Thermal Protection System Tiles Research Articles

Uncertainty-Based Comprehensive Optimization Design for the Thermal Protection System of Hypersonic Wing Structure

Due to the inherent uncertainties in material properties, loads, geometric dimensions, et al., the uncertainty-based optimization design method has become increasingly important for the design of the thermal protection system (TPS) by carefully considering the influences of uncertainties. In this study, an uncertainty-based comprehensive optimization design method, which sequentially performs the robust design of aerodynamic shape and structure size for the TPS of a hypersonic wing is proposed, on the presence of uncertain-but-bounded parameters. The robust design of the TPS’s aerodynamic shape is firstly carried out. The results show that the proposed method decreases the fluctuation of the lift-to-drag ratio by 5.7%, with a small increase of heat flux on the stagnation point by only 0.13% when compared with the conventional deterministic optimization method. After that, based on the optimized aerodynamic shape and heating loads, the robust design of the multilayer TPS tile is conducted. The results show that the mass of the TPS tile efficiently deceased from 2.713 kg to 2.445 kg by 9.89%, and the robustness of the optimized design is better than the initial design. Finally, the effectiveness of the proposed optimization method is validated by the heat insulting experiment of the typical multilayer TPS tiles.

Optimal Sensor Placement of Health Management for Thermal Protection System on RLV

When a Reusable Launch Vehicle (RLV) travels through the atmosphere there should be strong aero heating around it, so a Thermal Protection System (TPS) would be used to protect the vehicle structures from aero thermal ablation. However, the TPS structure may fail more frequently than normal vehicle structures because of the severe environment and load conditions, while a TPS failure can even lead to a catastrophic breakdown of vehicle structures. To keep the RLV safe, the TPS must be monitored real-time. To learn the heat transfer inside a ceramic TPS tile finite element models are built, and computation made with them indicates that temperature monitoring and frequency monitoring of the TPS structure should be effectual for fault detection. Health management for TPS is one of the key technologies for space vehicle. The paper puts forward a method that analyze and valuation the system states by monitoring the key function index sign of TPS. The concept of health management for Thermal Protection System of space shuttle and the principles of health monitoring are introduced. The principles of sensor selection and layout optimization method for TPS Health Management are discussed.

Ablative Heating Technology in Hypersonic Re-entry Vehicles and Cruise Aircrafts

The aim of researches conducted in thermal protection systems in aeronautics and astronautics field of engineering is to generally defend the craft from high heat loads during operation while operating at hypersonic regimes in air and space. The motive of the following composition is to draft a review analysis on ablative heating materials as thermal protective equipment on reusable planetary/atmospheric re-entry vehicles such as a space shuttle, an inter-continental ballistic missile, or a hypersonic cruise missile. The heat liberation can cause much damage to the aircraft/spacecraft whilst operation which is generally beyond repair. It is therefore of utmost importance to research multiple strategy to reduce the effect of shockwaves damage to spacecraft/aircraft materials. We shall initiate the analysis by mentioning some re-usable tile thermal protection system types such as high temperature reusable surface insulation tiles (H.R.S.I), fibrous refectory composite insulation tiles (F.R.C.I), low temperature reusable surface insulation tiles (L.R.S.I) and gradually move on to ablative thermal protection systems with the advent of reinforced carbon-carbon’s application in astronautics and aeronautics respectively.

The Understanding of Thermomechanical Analysis of a Thermal Protection System

Nowadays, reusable launch vehicle (RLV) has become a research focus of the transportation system between ground and space and the space weapon system. RLV plays an important role in controlling the cost of space transportation and performing the orbital mission. Since RLV would suffer from the aerodynamic heating inevitably during reentry, the thermal protection system (TPS) is designed to prevent too much heat transmitting to the vehicle structure and maintain the vehicle structure below a specified temperature limit. Several studies were performed to develop an understanding of not only the thermal and structural analysis of ceramic tile thermal protection system on the space shuttle but also the controlling factors of TPS. The TPS is subjected to the reentry heating and pressure profile of the Access to Space vehicle, and the transient temperature distribution and the resultant thermal stresses in the system are computed. Comparisons between various studies based on different assumptions were examined. By comparing these results with more realistic ones, the differences are evaluated. Results suggest that the TPS analysis must be based on reasonable and realistic parameters. Thus, engineers have to keep in mind that all parameters should be chosen very carefully to achieve results that close to practical ones.

Spray-on foam insulations for launch vehicle cryogenic tanks

Spray-on foam insulation (SOFI) has been developed for use on the cryogenic tanks of space launch vehicles beginning in the 1960s with the Apollo program. The use of SOFI was further developed for the Space Shuttle program. The External Tank (ET) of the Space Shuttle, consisting of a forward liquid oxygen tank in line with an aft liquid hydrogen tank, requires thermal insulation over its outer surface to prevent ice formation and avoid in-flight damage to the ceramic tile thermal protection system on the adjacent Orbiter. The insulation also provides system control and stability throughout the lengthy process of cooldown, loading, and replenishing the tank. There are two main types of SOFI used on the ET: acreage (with the rind) and closeout (machined surface). The thermal performance of the seemingly simple SOFI system is a complex array of many variables starting with the large temperature difference of 200–260K through the typical 25-mm thickness. Environmental factors include air temperature and humidity, wind speed, solar exposure, and aging or weathering history. Additional factors include manufacturing details, launch processing operations, and number of cryogenic thermal cycles. The study of the cryogenic thermal performance of SOFI under large temperature differentials is the subject of this article. The amount of moisture taken into the foam during the cold soak phase, termed Cryogenic Moisture Uptake, must also be considered. The heat leakage rates through these foams were measured under representative conditions using laboratory standard liquid nitrogen boiloff apparatus. Test articles included baseline, aged, and weathered specimens. Testing was performed over the entire pressure range from high vacuum to ambient pressure. Values for apparent thermal conductivity and heat flux were calculated and compared with prior data. As the prior data of record was obtained for small temperature differentials on non-weathered foams, analysis of the different methods is provided. Recent advancements and applications of SOFI systems on future launch vehicles and spacecraft are also addressed.

Thermomechanical Behavior of a Damaged Thermal Protection System: Finite-Element Simulations

The thermomechanical behavior of damaged space shuttle tile thermal protection system (TPS) is simulated using the finite-element method. The effects of damage on the thermal protection capability and the induced thermal stresses in the TPS are evaluated by comparing the thermal and structural response of the damaged configurations with the undamaged configurations. The TPS consists of three components (the LI-900 tile, the strain isolation pad, and the underlying structure), and it is modeled as a discrete, three-layer structure. The TPS is subjected to the reentry heating and pressure profile corresponding to the Access to Space vehicle. The transient temperature distribution and the resultant thermal stresses are then computed. Three different sizes of damage simulating hypervelocity impact are considered. The validity of the simplifying assumptions used in an earlier study is systematically examined to determine the correct combination of assumptions required for modeling this class of problems. Influence of damage location on TPS response is also studied and it is concluded that centrally located damage represents a worst-case scenario.

Thermomechanical behaviour of a damaged thermal protection system: experimental correlation and influence of hypersonic flow

Abstract This paper describes a combined experimental and numerical study on damaged and undamaged space shuttle tile thermal protection system (TPS). The principal objective of the study is to determine its thermomechanical behaviour and assess the structural integrity of the TPS. The TPS tile specimens are subjected to a temperature profile corresponding to the thermal loads of the Access to Space reference vehicle. Experiments are conducted in a vacuum chamber that allows re-entry static pressure to be simulated. Temperatures on the top and bottom surfaces of the specimen, and the strains in the underlying structure are recorded. The experimental results are used to guide the development of a refined finite element model, which is subsequently used to simulate the interactions between the high speed external flow past the cavity that represents damage. Using this model, the relative effects of damage on the thermal protection capability and the induced thermal stresses are determined by comparing the response of the damaged configurations with the undamaged configuration. Damage increases the thermal loads and significantly reduces the radiation heat loss from the surface of the tile, resulting in elevated temperatures. Results indicate that damage can raise the maximum temperature in the tile to values that exceed its melting point.

Damage Detection through Analysis of Modes in a Partially Constrained Plate

The need for aircraft, both military and civilian, to serve longer and cost less to operate is ever present. The ability to potentially extend service life and reduce operating and maintenance costs are key factors in the many choices with aircraft programs. The field of structural health monitoring attempts to reduce labor and cost by allowing technicians to monitor selected properties of an aircraft’s structure to detect impending failure. This research examines methods to detect damage to a thermal protection system tile using representative aluminum plates. Plates are subjected to modal analysis in single and joined conditions in an attempt to provide the capability of sensing damage to a tile on the surface of a vehicle whereas the sensors remain on the substructure of the airframe. Jointly, the development of a means to model the system using finite-element techniques is explored. It is found that the finite-element modeling technique produces correlating modal frequencies within a 7.19% worst case ...

Minimum-mass design of sandwich aerobrakes for a lunar transfer vehicle

A structural mass optimization study of a sandwich aerobrake for a lunar transfer vehicle was conducted. The proposed spherical aerobrake had a base diameter of 15.2 m and radius of 13.6 m. A hot thermal protection system (TPS) and cold structure were used in the design. Honeycomb sandwich aerobrake structures made up of four different materials—aluminum alloy, titanium alloy, graphite-epoxy, and graphite-polyimide—were considered. Cases of aerodynamic load, equivalent uniform pressure, and aerodynamic plus thermal load were analyzed. Both linear stress and buckling analyses were conducted for a range of skin and core thicknesses. A graphical optimization procedure was used to determine the skin and core thicknesses of a minimum-mass aerobrake. The design criteria used were material strength, global buckling, and TPS tile deformation. Among them, the TPS deformation criterion was the most critical. The graphite-epoxy aerobrake was the lightest among the four materials studied. Its total mass is about 12.3% of the LTV mass, for supports at 75% span. Equivalent uniform loading produced smaller deformations, stresses, and buckling loads than did the more realistic aerodynamic loading for the same aerobrake configuration. Thermally induced stresses countered the aerodynamically induced stresses and hence had a beneficial effect on the deformation and buckling of the aerobrake.

Experimental aerodynamic heating to simulated Shuttle tiles

The heat transfer to simulated Shuttle thermal protection system tiles was investigated experimentally using a highly instrumented metallic thin wall tile arranged with other metal tiles in a staggered tile array. Cold-wall heating rate data for laminar and turbulent flow were obtained in the Langley 8-Foot High-Temperature Tunnel at a nominal Mach number of 7, a nominal total temperature of 3300 deg R, free-stream unit Reynolds number from 340,000 to 2,200,000 per foot, and free-stream dynamic pressure of 1.8 to 9.1 psia. Experimental data are presented to illustrate the effects of flow angularity and gap width on both local peak heating and overall heating loads.

Analytical and experimental evaluations of space shuttle TPS tile vibration response

Analytical studies and laboratory experiments have been performed to evaluate the vibration response of the Space Shuttle Thermal Protection System (TPS) tiles due to the intense rocket generated acoustic noise during lift-off. The TPS tiles are mounted over the exterior of the Space Shuttle orbiter structure through Strain Isolation Pads (SIP) which protect the tiles from thermal induced shear loads at their interface. The analytical predictions indicate that the response of a typical tile is governed by the structural vibration inputs through the SIP under the tile at frequencies below 250 Hz, and by the direct acoustic excitation over the exterior surface of the tile at frequencies above 250 Hz. An evaluation of the laboratory test data for this same tile, in which conditioned (partial) coherent output spectral analysis procedures were used, leads to exactly the same conclusion. The results demonstrate the power of conditioned spectral analysis procedures in identifying vibration response mechanisms when two or more of the inputs are highly correlated.