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Within the RETALT project, multiple Thermal Protection System (TPS) materials have been tested, using DLR ́s arc heated facility L2K in Cologne, Germany. The high enthalpy flow was used to simulate the aerothermal heating during reentry of the 1st stage of the RETALT 1 launcher. In total, 21 samples of five materials in stagnation point configuration have been tested, having two ablative cork materials and three high temperature ceramics. The cork materials by Amorim Cork Composites featured an existing material (P50) and a newly developed, trowelable cork based TPS. The ceramics were provided in cooperation with the LightCoce project. The samples were exposed to cold wall heat fluxes in the range of 200 to 800 kW/m2. In addition to tests with only one heat exposure, some samples were tested 3 times at the same cold wall heat flux and same total test time, named cycle tests. Those cycle tests are intended to yield more information on a possible reuse of the materials. Surface recession and mass change is measured.

In recent years, the space market has been pushing towards decreasing costs of launching spacecraft by reusing parts of the launchers. The purpose of this article is to present a feasibility study of a recovery system for the engine and engine frame of an existing, expendable heavy launch vehicle and present recommendations for further research. The concept developed is verified based on the Ariane 6. The recovery of the Vulcain Aft Bay (VuAB) is initialised by separation from the first stage at 157.7 km altitude while travelling at 6930 m/s. The study investigates an inflatable aeroshell for protection and deceleration during re-entry, after which it is proposed to further decelerate the VuAB using drogue parachutes. The final part of the concept entails retrieval of the VuAB by a helicopter in mid-air. To enable a controlled gliding flight during retrieval a parafoil is proposed. At launch, the recovery system will weigh 2789 kg with a payload penalty of 720 kg. The system can be integrated into the existing design of the launcher and will not interfere with nominal operations of the launcher. Implementing the recovery system can reduce the cost per launch of an Ariane 6 by 15%

Accepted: 2018-12-18, 資料番号: SA1180428000

Thermal Protection Materials (TPM) such as carbon/phenolic composites are used to protect spacecraft structures from extreme conditions. This protection is, in part, achieved by the decomposition via pyrolysis of the phenolic resin. Finite rate chemistry models are however still unable to predict the chemical production rates and composition of the pyrolysis products accurately. This is mostly due to the scarcity of experimental data for model validation. In this work, the decomposition of a phenolic material representative of thermal protection material is studied in a unique micro-pyrolysis unit for the temperature range 300-800 °C. This unit is equipped with highly sensitive detectors allowing us to identify and quantify products in a broad range of molecular weights up to 240 g mol−1. More than 50 different products of the pyrolysis of phenolic resin have been quantified with a mass balance closure greater than 80%. The major compound groups found are permanent gases, phenols as well as larger molecules such as diphenols and naphthalenes. In addition, the char yield obtained at the fast heating rates employed in our apparatus was found ∼ 5 % -points lower compared to traditional thermogravimetry.

平成30年度衝撃波シンポジウム (2019年3月5日-7日. 横浜国立大学経済・経営学部講義棟1号館 ）横浜市, 神奈川県, Symposium on Shock Waves in Japan, FY2018 (March 5-7, 2019. Yokohama National University), Yokohama, Kanagawa, 資料番号: SA6000134047, レポート番号: 1C4-2

Accepted: 2015-12-07, 資料番号: SA1150270000

第1回観測ロケットシンポジウム（2018年7月17日-18日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県, 1st Sounding Rocket Symposium (July 17-18, 2018. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan, 資料番号: SA6000127008, レポート番号: Ⅲ-2

Accepted: 2014-01-09, 資料番号: SA1150002000

A new class of lightweight carbon fiber reinforced plastic (CFRP)–-the lightweight ablator series for transfer vehicle systems (LATS)–-has recently been developed. The LATS is fabricated by heating and pressurizing a material in which resin is impregnated in the laminated carbon fiber felt. A characteristic required to ensure the excellence of a conventional lightweight CFRP ablator of the LATS is the simplicity of the resin impregnation process. Since dried bulk density can be easily controlled, this manufacturing method is beneficial for use in the aerospace industry. Here, the ablation characteristics of this material under high-enthalpy airflow are described by a recently developed computer code to simulate the one-dimensional transient thermal behavior. The validity of the mathematical model and the applicability of the ablation code are then discussed by comparing the simulated and experimental results of arc-heated tests with the LATS. A new index adopted in this study predicted the mass loss rate; the measured and estimated values of the total mass loss rate in various test conditions are in good agreement. Thus, a heated LATS material shows excellent performance characteristics for use in re-entry vehicles, and its surface and in-depth temperatures can be estimated using the developed analysis code.

Results from plasma wind tunnel testing of a lightweight carbon phenolic ablative material are presented. The tests were carried out in a high-enthalpy subsonic flow at a condition corresponding to the Hayabusa re-entry condition in 78 km altitude. The samples were produced from ZURAM which is a carbon-fiber structure impregnated with phenolic resin. The focus of the investigation was on the measurement method of the recession. Stereoscopic photogrammetry was applied to reconstruct the sample surface in three-dimensional space over the testing time. Two cameras were used to capture images simultaneously. Using a three-step process the sample surface for each image pair was reconstructed and the recession between the image pairs was calculated. The recession resolution is in the range of 45 µm. The advantage of the method is that the time-resolved in-situ surface recession measurement gives insight into the local recession distribution over the sample surface. This means that recession differences due to the sample geometry or due to material inhomogeneities can be identified. In order to verify this, samples were prepared that include distinct geometric features which influence the recession. Starting with a flat surface cylinder, samples with a rounded surface were prepared as well as a cone-sphere sample and one with a defined indentation in the form of a half-sphere. The temporally and locally resolved recession was successfully measured and related to the sample geometric features.

Results from an in situ analysis of ablative heat shield materials based on stereoscopic photogrammetry during plasma wind tunnel testing are presented. A carbon fiber preform material sample has been investigated in a high-enthalpy airflow corresponding to a Hayabusa reentry condition at a 78 km altitude. The applied photogrammetry uses two digital single lens reflex cameras triggered at 3 frames/s to acquire images simultaneously from two different perspectives of the sample exposed to the high-speed flow. The photogrammetric analysis resolves the surface with 25,000 px/cm 2 , approximately 400 dpi, allowing the in situ analysis of the recession phenomena of ablators. Material inhomogeneities are identified during the ablation process, and surface recession is derived with an accuracy of 21 μm . The method is a useful tool for the investigation of temporally resolved volumetric ablation phenomena in low-density carbon phenolic materials.

第59回宇宙科学技術連合講演会（2015年10月7日-9日. かごしま県民交流センター), 鹿児島市, 鹿児島県, An arc-heated wind tunnel test was conducted to confirm the measurement method of stagnation pressure on JAXA lightweight ablator in JAXA 750 kW Arc-heated Wind Tunnel. In this test we confirmed pressure measurement system installed on the small size re-entry capsule will work and this will not cause ablator breaks. And we found two problems for more accurate measurement: 1) the influences of ablator recession around pressure port and 2) ablator-gas on the measurement., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AC1600022000, レポート番号: JSASS-2015-4635

An integrated numerical method to obtain thermal response of ablator under aerodynamic heating environment is briefly described. In this method, a two-dimensional thermal response analysis code, namely SCMA2, is loosely coupled with a thermochemical nonequilibrium CFD code. The method is applied to reproduce the thermal response of ablator obtained by the heating tests conducted in the 1 MW arc-jet wind tunnel facility at ISAS/JAXA. Surface temperature, heat flux variation and surface shape change of the ablative test piece under arc-jet flow condition are calculated by the integrated code. These obtained results are compared with those given by one-dimensional SCMA code, and also with the experimental data for the purpose of validation. As to the comparison of calculated surface temperature with experimental data, quantitative agreement is yet to be accomplished because of the uncertainties in freestream conditions and material surface conditions. However, obtained results show that the present coupled approach is still available to predict the thickness of ablator required for space vehicles in future missions., 資料番号: AA0049212014, レポート番号: JAXA-SP-05-017

REV was the first reentry capsule developed through pure domestic technology and successfully recovered. REV is thermally protected against the severe aerodynamic heating by carbon-phenolic ablator material. The present paper describes the feature of the heating test in the arc wind tunnel and the ablator characteristics clarified during the research and development process of the REV heat shield. Because the heat flux anticipated on the stagnation surface is about 2 MW/sq m, the dominant surface recession mechanism is identified to be in the reaction control region. The thermochemical reaction data have been acquired through three arc-heater facilities with different enthalpy level for calibration and tuning of the ablation analysis code. The thermomechanical behavior of the ablator under the high heat flux environment such as delamination, or spallation also has been investigated for functional safety of the heat shield., 資料番号: AA0048710005, レポート番号: JAXA-RR-04-045

The re-entry capsule USERS/REV was successfully recovered in May 2003. REV has a heat shield system to protect inner equipments against the severe heating environment during the re-entry. The heat shield system is mainly made of three kinds of layers, i.e., CFRP ablator, felt type insulation material with thermal anchor and an aluminum shell structure. In the design of the heat shield system, contradictory requirements of heat ejection in orbit and heat protection during the re-entry had to be satisfied. Thermal protection performance of the heat shield system was calculated using an ablation analysis code. Thermal stress analysis of the ablator during the reentry was carried out using a large deformation FEM that considers materials non-linearity. For the development of the heat shield system, various kinds of tests such as the arc-heated tests of the ablator and the strain survey test of the heat shield system were carried out. After the recovery of REV, evaluation of the recovered heat shield system was carried out. The temperature time histories in the ablator obtained during the re-entry and the damage of the heat shield system measured after the recovery were compared with the results of ablation analysis and structural analysis, from which the heating environment during the re-entry and the cause of the damage were estimated. The evaluation of the heat shield system design based on the flight results is also described., 資料番号: AA0048710006, レポート番号: JAXA-RR-04-045

資料番号: AA0048089013

資料番号: AA0048089008

資料番号: AA0047120005

資料番号: AA0047120039

資料番号: AA0047120006

The effects of heat load, ablator density, and backup structure, etc. upon the heat shield performance of the lightweight phenolic carbon ablators named LATS were investigated using a one-dimensional ablation analysis code. The ablator density was assumed to be from about 260 to 1000 kg/m3. Heat flux time histories of a rectangular pattern were assumed, where cases of constant heating duration time and constant accumulated heat load (up to 600 MJ/m2) were considered. The heating level was assumed to be from 1 to 10 MW/m2, which means that the ablator surface is in the region of diffusion control oxidation/sublimation. The materials of the backup wall are assumed to be aluminum, stainless steel and high density CFRP. Main findings are: (1) For a low heat flux q with the same heating duration time tq, the necessary thickness, with which the maximum back surface temperature equals to the pre-determined allowable temperature, is nearly constant as the density ρv changes. On the other hand, the necessary thickness increases largely when q is larger and ρv is smaller. The ablator necessary mass increases with the increase of ρv and q for the same tq. (2) When a backup wall is attached, the necessary thickness decreases and the necessary mass including the wall mass increases. (3) For a constant accumulated heat load, necessary thickness and mass decrease for a higher heat flux q especially when ρv is high. (4) A lower density ablator with a CFRP backup wall gives the lightest mass of the heat shield system for most of the parameter range among the three wall materials. (5) For a high heat flux, selection of a lower density ablator gives a larger necessary thickness., 30th International Symposium on Space Technology and Science (ISTS), July 4th-July 10th, 2015, Kobe-Hyogo, Japan

資料番号: SA1160210000

宇宙航行の力学シンポジウム 平成16年度 宇宙航空研究開発機構 20041213-20041214 神奈川 日本, Symposium on Flight Mechanics and Astrodynamics 2004 Japan Aerospace Exploration Agency 20041213-20041214 Kanagawa Japan, 資料番号: AA0048089000

The effects of density, thickness and heat load upon the heat shield performance of the lightweight phenolic carbon ablators named LATS (Lightweight Ablator series for Transfer vehicle Systems) were examined quantitatively for both arc-heated test and re-entry heating conditions using a one-dimensional ablation analysis code. Thermal conductivity values of the ablator were tuned based on the arc-heated test results by matching the calculated temperatures to the measured data. Main findings are: (1) For both heating conditions, the heat shielding performance of the ablator has the same tendencies with respect to parameters of the ablator density, thickness and heat load. (2) The dependency of the back surface temperature upon the ablator density is small especially for a large ablator thickness. (3) The surface recession decreases with the increase of the density. However, the mass loss increases almost linearly with the increase of the density. (4) The ablator necessary thickness, with which the maximum back surface temperature equals to the pre-determined allowable temperature value, is nearly constant as the density changes. The ablator necessary mass increases almost linearly, with the increase of the density. (5) In considering the mechanism of nearly equal necessary thickness of the ablator, it is very important that the thermal diffusivity does not vary much with different densities of the LATS ablator. (6) From the point of the reduction of the ablator weight, the selection of a lower density ablator is more advantageous than that of a higher density ablator.

A space vehicle which undergoes the atmospheric re-entry or a planetary entry needs the heat shield system to protect inner equipments against severe aerodynamic heating environments. Charring ablator is usually used for the heat shield system. In order to design the heat shield system, it is necessary to predict the thermal behavior under aerodynamic heating by ablation analysis. A computer code for charring ablation and thermal response analysis is newly developed for simulation of one-dimensional transient thermal behavior of charring ablation materials. The mathematical model for the charring ablation including basic equation and computational method of ablation analysis is briefly described. A new ultra light weight phenolic carbon ablator called LATS (Lightweight Ablator series for Transfer vehicle) was recently developed. Arc-heated tests of the LATS ablator were carried out and measured results of the temperature response and surface mass loss are compared with the simulation results of the ablation analysis program. The agreement between the results of simulation and measurement is found to be good. It is also found that the mathematical model used in the ablation code can be applied to the ablation analysis of the low density LATS ablator., 28th International Symposium on Space Technology and Science (28th ISTS), 5-12 June 2011, Okinawa, Japan

The ablative material ZURAM has been developed by the German Aerospace Center DLR. It is manufactured from a rigid carbon felt impregnated with phenolic resin. It has a low density of approximately 0.36 – 0.4 g/cm³. Arc jet tests were carried out up to heat fluxes of 13.5 MW/m² indicating the potential of the material with recession rates of approximately 0.3 mm/s. To characterize the material in detail to facilitate the design for space vehicles, the complete set of thermal and mechanical material properties has to be established. This paper will present the determination of the mechanical properties related to the microstructure. A characterization in tension, compression and shear was carried out resulting in the respective moduli and strength values. It was found that the material has an anisotropic behavior. Since the material is rather delicate in comparison to other structural materials, stress and strain measurements were in part also based on optical measurements to not disturb the material response.