the initial IRVE inflatable system to be launched on a Terrier/Orion sounding rocket in the fourth quarter of CY2005. The experiment will demonstrate achievable packaging efficiency of the inflatable aeroshell for launch, inflation, leak performance of the inflatable system throughout the flight regime, structural integrity when exposed to a relevant dynamic pressure and aerodynamic stability of the inflatable system. Structural integrity and structural response of the inflatable will be verified with photogrammetric measurements of the back side of the aeroshell in flight. Aerodynamic stability as well as drag performance will be verified with on board inertial measurements and radar tracking from multiple ground radar stations. The experiment will yield valuable information about zero-g vacuum deployment dynamics of the flexible inflatable structure with both inertial and photographic measurements. In addition to demonstrating inflatable technology, IRVE will validate structural, aerothermal, and trajectory modeling techniques for the inflatable. Structural response determined from photogrammetrics will validate structural models, skin temperature measurements and additional in-depth temperature measurements will validate material thermal performance models, and on board inertial measurements along with radar tracking from multiple ground radar stations will valid...

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“…2ङ arc-jet testing for aeight qualiaecation of the heatshield, 152,153 3ङ predicting the aerodynamic stability of the vehicle during entry, 154,155 and 4ङ estimating the landing footprint for vehicle recovery. 156 The aeow e n vironment for the high speed Stardust entry is complex; the sample return capsule heatshield is ablating and the aeow is radiating.…”

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confidence: 99%

Computational aerothermodynamic design issues for hypersonic vehicles

Olynick

Venkatapathy

1997

32nd Thermophysics Conference

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A brief review of the evolutionary progress in computational aerothermodynamics is presented. The current status of computational aerothermodynamics is then discussed, with emphasis on its capabilities and limitations for contributions to the design process of hypersonic vehicles. Some topics to be highlighted include: घ1ङ aerodynamic coeaecient predictions with emphasis on high temperature gas eaeects; घ2ङ surface heating and temperature predictions for thermal protection system घTPSङ design in a high temperature, thermochemical nonequilibrium environment; घ3ङ methods for extracting and extending computational aeuid dynamic घCFDङ solutions for eaecient utilization by all members of a multidisciplinary design team; घ4ङ physical models; घ5ङ validation process and error estimation; and घ6ङ gridding and solution generation strategies. Recent experiences in the design of X-33 will be featured. Computational aerothermodynamic contributions to Mars Pathaender, METEOR, and Stardust घComet Sample returnङ will also provide context for this discussion. Some of the barriers that currently limit computational aerothermodynamics to a predominantly reactive mode in the design process will also be discussed, with the goal of providing focus for future research.

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Stardust backshell and back interface plate design verification tests in the NASA Ames Arc Jet facilities

Cited by 10 publications

References 1 publication

Numerical Investigation of Geometric Effects of Stardust Return Capsule Heat Shield

Numerical Investigation of Geometric Effects of Stardust Return Capsule Heat Shield

Inflatable Re-entry Vehicle Experiment (IRVE) Design Overview

Computational aerothermodynamic design issues for hypersonic vehicles

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