THERMAL CONDUCTIVITY CHARACTERIZATION OF A CFRP SINGLE-LAP JOINT

Fiber reinforced plastics (FRP), especially Carbon-FRPs, are a frequently used material for spacecraft’s primary and secondary structural design. Optimal results are achieved when the distinctive orthotropic mechanical properties are considered in the composite structures’ design process. Besides their excellent mechanical properties, FRPs offer also a high potential for thermal applications. In order to allow a partially coupled analysis, Lange [1] proposed a semi-analytic formula which connects the structural and thermal analysis of loadbearing single-lap joints (SLJ). For its validation, a thermal vacuum test was conducted [1] which showed non-conclusive results. The present paper presents shortcomings identified in [1] and how they are resolved. Next to improvements on the setup an additional experiment on material basis was conducted. It not only allowed the precise confirmation of the calculated CFRP material’s thermal conductivity lamda_11, but also to validate the whole setup for the SLJ experiment. The latest test results revealed that after the implemented setup changes and even though the temperature gradients are strictly limited, the experiment is very sensitive to radiation effects. This is shown by an analytical approximation of the radiative heat loss from the specimen to the environment and comparing it to the experimental results. [1] M. Lange, V. Baturkin, C. Hühne, O. Mierheim (2018). Validation of an analytical model describing the heat flux distribution in load-bearing CFRP single-lap joints_v1. In Proc. of 15th European Conference on Spacecraft Structures, Materials & Environmental Testing, Noordwijk, The Netherlands

ON ORBIT DEPLOYMENT OF THE EU:CROPIS SOLAR PANEL BY GFRP TAPE SPRING HINGES

Eu:CROPIS is a compact satellite featuring a biological payload. The Satellite was launched on December 3rd 2018. The cylindrical Satellite of 1m diameter has four deployable panels for power generation. Those panels are connected to the main structure by glass fibre reinforced polymer (GFRP) tape spring hinges. The hinges, comparable to curved metallic measuring tapes, have elastic energy stored when flattened and folded and thus deploy the panels by simply unfolding. When unfolded the hinges snap into their original shape and support the panels with considerable stiffness. No friction or mechanical locking is involved in the deploying process. The presented paper focuses on the practical handling of the hinges and the mechanisms during the final integration and the deployment process. The integration of the panels requires some special consideration. The hinges are not able to support the panels under gravity. The release mechanisms only work at a correct positioning of the panels. The measures taken to ensure the integrity and functionality of the hinges and mechanisms are described and examples are given for a correct and a false outcome. The separation is done by breaking a bolt with a heated bushing from shape memory alloy. Though reliable the separation cannot be timed down to the second and there is no direct feedback of the separation. To prevent an uneven opening of the panels several on orbit pre-tests are performed to ensure the functionality of the mechanisms for the actual deployment. At the actual separation the heating is monitored to ensure that all mechanisms are activated and the separation is working as proposed. Furthermore, a method was developed to detect the successful breaking of the bolts by use of the heating temperature data. The paper describes these checks and surveillance methods. As not all things go as planned some decisions were to be made before and at panel deployment. Also, the unfolding of the hinges was slower than during the on ground. Tests were made to simulate and understand the on-orbit behaviour. Lessons learned for further use of the mechanisms are presented

Mobile Asteroid Surface Scout (MASCOT) - Design, Development and Delivery of a Small Asteroid Lander Aboard Hayabusa2

MASCOT is a small asteroid lander launched on December 3rd, 2014, aboard the Japanese HAYABUSA2 asteroid sample-return mission towards the 980 m diameter C-type near-Earth asteroid (162173) 1999 JU3. MASCOT carries four full-scale asteroid science instruments and an uprighting and relocation device within a shoebox-sized 10 kg spacecraft; a complete lander comparable in mass and volume to a medium-sized science instrument on interplanetary missions. Asteroid surface science will be obtained by: MicrOmega, a hyperspectral near- to mid-infrared soil microscope provided by IAS; MASCAM, a wide-angle Si CMOS camera with multicolour LED illumination unit; MARA, a multichannel thermal infrared surface radiometer; the magnetometer, MASMAG, provided by the Technical University of Braunschweig. Further information on the conditions at or near the lander‘s surfaces is generated as a byproduct of attitude sensors and other system sensors. MASCOT uses a highly integrated, ultra-lightweight truss-frame structure made from a CFRP-foam sandwich. It has three internal mechanisms: a preload release mechanism, to release the structural preload applied for launch across the separation mechanism interface; a separation mechanism, to realize the ejection of MASCOT from the semi-recessed stowed position within HAYABUSA2; and the mobility mechanism, for uprighting and hopping. MASCOT uses semi-passive thermal control with Multi-Layer Insulation, two heatpipes and a radiator for heat rejection during operational phases, and heaters for thermal control of the battery and the main electronics during cruise. MASCOT is powered by a primary battery during its on-asteroid operational phase, but supplied by HAYABUSA2 during cruise for check-out and calibration operations as well as thermal control. All housekeeping and scientific data is transmitted to Earth via a relay link with the HAYABUSA2 main-spacecraft, also during cruise operations. The link uses redundant omnidirectional UHF-Band transceivers and patch antennae on the lander. The MASCOT On-Board Computer is a redundant system providing data storage, instrument interfacing, command and data handling, as well as autonomous surface operation functions. Knowledge of the lander’s attitude on the asteroid is key to the success of its uprighting and hopping function. The attitude is determined by a threefold set of sensors: optical distance sensors, photo electric cells and thermal sensors. A range of experimental sensors is also carried. MASCOT was build by the German Aerospace Center, DLR, with contributions from the French space agency, CNES. The system design, science instruments, and operational concept of MASCOT will be presented, with sidenotes on the development of the mission and its integration with HAYABUSA2

Towards a Reusable First Stage Demonstrator: CALLISTO - Technical Progresses & Challenges

In order to investigate the capabilities of a reusable launch system, JAXA, CNES and DLR have jointly initiated the project CALLISTO ("Cooperative Action Leading to Launcher Innovation for Stage Toss-back Operations"). The goal of this cooperation is to launch, recover and reuse a first stage demonstrator to increase the maturity of technologies necessary for future operational reusable launch vehicles (RLV) and to build up know-how on such vehicles under operational and developmental aspects. As the project has now turned into the detailed design phase, significant technical progresses have been made in definition, analysis and testing of systems and subsystems. The CALLISTO vehicle itself constitutes a subscale vertical take-off vertical landing (VTVL) stage with an overall length of 13.5 m and a take-off mass of less than 4 tons, which is propelled by a throttleable LOX/LH2 engine. It is capable to perform up to 10 consecutive flights during the planned flight campaign in French Guiana. Globally, the development effort on this system is equally shared between the three project partners. This paper presents the recent achievements in development of the key technologies for the reusable launch vehicle. While the design of critical subsystems has reached PDR level, detailed analyses and first breadboard tests have been performed successfully. These results are presented and discussed within the perimeter of the CALLISTO development roadmap. Possible technical challenges are indicated and their resolution methods are examined. Finally, the upcoming development steps are described which are foreseen to move forward to the qualification and maiden flight campaign

Simulation and measurement of thermal fluxes in load-bearing bonded FRP single-lap joints

Fibre reinforced plastics are a commonly used material for primary and secondary spacecraft structures. Besides their excellent mechanical properties, they offer a high potential for thermal applications. At the moment this combined potential of thermal and mechanical material properties is usually limited to the design and simulation of secondary composite’s structures. If both potentials would be combined on full spacecraft level, too, further mass and savings are possible. Hence, it is sought to connect the thermal and structural design in a combined semi-analytical mechanical and thermal 2D/3D FE simulation technique. As a prerequisite this paper focuses on the investigation of an exemplarily load-bearing bonded single-lap joint coupon of two CFRP laminates. The temperature distribution in the coupon is calculated analytically, numerically and subsequently measured in a corresponding thermal vacuum test. It is discussed what are the differences in the analytical and numerical solution and what is required before correlating the results

A trade-off study on the mechanical support structure of the MASCOT-2 small body lander package

The Mobile Asteroid surface SCOuT (MASCOT) is a 11 kg small body lander package developed under DLR lead for the Japanese space probe Hayabusa2. Based on MASCOT a modified MASCOT-2 lander package was studied for the AIM (Asteroid Impact Mission) mission study. The objective of the study was to develop a lander package, which maximizes heritage and reuse from the MASCOT predecessor. Thus the landing module’s framework structure retained the MASCOT design variant, being up-scaled by approx. 20% (MASCOT-2: 330mm x 300mm x 210mm, 13 kg). In contrast, the interface structure called Mechanical and Electrical Support Structure (MESS) experienced a more significant re-design, due to the need of interface simplification and the peculiarities of the lander deployment for AIM mission and the Didymos system. This paper introduces three possible design variants for the MESS, which are later narrowed down to two and presented in greater detail: a CFRP-honeycomb sandwich plate with additional unidirectional stiffening plies and an X-shaped solid CFRP hat beam structure. Both variants are much simplified compared to the MASCOT support structure, but retain the overall mounting variant. This conceptual discussion is followed by a detailed structural analysis of both mechanical support structures. Provided by a set of mechanical loads and stiffness requirements, the sandwich and beam interface structures are separately simulated in a finite element model, consisting of shell and beam elements, respectively. The attached landing module is modelled with both, shell and beam elements, allowing a coupled structural analysis of the system. By varying geometrical and material parameters in the structural MESS models, a trade-off between the resulting minimal masses, the stiffness and the strength requirements is performed. Specifically considering also development risks it is concluded that the sandwich design variant shows an overall better performance

COMPARISON OF TWO CARBON FIBER REINFORCED POLYMER GRID TYPES WITH RESPECT TO STABILITY, STIFFNESS AND MASS FOR THE USE AS INTERSTAGE STRUCTURE

In this paper, two grid types are discussed for the use as interstage 1-2 structure of the Veículo Lançador de Microssatélites (VLM). VLM is a launcher for microsatellites made of 3 stages arranged in tandem. The interstage 1- 2 is an open grid structure because of the hot separation of the first stage. The grid is made of light yet stiff carbon fiber reinforced polymer (CFRP). The two discussed grid types for the interstage are the axial grid and the hoop grid. The axial grid consists of axial stiffeners, which are parallel to the axis of the launcher, and helical stiffeners arranged at a +θ angle and at a – θ angle. The hoop grid is made of hoop stiffeners which are circular and helical stiffeners similar to those of the axial grid. The mechanical behavior of interstages made of these two grid types is discussed with respect to stiffness and stability. Finite element modelling (FEM) is used to determine the influence of the grid components on the eigenfrequency of the launcher. In this analysis, the stiffeners of the interstage are smeared into a cylindrical shell and implemented into a shell element model of VLM. The net thickness of the smeared stiffeners is varied to determine the influence of each stiffener type on the modal analysis. The smeared stiffeners are then discretized into individual stiffeners and the resulting gird barrels are analyzed via FEM using beam elements to determine the stability behavior of the grids. Here the influence of the width to thickness ratio of the stiffeners having rectangular cross sections is inspected. Having established an understanding of the mechanical behavior of the two grids, an interstage design for both grid types is developed using the above described analyses. It is found that the minimum mass of the axial grid interstage, which fulfils the stiffness and stability requirements, is lighter than the hoop grid interstage

Simulation von Bauweisen-Konzepten eines AKE Luftfrachtcontainers und deren Bodenplatte

Um neue Werkstoffe und Bauweisen für Luftfrachtcontainer einsetzen zu können, müssen diese qualifiziert werden. Hierfür stehen entweder Tests oder Simulationen zur Verfügung. Für Luftfrachtcontainer sind empfohlene Tests aus ULD Regulations [3] (ehemals dem IATA ULD Technical Manual [2]) bekannt. Da teilweise der komplette Luftfrachtcontainer als ein Bauteil getestet wird, sind nicht alle geforderten Tests umsetzbar. Eine FEM-Simulation ist hier die kostengünstigere und schnellere Alternative