Due to the Columbia Space Shuttle Accident of February 2003, the Columbia Accident Investigation Board determined the need for an on-orbit inspection system for the Thermal Protection System that accurately determines damage depth to 0.25". NASA contracted the Spacecraft Technology Center in College Station, Texas, for a proof-of-concept photogrammetric system. This system involves a high quality digital camera placed on the International Space Station, capable of taking high fidelity images of the orbiter as it rotates through the Rendezvous Pitch Maneuver. Due to the pitch rotation, the images are tilted at different angles. The tilt causes the damage to exhibit parallax between multiple images. The tilted images are therefore registered to the near-vertical images using visually striking features on the undamaged surface of the Thermal Protection System that appear in multiple images taken at different tilt angles. The images become relatively oriented after registration, and features in one image are ensured to lie on the epipolar line in the other images. Features that do not lie on the undamaged surface, however, are shifted in the tilted images. These pixels are matched to the near-vertical image using a sliding-window area-matching approach. The windows are matched using a least-squares error method. The change in location for a pixel in a tilted image from its expected location on the undamaged surface is called the pixel disparity. This disparity is linearly scaled using the tilt angle and the pixel sampling to determine the depth of the damage at that pixel location. The algorithm is tested on a set of damaged tiles at the Johnson Space Center in Houston and the photogrammetric damage depth is then compared to a set of truth data provided by NASA. The photogrammetric method shows promise, with the 0.25" error limit being exceeded in only a few pixel locations. Once the camera properties are fully known from calibration, this systematic error should be reduced.

Inductively Coupled Plasma (ICP) facilities create high enthalpy ows to recreate atmospheric entry conditions. Although no condition has been duplicated exactly in a ground test facility, it is important to characterize the condition to understand how close a facility can come to doing so. An ICP facility was constructed at the University of Vermont for aerospace material testing in 2010. The current setup can operate using air, carbon dioxide, nitrogen, and argon to test samples in a chamber. In this work we investigate di erent ways to increase measured heat ux and expand our facility to operate supersonically. To do so, a water cooled injection system was designed to overcome failure points of the prior system. An investigation of heat ux methods that provide a baseline for the facility were also examined and tested. A nozzle con guration was also developed with an overall goal of increasing the plasma ow to reach sonic and supersonic velocities, allowing it to be compared with the existing subsonic system. An iterative approach was taken to develop a nozzle design that is robust enough to handle the harsh environment, yet adaptable to the pre-existing facility components. The current design uses interchangeable sonic and supersonic nozzles which also allow for appropriate plasma gas expansion. Data are taken through retractable and goose-neck probe sample holders during testing. Heat ux can be determined by use of a Gardon gage, slug calorimeter, and water cooled calorimeter. Total and static pressure are determined from a pitot tube and pressure tap, which are then manipulated into a velocity measurement. A comparison between subsonic and supersonic operation is then made with these data. Existing literature uses correlations between jet diameter and velocity gradients to determine the e ective heat ux. This investigation found that the experimental and theoretical heat ux results scale correctly according to the correlations.

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Flexible thermal protection systems are of interest due to their necessity for the success of future atmospheric entry vehicles. Current non-ablative flexible designs incorporate a two-dimensional woven fabric on the leading surface of the vehicle. The focus of this research investigation was to characterize the aerothermal performance of silicon carbide fabric using the 30 kW Inductively Coupled Plasma Torch located at the University of Vermont. Experimental results have shown that SiC fabric test coupons achieving surface temperatures between 1000°C and 1500°C formed an amorphous silicon dioxide layer within seconds after insertion into air plasmas. The transient morphological changes that occurred during oxidation caused a time dependence in the gas / surface interactions which may detrimentally affect the in-flight performance. Room temperature tensile tests of the SiC coupons have shown a rapid strength loss for durations less than 240 seconds due to oxidation. Catastrophic failure and temperature spikes were observed on almost all SiC coupons when exposed to air plasmas at heat fluxes above 80 W/cm2. Interestingly, simulation of entry into the Mars atmosphere using a carbon dioxide plasma caused a material response that was vastly different than the predictable silica layer observed during air plasma exposure.

The objective of this thesis is to predict shape change due to ablation and to find temperature distribution of the thermal protection system of a supersonic vehicle under aerodynamic heating by using finite element method. A subliming ablative is used as thermal protection material. Required material properties for the ablation analyses are found by using DSC (Differential Scanning Calorimetry) and TGA (Thermogravimetric Analysis) thermal analysis techniques. DSC is a thermal analysis technique that looks at how a material&#039 / s specific heat capacity is changed by temperature and TGA is a technique in which the mass of a substance is monitored as a function of temperature. Moreover, oxyacetylene ablation tests are conducted for the subliming ablative specimens and measured recession values are compared with the analytically calculated values. Maximum difference between experimental results and analytical results is observed as 3% as seen in Table 7. For the finite element analyses, ANSYS Software is used. A numerical algorithm is developed by using programming language APDL (ANSYS Parametric Design Language) and element kill feature of ANSYS is used for simulation of ablation process. To see the effect of mesh size and time step on the solution of analyses, oxyacetylene test results are used. Numerical algorithm is also applied to the blunt-nosed section of a supersonic rocket which is made from subliming ablative material. Ablation analyses are performed for the nose section because nose recession is very important for a rocket to follow the desired trajectory and nose temperature is very important for the avionics in the inner side of the nose. By using the developed algorithm, under aerodynamic heating, shape change and temperature distribution of the nose section at the end of the flight are obtained. Moreover, effects of ablation on the trajectory of the rocket and on the flow around the rocket are examined by Missile DATCOM and CFD (computational fluid dynamics) analysis tools.

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The Space Shuttle Columbia accident and the recent excitement surrounding Discovery's return to space brought excessive media attention to the foam products used on the External Tank (ET). In both cases, videos showed chunks of foam or ablative material falling away from the ET during lift off. This led to several years of investigation and research into the exact cause of the accident and potential solutions to avoid the problem in the future. Several design changes were made prior to the return to flight this year, but the ET still shed foam during lift off. Since the Columbia accident, the loss of foam on ETs has been a significant area of interest for NASA, United Space Alliance, and Lockheed Martin. The Columbia Accident Investigation Board did not evaluate alternative materials but certainly highlighted the need for change. The majority of the research previously concentrated on improving the design and/or the application process of the current materials. Within recent years, some research and testing has been done to determine if a glass microsphere composite foam would be an acceptable alternative, but this work was overcome by the need for immediate change to return the shuttle to flight in time to deliver supplies to the International Space Station. Through a better understanding of the foam products currently used on the ET, other products can be evaluated for future space shuttle flights and potential applications on new space vehicles. The material properties and the required functionality of alternative materials can be compared to the current materials to determine if suitable replacement products exist. This research also lends itself to the development of future space flight and unmanned launch vehicles. In this paper, the feasibility of alternative material for the space shuttle's external tank will be investigated. Research on what products are used on the ET and a set of functional requirements driving the selection of those materials will be presented. The material properties of the current ET foam products will be collected and an evaluation of how those materials' properties meet the functional requirements will be accomplished. Then significant research on polymeric foams and ablative materials will be completed to learn how these various products can be applied in this industry. With this research and analysis, the knowledge gained will be used to select and evaluate the effectiveness of an alternate product and to determine feasibility of a product change with the current ET and the importance of maintaining the shuttle launch schedule. This research will also be used to evaluate the potential application of the alternative product on future platforms. There are several possible outcomes to this research. This research could result in a recommended change to the ET foam material or a perfectly acceptable alternative material that could result in a cost or schedule impact if implemented. It is also possible that there exists no suitable alternative material given the existing functional requirements. In any case, the alternative material could have future applications on new space vehicles. A set of results from the research and analysis will be provided along with a recommendation on a future material for use on space vehicles.

This study researched the history of Space Shuttle Reusable Surface Insulation which was designed and developed for use on the United States Orbiter fleet to protect from the high heating experienced during reentry through Earth's atmosphere. Specifically the tile system which is attached to the structure by the means of an RTV adhesive has experienced situations where the bonds are identified as subnominal. The history of these subnominal conditions is presented along with a recent identification of a subnominal bond between the Strain Isolation Pad and the tile substrate itself. Tests were run to identify the cause of these subnominal conditions and also to show how these conditions were proved to be acceptable for flight. The study also goes into cases that could be used to identify subnominal conditions on tile as a non-destructive test prior to flight. Several options of non-destructive testing were identified and recommendations are given for future research into this topic. A recent topic is also discussed in the instance where gap fillers were identified during the STS-114 mission that did not properly adhere to the substrate. The gap fillers were found protruding past the Outer Mold Line of the vehicle which required an unprecedented spacewalk to remove them to allow for a safe reentry through the atmosphere.

During hypersonic atmospheric entry, spacecraft are exposed to enormous aerodynamic heat. To prevent the payload from overheating, charring ablative materials are favored to be applied as the heat shield at the exposing surface of the vehicle. Accurate modeling not only prevents mission failures, but also helps reduce cost. Existing models were mostly limited to one-dimensional and discrepancies were shown against measured experiments and flight-data. To help improve the models and analyze the charring ablation problems, a multidimensional material response module is developed, based on a finite volume method framework. The developed computer program is verified through a series of test-cases, and through code-to-code comparisons with a validated code. Several novel models are proposed, including a three-dimensional pyrolysis gas transport model and an orthotropic material model. The effects of these models are numerically studied and demonstrated to be significant.

Space vehicles, while entering the planetary atmosphere, experience high loads of heat. Ablative materials are commonly used for a thermal protection system, which undergo mass removal mechanisms to counter the heat rates. Spallation is one of the ablative processes, which is characterized by the ejection of solid particles from the material into the flow. Numerical codes that are used in designing the heat shields ignore this phenomenon. Hence, to evaluate the effectiveness of spallation phenomenon, a numerical model is developed to compute the dynamics and chemistry of the particles. The code is one-way coupled to a CFD code that models high enthalpy flow field around a lightweight ablative material. A parametric study is carried out to examine the variations in trajectories with respect to ejection parameters. Numerical results are presented for argon and air flow fields, and their effect on the particle behavior is studied. The spallation code is loosely coupled with the CFD code to evaluate the impact of a particle on the flow field, and a numerical study is conducted.