Hypersonic Aerothermochemistry Duplication in Ground Plasma Facilities: A Flight-to-Ground Approach

peer reviewedHigh conservative safety margins, applied to the design of spacecraft thermal protection systems for planetary entry, need to be reduced for higher efficiency of future space missions. Ground testing of such protection systems is of great importance during the design phase. This study covers a methodology for simulating the complex hypersonic entry aerothermochemistry in a plasma wind tunnel for a given spacecraft geometry without any assumption on axisymmetry or bluntness. A demonstration of this proposed methodology is made on the Qubesat for Aerothermodynamic Research and Measurements on AblatioN, QARMAN mission, which is a rectangular reentry CubeSat with a cork-based ablative thermal protection system in the front unit. The reacting boundary-layer profiles of the hypersonic entry probe compare well with the ones developing at the stagnation region of the plasma test model, defined with the proposed flight-to-ground duplication method

Qubesat for Aerothermodynamic Research and Measurement on AblatioN

International audienceThe preliminary design of the QARMAN re-entry CubeSat developed by the von Karman Institute is presented in this paper from de-orbiting to payload choices. It represents an ideal cost-efficient platform for re-entry flight test and validation of thermal protection system (TPS) materials with a demonstration flight scheduled for June 2015. The CubeSat comprises a standard double-unit platform with sensors for atmospheric research and a functional unit for essential satellite operations. A third unit accommodating an ablative heat shield is added to protect the vehicle against the extreme aerothermal conditions of the re-entry. The challenging aspect of the project lies on the constraining mass and form factor from the CubeSat standard, 3kg and 34x10x10 cm 3. Finally, the preliminary design of the vehicle results in a payload of 400 g collecting data all along the re-entry trajectory including the maximal heat flux conditions

DEVELOPMENT OF A DEPLOYABLE DECELERATOR CONCEPT FOR SMALL MARS LANDERS

Small exploration spacecraft and landers have proven to be scientifically useful and capable with missions such as Philae or Hayabusa 2. For the exploration of planetary bodies with atmospheres, novel and efficient entry, descent and landing (EDL) technologies are being explored. One of these concepts is the rigid deployable decelerator, which would be an alternative to existing EDL systems if proven to be feasible. For a Mars micro lander mission with an entry mass of 25 kg and a ballistic coefficient of 3:5 kg=m2, a concept for a deployable decelerator was developed. First, a flow-field analysis of different possible geometries of the deployed structure with Ansys Fluent was performed. From this, the pressure and temperature distribution and qualitatively the heat flux density along the profile wall of each geometry were determined. Subsequently, for a conical geometry, a design for a deployment mechanism was developed based on the umbrella concept, where a deployable structure spans a flexible thermally resistant cloth. The mechanism developed is a combination of folding parallelised struts, similar to an umbrella, and telescopic rods. Focusing on the strut structure and based on the results of the flow field analysis, with Ansys it was then investigated whether the design can in principle withstand the mechanical loads generated by the maximum dynamic pressure and how the temperature behind the deployed cloth is distributed under the maximum thermal load

Enabling and Enhancing Science Exploration Across the Solar System: Aerocapture Technology for SmallSat to Flagship Missions

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Aerothermodynamics of Pre-Flight and In-Flight Testing Methodologies for Atmospheric Entry Probes

Spacecraft, returning back to Earth, experience a very harsh environment during the encounter with the particles of the atmosphere. One of the major issues of the atmospheric entry is the extreme aerodynamic heating and the exothermic chemical reactions due to the gas-surface interaction at hypersonic free stream velocities. There is a constant effort by the space agencies to increase the understanding of the re-entry flight dynamics to optimize the spacecraft and especially its thermal protection system design. During the design process, ground tests and numerical tools are extensively used for their low cost and controlled environment abilities. However, real flight tests are indispensable for ground test and numerical tools validation. Due to high costs, such missions are rarely launched and thus there is an increasing interest in small affordable entry probes. Such platforms, once matured enough, may serve as an easily accessible tool to produce experimental data. It is the aim of this dissertation to propose tools to improve ground test capabilities and on the other hand to present the design, and using the developed tools, the testing of aerothermodynamic experimental payloads to collect flight data with a small entry probe. QARMAN (QubeSat for Aerothermodynamic Research and Measurements on AblatioN) is a triple unit CubeSat with ablative and ceramic thermal protection systems. It will perform an atmospheric entry with 7.7 km/s and a peak heat flux of 1.7 MW/m2. The aim of the in-flight experiments is to retrieve real flight data on ablator efficiency (temperature, pressure, recession) and temperature-pressure measurements for transition on the side panels. The peculiar squared geometry of QARMAN led to the development of a Flight-to-Ground Duplication methodology accounting for spacecraft geometries. It allows duplicating fully the stagnation region of a spacecraft with an arbitrary geometry in subsonic plasma wind tunnels. As a requirement of this methodology, free stream characterization techniques, specifically enthalpy measurement techniques are introduced. Experimental and numerical databases are built. A thorough ablation characterization campaign in VKI Plasmatron is conducted to provide input for building material response models. The cork P50 ablator is studied in terms of surface and sub-surface temperatures, emissivity, mass loss, char-pyrolysis layers, outgassing species and recession and swelling profiles. Similar in-flight experiments are proposed for QARMAN flight for in-depth temperature and pressure. Methods to build models for advanced data treatment are proposed. A full picture of post-flight analysis strategy is described for each study to relate the ground tools and flight data

Experimental investigation of passive/active oxidation behavior of SiC based ceramic thermal protection materials exposed to high enthalpy plasma

The passive/active oxidation behavior of a CVD-SiC coated C/C-SiC thermal protection material of the hypersonic SpaceLiner vehicle is investigated experimentally. For the safety of spacecraft and its passengers, it is important to know whether the thermal protection system will experience passive/active oxidation during the vehicle’s entry into Earth’s atmosphere. Active oxidation can promote material loss whereas passive oxidation forms a protective film. The high enthalpy flight conditions of SpaceLiner vehicle are duplicated in VKI Plasmatron, where the samples are exposed to high enthalpy plasma and the surface temperatures increase up to 2800 K at various total pressure (2–20 kPa) conditions. Surface temperature profiles, visual characteristics, mass changes, emissivity, spectrometer and SEM/EDX data are examined to identify the oxidation transition border of the tested material. A temperature jump is observed in all active oxidation regimes. The experimental results are found to be in good agreement with correlations from the literature.FAST20X

IRAS III Abschlussbericht

Dieser interne Bericht enthält die Aktivitäten des IRAS III-Projekts aller Partner, die vom Projektbeginn am 01.01.2020 bis zum Projektende am 20.06.2022 stattfanden