Thermal protection system integrating graded insulation materials and multilayer ceramic matrix composite cellular sandwich panels

2019 ◽  
Vol 209 ◽  
pp. 523-534 ◽  
Author(s):  
Xiuwu Wang ◽  
Kai Wei ◽  
Yong Tao ◽  
Xujing Yang ◽  
Hao Zhou ◽  
...  
Keyword(s):  
Matrix Composite ◽  
Thermal Protection ◽  
Ceramic Matrix Composite ◽  
Sandwich Panels ◽  
Ceramic Matrix ◽  
Thermal Protection System ◽  
Protection System ◽  
Insulation Materials ◽  
Multilayer Ceramic

scholarly journals Hot Structure Flight Data of a Faceted Atmospheric Reentry Thermal Protection System

2019 ◽  
Vol 2019 ◽  
pp. 1-16 ◽  
Author(s):  
Hannah Boehrk ◽  
Hendrik Weihs ◽  
Henning Elsäßer
Keyword(s):  
Matrix Composite ◽  
Heat Balance ◽  
Thermal Protection ◽  
Ceramic Matrix Composite ◽  
Ceramic Matrix ◽  
Thermal Protection System ◽  
Protection System ◽  
Maximum Temperature ◽  
Nose Radius ◽  
Maximum Heat

The second sharp-edged flight experiment is a faceted suborbital reentry body that enables low-cost in-flight reentry research. Its faceted thermal protection system consisting of only flat radiation-cooled thermal protection panels is cost-efficient since it saves dies, manpower, and storage. The ceramic sharp leading edge has a 1 mm nose radius in order to achieve good aerodynamic behaviour of the vehicle. The maximum temperature measured during flight was 867°C just before transmission ended and was predicted with an accuracy of the order of 10%. The acreage thermal protection system is set up by 3 mm fiber-reinforced ceramic panels isolated by a 27 mm alumina felt from the substructure. The panel gaps are sealed by a ceramic seal. Part of the thermal protection system is an additional transpiration-cooling experiment in which nitrogen is exhausted through a permeable ceramic matrix composite to form a coolant film on the panel. The efficiencies at the maximum heat flux are 58% on the porous sample and 42% and 30% downstream of the sample in the wake. The transient load at each panel location is derived from the trajectory by oblique shock equations and subsequent use of a heat balance for both cooled and uncooled structures. The comparison to the heat balance HEATS reveals heat sinks in the attachment system while the concurrence with the measurement is good with only 8% deviation for the acreage thermal protection system. Aerodynamic control surfaces, i.e., canards, have been designed and made from a hybrid titanium and ceramic matrix composite structure.

scholarly journals Investigating the Thermo-Mechanical Behavior of a Ceramic Matrix Composite Wing Leading Edge by Sub-Modeling Based Numerical Analyses

Computation ◽  
2020 ◽  
Vol 8 (2) ◽  
pp. 22 ◽  
Author(s):  
Michele Ferraiuolo ◽  
Concetta Palumbo ◽  
Andrea Sellitto ◽  
Aniello Riccio
Keyword(s):  
Mechanical Behavior ◽  
Matrix Composite ◽  
Thermal Protection ◽  
Ceramic Matrix Composite ◽  
Leading Edge ◽  
Ceramic Matrix ◽  
Thermal Protection Systems ◽  
Protection Systems ◽  
Very High ◽  
Composite Wing

The thermo-structural design of the wing leading edge of hypersonic vehicles is a very challenging task as high gradients in thermal field, and hence high thermal stresses, are expected. Indeed, when employing passive hot structures based thermal protection systems, very high temperatures (e.g., 1400 °C) are expected on the external surface of the wing leading edge, while the internal structural components are required to not exceed a few hundred degrees Celsius (e.g., 400 °C) at the interface with the internal cold structure. Hence, ceramic matrix composites (CMC) are usually adopted for the manufacturing of the external surface of the wing leading edge since they are characterized by good mechanical properties at very high temperatures (up to 1900 °C) together with an excellent thermal shock resistance. Furthermore, the orthotropic behavior of these materials together with the possibility to tailor their lamination sequence to minimize the heat transferred to internal components, make them very attractive for hot structure based thermal protection systems applications. However, the numerical predictions of the thermo-mechanical behavior of such materials, taking into account the influence of each ply (whose thickness generally ranges between 0.2 and 0.3 mm), can be very expensive from a computational point of view. To overcome this limitation, usually, sub-models are adopted, able to focus on specific and critical areas of the structure where very detailed thermo-mechanical analyses can be performed without significantly affecting the computational efficiency of the global model. In the present work, sub-modeling numerical approaches have been adopted for the analysis of the thermo-mechanical behavior of a ceramic matrix composite wing leading edge of a hypersonic vehicle. The main aim is to investigate the feasibility, in terms of computational efficiency and accuracy of results, in using sub-models for dimensioning complex ceramic matrix components. Hence, a comprehensive study on the size of sub-models and on the choice of their boundaries has been carried out in order to assess the advantages and the limitations in approximating the thermo-mechanical behavior of the investigated global ceramic matrix composite component.

Performance of cork and ceramic matrix composite joints for re-entry thermal protection structures

2017 ◽  
Vol 108 ◽  
pp. 270-278 ◽  
Author(s):  
K. Triantou ◽  
B. Perez ◽  
A. Marinou ◽  
S. Florez ◽  
K. Mergia ◽  
...  
Keyword(s):  
Matrix Composite ◽  
Thermal Protection ◽  
Ceramic Matrix Composite ◽  
Ceramic Matrix ◽  
Composite Joints ◽  
Protection Structures

scholarly journals Fatigue failure warning method for needled ceramic matrix composite by acoustic emission monitoring

2021 ◽  
Vol 2125 (1) ◽  
pp. 012039
Author(s):  
Geng Hou ◽  
De-Guang Shang ◽  
Lin-Xuan Zuo ◽  
Lin-Feng Qu ◽  
Ming Xia ◽  
...  
Keyword(s):  
Acoustic Emission ◽  
Fatigue Failure ◽  
Matrix Composite ◽  
Thermal Protection ◽  
Ceramic Matrix Composite ◽  
Ceramic Matrix ◽  
Threshold Value ◽  
High Temperature Mechanical Properties ◽  
Ae Energy ◽  
Aircraft Propulsion System

Abstract Ceramic matrix composite is a kind of mechanical engineering material with excellent high temperature mechanical properties, which has been widely used in aircraft propulsion system and thermal protection system. Therefore, it is of great significance to study the fatigue failure of needled ceramic matrix composite. In this investigation, based on the realtime acoustic emission (AE) monitoring of needled C/SiC ceramic matrix composite, the characteristics of AE energy during the fatigue damage process were obtained. In addition, considering the emission coefficient of AE energy and the threshold value of AE energy in single cycle, a method for judging the imminent fatigue failure of needled composite was proposed. By comparing the cycle of failure warning by proposed method with the experimental fatigue life, the proposed method can provide fatigue failure warning near and before fatigue failure.

An Advanced CMC Thermal Protection System Flight Experiment on the FOTON-M2 Mission

2007 ◽  
Author(s):  
T. Reimer
Keyword(s):  
Thermal Protection ◽  
Ceramic Matrix Composite ◽  
Thermal Protection System ◽  
Protection System ◽  
Structural Components ◽  
Excessive Heat ◽  
Flight Direction ◽  
Surface Diameter ◽  
German Aerospace ◽  
Structural Aspects

In June 2005 the re-entry module of the Russian FOTON-M2 spacecraft made a successful landing after a microgravity mission in space. On board was also the “KERAMIK” experiment by the Institute of Structures and Design of the German Aerospace Center (DLR) in Stuttgart, Germany. The re-entry technology experiment had performed a flight of a Thermal Protection System (TPS) with it’s structural components designed and manufactured fully in C/C-SiC, a Ceramic Matrix Composite (CMC) material developed by DLR. The experiment had a surface diameter of 340 mm and was located on the exterior of the re-entry spacecraft embedded into the ablative heat shield. The emphasis of the experiment was on the system aspects of the TPS design. It included two stiffened surface panels each fixed to three structural posts with a special fastener type, all components made from the C/C-SiC material. The experiment deliberately included a gap between the panels and a surrounding close-out ring to test the performance of a dedicated seal in that area. In addition to the structural aspects of the experiment, a set of different oxidation protection coatings was applied to the surface of one of the panels to conduct a comparative test under flight conditions. The position of the experiment was carefully selected with regard to the aerothermodynamic environment. Since the experiment technology aims at re-usable vehicles, a position in the stagnation area of the ballistic re-entry module would have resulted in excessive heat loads. Therefore a location was preferred at an angle of almost 90° relative to the flight direction. The temperature data that was measured during re-entry shows that the surface temperature was close to 1500°C, which was in the targeted range. The structural components on the surface of the experiment were in an excellent condition as a visual inspection immediately after landing revealed.

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