Development Of Thermal Protection Systems Research Articles

High-Temperature DIC Deformation Measurement under High-Intensity Blackbody Radiation

During the high-speed flight of a vehicle in the atmosphere, surface friction with the air generates aerodynamic heating. The aerodynamic heating phenomenon can create extremely high temperatures near the surface. These high temperatures impact material properties and the structure of the aircraft, so thermal deformation measurement is essential in aerospace engineering. This paper revisits high-temperature deformation measurement using the digital image correlation (DIC) technique under high-intensity blackbody radiation with a precise speckle pattern fabrication and a heat haze reduction method. The effects of the speckle pattern on the DIC measurement have been thoroughly studied at room temperature, but high-temperature measurement studies have not reported such effects so far. We found that the commonly used methods to reduce the heat haze effect could produce incorrect results. Hence, we propose a new method to mitigate heat haze effects. An infrared radiation heater was employed to make an experimental setup that could heat a specimen up to 950 °C. First, we mitigated image saturation using a short-wavelength bandpass filter with blue light illumination, a standard procedure for high-temperature DIC deformation measurement. Second, we studied how to determine the proper size of the speckle pattern in a high-temperature environment. Third, we devised a reduction method for the heat haze effect. As proof of the effectiveness of our developed experimental method, we successfully measured the deformation of stainless steel 304 specimens from 25 °C to 800 °C. The results confirmed that this method can be applied to the research and development of thermal protection systems in the aerospace field.

Design, Fabrication, and Testing of Metallic Thermal Protection Systems for Spaceplane Vehicles

This study presents the design and testing of a metallic thermal protection system (TPS) for future spaceplane vehicles. A conceptual design of the developed metallic TPS panel consisting of an outer sandwich structure, thermal insulation, inner support brackets, and a base framelike structure was redesigned from the metallic TPS panel developed by NASA. The thermal load was selected from a previous study, which was equivalent to the highest surface temperature of 1050 K. Nine metallic TPS prototype panels and supporting hardware structures made of 304 stainless steel were initially prepared for thermomechanical testing in room conditions. A noncontact thermal–mechanical measurement method was employed to measure full-field temperatures and deformations of the inner base structure and supporting hardware structures. The results indicated that the proposed design satisfied the requirements of temperature and permanent deformation limits. Panel-to-panel gaps showed a significant effect on the temperature rise in the inner base structure. The performance of the prototype proved the feasibility, reusability, and durability of using low-cost 304 stainless steel for TPS structures up to four thermal cycles, and may achieve more in further tests. Results provide valuable data for evaluating designs at the early stage of TPS development and testing in space conditions.

A Novel Zirconium Modified Arylacetylene Resin: Preparation, Thermal Properties and Ceramifiable Mechanism

With the rapid development of thermal protection systems for the aerospace industry and power electronics, polyarylacetylene (PAA) resin plays an important role because of its good mechanical properties, high glass transition temperature (Tg), low water absorption, high char yield (Yc), and the fact that there is no byproduct released in the curing process. In order to further improve the thermal property of PAA based FRP for the thermal protection field, the introduction of a zirconium element into arylacetylene is promising. In this paper, zirconium modified arylacetylene (ZAA) resin was prepared by two-step synthesis. The FTIR analysis characterized its molecular structure and confirmed the products. The viscosity of ZAA was about 6.5 Pa·s when the temperature was above 120 °C. The DSC analysis showed that the ZAA had a low curing temperature, and its apparent activation energy was 103.86 kJ/mol in the Kissinger method and 106.46 kJ/mol in the Ozawa method. The dielectric constant at 1 MHz of poly(zirconium modified arylacetylene) (PZAA) was 3.4. The TG analysis showed that the temperatures of a weight loss of 5% (Td5) and char yield (Yc) at 800 °C of PZAA were 407.5 °C and 61.4%, respectively. The XRD results showed the presence of SiO2 and ZrO2 in the PZAA residue after ablation. The XRF results showed that the contents of SiO2 and ZrO2 in PZAA residual after ablation were, respectively, 15.3% and 12.4%. The SEM showed that the surface of PZAA after ablation had been covered with a dense and rigid ceramic phase composed of ZrO2 and SiO2. Therefore, the introduction of Zr into arylacetylene greatly improved the densification of the surface after ablation, and improved the heat resistant property.

A review of the in situ probe designs from recent Ice Giant mission concept studies.

For the Ice Giants, atmospheric entry probes provide critical measurements not attainable via remote observations. Including the 2013-2022 NASA Planetary Decadal Survey, there have been at least five comprehensive atmospheric probe engineering design studies performed in recent years by NASA and ESA. International science definition teams have assessed the science requirements, and each recommended similar measurements and payloads to meet science goals with current instrument technology. The probe system concept has matured and converged on general design parameters that indicate the probe would include a 1-meter class aeroshell and have a mass around 350 to 400-kg. Probe battery sizes vary, depending on the duration of a post-release coast phase, and assumptions about heaters and instrument power needs. The various mission concepts demonstrate the need for advanced power and thermal protection system development. The many completed studies show an Ice Giant mission with an in situ probe is feasible and would be welcomed by the international science community.

Software Package for the Development of Thermal Protection Systems for Spacecraft Descent into the Atmospheres of Planets

The paper presents a mathematical model and a computational algorithm and describes a programming and computing suite (PCS) for the development of thermal protection systems for spacecraft designed to descend into the atmospheres of planets. The mathematical model and the PCS are implemented based on the development of thermal protection for the descent module of the EXOMARS 2020 surface platform.

Review of High Temperature Ceramics for Aerospace Applications

This paper presents a review of high temperature ceramics research for aerospace applications. Following a brief historical perspective, the paper reviews the effort to toughen ceramics for high temperature structural applications. These include: efforts to toughen zirconia-based ceramics, aluminum oxide, silicon carbide, silicon nitride, molybdenum disilicide and zirconium diborides and carbon-based composites. The development of thermal protection systems is also reviewed within the context of thermal barrier coatings (TBCs) and thermal protection systems for space vehicles. The paper concludes with a final section in which the implications of the results are then discussed for the thermostructural applications of ceramics in aerospace structures.

Adaptation of the in-cavity calibration method for high temperature heat flux sensors

The need for in situ heat flux measurements in hot structures, used in hypersonic vehicle thermal protection system development, combustion and propulsion research, and fire testing requires that heat flux sensors are characterized over their entire operating temperature range. The in-cavity heat flux sensor calibration technique has been adapted to accommodate elevated sensor temperatures, in an effort to develop a primary calibration scheme for high temperature heat flux sensors using an existing blackbody calibration system. The new scheme has been demonstrated through the calibration of a high temperature, thermopile-type heat flux sensor. The output temperature dependence of the high temperature heat flux sensor (HTHFS) has been successfully characterized over the range of 175–960 °C with acceptable uncertainty limits. The calibrated HTHFS sensitivity agrees well with a theoretical sensitivity model, suggesting that the extended in-cavity calibration technique is a viable choice for primary calibration of heat flux sensors at elevated sensor temperatures.

Comprehensive Flow Characterization in a 110-Kilowatt Inductively-Coupled-Plasma Heater

E LECTRICALLY heated wind tunnels are widely used as one of the ground-test facilities to evaluate the performance of thermal protection system (TPS) materials for atmospheric reentry vehicles. Recent interests in TPS development may be focused on accurate assessment of heat generation due to chemical reactions on the gas– surface interface, such as catalytic recombination on the surface, and oxidation and nitridation of the TPS material. To accurately assess the contributions of such processes to the net heat transfer rate, it is necessary to obtain detailed information about the flow properties such as temperature and concentration of the atomic species when TPS materials are tested. In the past studies [1,2], in an attempt to measure the thermochemical properties of the testflow in a 110-kWinductively-coupledplasma (ICP) heater at the Aerospace Research Center, Japan Aerospace Exploration Agency [3], emission spectroscopy associated with the line-by-line spectrum analysis was conducted by using the radiation code SPRADIAN2 [2]. The temperature and the chemical composition in the test flow were successfully determined under the representative operating conditions. More recently, the nitridation rate coefficient for the carbon surfacewasmeasured in the nitrogen test flow using the ICP heater [4,5]. In this experiment, to eliminate atomic oxygen remaining in the test section as an impurity, the ambient gas in the test section was replaced with pure nitrogen before ignition of the ICP heater. The experimental results indicated successful reduction of atomic oxygen by the gas replacement; however, its effectiveness has not been quantitatively assessed yet. In this study, following the preceding studies, comprehensiveflow characterization in a wide range of the operating conditions is conducted for future use in thermal protection materials testing and evaluation. A new imaging optical system is introduced to obtain emission spectra with less optical aberration in a wider wavelength range than before. The presence and source of impurities in the test flow are discussed more elaborately. Finally, effectiveness of the gas replacement on reduction of impurities is quantitatively assessed by obtaining the radial distribution of the impurities in the core flow of the test section.

Thermal protection system development, testing, and qualification for atmospheric probes and sample return missions: Examples for Saturn, Titan and Stardust-type sample return

The science community has continued to be interested in planetary entry probes, aerocapture, and sample return missions to improve our understanding of the Solar System. As in the case of the Galileo entry probe, such missions are critical to the understanding not only of the individual planets, but also to further knowledge regarding the formation of the Solar System. It is believed that Saturn probes to depths corresponding to 10 bars will be sufficient to provide the desired data on its atmospheric composition. An aerocapture mission would enable delivery of a satellite to provide insight into how gravitational forces cause dynamic changes in Saturn’s ring structure that are akin to the evolution of protoplanetary accretion disks. Heating rates for the “shallow” Saturn probes, Saturn aerocapture, and sample Earth return missions with higher re-entry speeds (13–15 km/s) from Mars, Venus, comets, and asteroids are in the range of 1–6 KW/cm 2. New, mid-density thermal protection system (TPS) materials for such probes can be mission enabling for mass efficiency and also for use on smaller vehicles enabled by advancements in scientific instrumentation. Past consideration of new Jovian multiprobe missions has been considered problematic without the Giant Planet arcjet facility that was used to qualify carbon phenolic for the Galileo probe. This paper describes emerging TPS technologies and the proposed use of an affordable, small 5 MW arcjet that can be used for TPS development, in test gases appropriate for future planetary probe and aerocapture applications. Emerging TPS technologies of interest include new versions of the Apollo Avcoat material and a densified variant of Phenolic Impregnated Carbon Ablator (PICA). Application of these and other TPS materials and the use of other facilities for development and qualification of TPS for Saturn, Titan, and Sample Return missions of the Stardust class with entry speeds from 6.0 to 28.6 km/s are discussed.

Evaluation of SCIROCCO plasma wind-tunnel capabilities for entry simulations in CO 2 atmospheres

A high-enthalpy wind-tunnel, named SCIROCCO, has been developed in order to perform tests related to the design and the development of thermal protection systems for atmospheric entry vehicles. The main interest of this numerical study is to investigate the capabilities of this facility to simulate flight conditions during entries into planet atmospheres containing CO 2. The objective is to estimate the performance map obtained on a sample in the facility test chamber. The flight values of these parameters need to be duplicated during ground tests of the thermal shield protecting the probe. First, a study at equilibrium of the mixture composition for a range of temperatures and pressures has been performed. Then, numerical simulations of the flow along the facility nozzle have been undertaken using a quasi one-dimensional method at equilibrium and a non-equilibrium Navier–Stokes solver. This has determined the theoretical performance envelope of the facility for a CO 2 atmosphere. The mass fraction of CO, a highly toxic explosive gas, in the test chamber has also been estimated.

Apollo thermal-protection system development

The development of the Apollo thermal-protection system is presented chronologically. The paper discusses the avenues explored, the problems encountered, and the rationale which resulted in a blunt vehicle protected by a low-density ablator in a honeycomb matrix bonded to a steel sandwich structure. Extensive ground and flight tests and analytical studies were performed to support this development. A flight test program was designed to provide data at extreme conditions within the flight corridor, and the results of these flights were used to verify system performance and to refine prediction capability. It is concluded that structural and manufacturing considerations often outweigh thermal-performance considerations in system design, that many important-appearing problems are not significant, and that the technology gained in the Apollo Program forms a reliable base for the development of future thermal-protection systems.