Modeling Hypervelocity Impact of Reinforced Carbon-Carbon Composite Thermal Protection System

2019 ◽  
Author(s):  
Alexander J. Carpenter ◽  
Sidney Chocron ◽  
James D. Walker
Keyword(s):  
Thermal Protection ◽  
Space Shuttle ◽  
Carbon Composite ◽  
Hypervelocity Impact ◽  
Protection System ◽  
Structural Stiffness ◽  
Hypervelocity Impacts ◽  
Nose Cone ◽  
Space Shuttle Columbia ◽  
Very High

Abstract Reinforced carbon-carbon (RCC) composite is used in applications where structural stiffness and strength must be maintained at very high temperatures that may reach 2000°C or more. For example, it was used on both the Space Shuttle’s nose cone and the leading edges of its wings. As exemplified by the Space Shuttle Columbia accident, the ability of these materials to survive impacts up to hypervelocity speeds can be critical for some applications. As computational modeling becomes an increasingly important component of the design process, the ability to accurately model RCC materials under impact conditions likewise becomes more and more important. This paper describes a computational model of the thermal protection used on the Space Shuttle orbiter. The model incorporates both the RCC comprising much of the protection system and its silicon carbide coating. The model was subjected to hypervelocity impacts with both steel and aluminum projectiles, and the results were compared to test data from the literature.

Manufacturing of Inconel 718 Based Honeycomb Panels for Metallic Thermal Protection Systems

2012 ◽  
Vol 710 ◽  
pp. 197-202 ◽  
Author(s):  
Hanamantray Baluragi ◽  
V. Anil Kumar ◽  
K. Narasaiah ◽  
Shibu Gopinath ◽  
P.P. Sinha
Keyword(s):  
Thermal Protection ◽  
Space Shuttle ◽  
Thermal Protection System ◽  
Protection System ◽  
Front Wall ◽  
Honeycomb Core ◽  
Thermal Protection Systems ◽  
Wide Range ◽  
Protection Systems ◽  
Three Point Bend Test

Metallic thermal protection system (MTPS) offers significant improvements over the ceramic based TPS for reentry applications. Space shuttle refurbishment time is estimated to be around 17000 man hours between flights. Metallic based TPS can be fabricated easily and provides wide range of design options for TPS. Adaptability and robustness of metallic thermal protection systems offers the potential for reusability. In this work, a unique manufacturing process has been evolved to realize light weight honeycomb panels through corrugation, laser welding and diffusion brazing of faceplates, where in 50 micron thick Inconel718 foil is used for making honeycomb core and 0.2mm thick Inconel718 foil as faceplates. The compression and three point bend test on these panels have shown no debond between faceplates and honeycomb core. 150x150x5mm size honeycomb panels were coated with YSZ and NiCrAlY based Thermal Barrier Coatings (TBC) and high temperature tests have shown thermal resistance of around 570 °C with front wall temperature of 1186 °C and back wall of 533 °C. Also these panels have been characterized for reusability by the testing of same panel at different heat flux levels. Though it is found that honeycomb panel has shown its integrity without debond a certain acceptable level of degradation in coating is observed. Thus Inconel718 based honeycomb panels with TBC coating are proved for use as thermal protection system for reusable launch vehicle systems.

Aerothermal Analysis of Small Payloads Delivered Into Low Earth Orbit From an Airborne Launch Platform

2008 ◽  
Author(s):  
Martin J. Guillot ◽  
Ian McNab
Keyword(s):  
Circular Orbit ◽  
Thermal Protection ◽  
Raw Materials ◽  
Space Station ◽  
Carbon Composite ◽  
Low Earth Orbit ◽  
Thermal Protection System ◽  
Protection System ◽  
Earth Orbit ◽  
Stable Circular Orbit

In recent years there has been an ever increasing need to launch small payloads (∼1–100 kg) into low earth orbit (LEO). Examples include the defense and telecommunications industries. Permanent human presence in LEO, such as the international space station, requires continual re-supply from earth. Additionally, NASA’s stated mission of launching a manned mission to Mars requires many tonnes of raw materials to be economically launched into LEO and assembled there. Conventional rocket launch from earth is prohibitively expensive for small mass payloads. Estimates range from $7000–$20,000 to launch 1 kg of mass into low earth orbit. Several concepts have been proposed to economically launch small payloads from earth, including light gas guns, electromagnetic launchers and the so called “slingatron” concept. The goal of these concepts is to reduce the cost per kg (to under $1000) to achieve LEO. Each of these concepts involves launching small payloads that traverse the atmosphere and then placed into low earth orbit using thrusters to turn the velocity vector into a stable circular orbit. As the launch vehicle traverses the dense lower portion of the atmosphere it experiences severe thermal heating loads that must be absorbed by a thermal protection system (TPS) if the payload is to survive the transit. The University of Texas is currently heading a multi-university research initiative (MURI) to study the feasibility of launching small payloads into low earth orbit from an electromagnetic gun housed in an airborne platform. As part of the study, the aerothermal issues associated with traverse through the atmosphere and propellant mass required to achieve a stable circular orbit are investigated. The effort focuses on quantifying the required parasitic mass of the thermal protection system (TPS) and propellant need to place a nominal 10 kg launch mass into a circular low earth orbit from an electromagnetic launcher at 16 km altitude. The TPS is assumed to be graphite or carbon-carbon composite. In this effort, we consider ballistic trajectories only. Circular orbit is achieved using rocket thrusters at the terminal altitude. Total parasitic mass (TPS + propellant) is estimated for various launch angles.

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