Engineering topics

Kontinos, Dean

Publications and source records attributed to Kontinos, Dean.

5 records

NASA NTRS2008

In-Field Diffuse Ultraviolet Spectroscopy and Imaging of the Stardust Sample Return Capsule

In-field diffuse Ultraviolet (UV) spectroscopy and imaging systems were developed for the purposes of evaluating the surface chemical composition of spacecraft thermal control coatings and materials. The investigation of these systems and the compilation of an associated UV reflectance and luminescence database were conducted using the Stardust Sample Return Capsule (SRC), located at the Johnson Space Center. Spectral responses of the surfaces of the Stardust forebody and aftbody in both reflectance and fluorescence modes were examined post-flight. In this paper, we report on two primary findings of in-field diffuse UV spectroscopy and imaging: (1) deduction of the thermal history of thermal control coatings of the forebody and (2) bond line variations in the aftbody. In the forebody, the thermal history of thermal control coatings may be deduced from the presence of particular semiconducting defect states associated with ZnO, a common emissivity constituent in thermal control coatings. A spatial dependence of this history was mapped for these regions. In the aftbody, luminescing defect states, associated with Si and SiO2 color centers were found along regions of bond variability.

Pugel, D. Elizabeth↗

NASA NTRS2001

Mechanical Properties and Microstructural Characterization of Particulate Reinforced Diboride Composites for High Temperature Leading Edge Applications

Previous work on refractory diboride composites has shown that these systems have the potential for use in high temperature leading edge applications for reusable reentry vehicles. Experiments in reentry environments have shown that these materials have multiple use temperatures greater than 1900 C. The work to be discussed focuses on three compositions: HfB2/SiC, ZrB2/SiC, and ZrB2/C/SiC. These composites have been hot pressed and their mechanical properties measured at room and elevated temperatures. Extensive microstructural characterization has been conducted on polished cross sections and the fracture surfaces have been examined to determine their failure origins.

Ellerby, Donald T.↗

NASA NTRS2001

Thermal Protection Materials for Reentry Applications

Thermal protection materials and systems (IRS) are used to protect spacecraft during reentry into Earth's atmosphere or entry into planetary atmospheres. As such, these materials are subject to severe environments with high heat fluxes and rapid heating. Catalytic effects can increase the temperatures substantially. These materials are also subject to impact damage from micrometeorites or other debris during ascent, orbit, and descent, and thus must be able to withstand damage and to function following damage. Thermal protection materials and coatings used in reusable launch vehicles will be reviewed, including the needs and directions for new materials to enable new missions that require faster turnaround and much greater reusability. The role of ablative materials for use in high heat flux environments, especially for non-reusable applications and upcoming planetary missions, will be discussed. New thermal protection system materials may enable the use of sharp nose caps and leading edges on future reusable space transportation vehicles. Vehicles employing this new technology would have significant increases in maneuverability and out-of-orbit cross range compared to current vehicles, leading to increased mission safety in the event of the need to abort during ascent or from orbit. Ultrahigh temperature ceramics, a family of materials based on HfB2 and ZrB2 with SiC, will be discussed. The development, mechanical and thermal properties, and uses of these materials will be reviewed.

Johnson, Sylvia M.↗

NASA NTRS1998

Aerothermal Effects of Cavities and Protuberances for High-Speed Sample Return Capsules

Extraterrestrial sample return is a growing component of solar system exploration. Currently, four missions, Stardust, 1 Muses-C, 2 Genesis, and Mars Sample Return, are under development that employ sample return as a prime component of the mission architecture. Respectively, these missions will return samples from the tail of a comet, an asteroid, the solar wind, and, Mars. An important component of these missions and the focus of this paper is the design of the sample return capsule (SRC). The purpose of the SRC is to safely return to Earth any gathered samples for terrestrial analysis. The two major design constraints for any SRC are as follows: 1) it must be able to survive a high-speed Earth entry (11 km/s to as a high as 15 km/s), 2) the mass of the SRC must be as small as possible. Because the SRC mass is carried from Earth to the sample sight and back, the SRC mass is a strong driver in the mission mass budget. Further, for the Mars Sample Return Capsule, planetary protection is another constraint. For this capsule, the probability of planetary contamination at Earth due to an SRC failure at entry must be minimal. For an SRC, a possible failure mechanism is severe local heating as a result of cavities and or protuberances in the SRC forebody heatshield. For example, the Apollo Command Module had a number of cavities and protuberances as part of the baseline designs Wind-tunnel tests of models containing small cavities and protuberances showed severe local heating augmentations in the vicinity of these surface discontinuities.4-5 As another example, the Genesis SRC forebody heat-shield contains penetrations (cavities) to mount the vehicle to the carrier bus. It is expected that these penetrations will also experience a severe local heating environment. A concern is that the large thermal gradients may produce sufficient thermal stress to cause local mechanical failure of the heatshield. Penetrations to the forebody heat-shield can also result from damage at vehicle integration, during launch, or during transportation of the sample return capsule from earth to the sample site and back. For example, the Starting SRC was damaged near the shoulder during the heatshield integration process producing a local surface discontinuity. Also, the Starting SRC traverses through the tail of a comet and is in space for 7 years. Thus, damage to the heatshield as a result of micrometeroid impact is a concern. Finally, it is difficult to characterize the effects of these potential heatshield singularities with ground-test facilities. Either detailed simulation or a dedicated flight test is required.

Olynick, David↗

NASA NTRS1997

The Effect of Metallic TPS Panel Bowing on the Surface Heating of the X-33 Vehicle

The thermal protection system of the windward surface of the X-33 vehicle consists of metallic honeycomb sandwich panels. Thermal gradients experienced during the descent phase of the trajectory result in a different rate of thermal expansion between the inner and outer face sheets of the metallic panels. This causes the panels to bow outward when the temperature of the outer face sheet is larger than that of the inner face sheet and inward when the temperature of the outer face sheet is less than that of he inner face sheet. This results in a quilted-type body surface. Using computational fluid dynamic analysis, this study will determine the effect the metallic TPS panel bowing has on the surface heating.

Palmer, Grant↗