Experience of passive thermal control of long-term near- Earth small satellite mission

The microsatellite BIRD (Bispectral InfraRed Detection) with mass of 94 kg and overall sizes 0.55 x 0.61 x 0.62 m operates on near-earth sun-synchronous orbit more than 11 years. The temperature range -10…+30 oC for payload and housekeeping equipment with average power of 35 W and peak power of 200 W in the observation mode (10…20 min) is provided by a passive thermal control system (TCS). The TCS supports a thermal stability of the payload structure by use of heat transfer elements – grooved heat pipes, thermally jointing the satellite segments. Two radiators, multilayer insulation (MLI) and low-conductive stand-offs provide the required temperature level. An analysis of TCS performance includes the definition of minimal, maximal and average temperatures of satellite units and their comparison with the designed parameters. The elaborated passive TCS successfully keeps the nominal temperature level of satellite components during one-year designed period of exploitation and sequent 10 years

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

A technical description of the Balloon Lidar Experiment (BOLIDE)

The Balloon Lidar Experiment (BOLIDE) was the first high-power lidar flown and operated successfully onboard a balloon platform. As part of the PMC Turbo payload, the instrument acquired high resolution backscatter profiles of Polar Mesospheric Clouds (PMCs) from an altitude of ∼38 km during its maiden ∼6 day flight from Esrange, Sweden, to Northern Canada in July 2018. We describe the BOLIDE instrument and its development and report on the predicted and actual in-flight performance. Although the instrument suffered from excessively high background noise, we were able to detect PMCs with a volume backscatter coefficient as low as 0.6 × 10^−10 m^−1 sr^−1 at a vertical resolution of 100 m and a time resolution of 30 s

PASSIVE THERMAL CONTROL SYSTEMS FOR SPACE INSTRUMENTS MAKING – SCIENTIFIC BACKGROUND, QUALIFICATION, EXPLOITATION IN SPACE

Passive thermal control systems (TCS) are one of obligatory system of any space mission, used as on large spacecraft and microsatellites Supporting of required temperature range for space instruments is supported by rational design of TCS with optimal choice of main thermal control components such as multilayer insulation, optical coatings, heat conductive elements, heat insulation supports, thermal conductive gaskets, radiating surfaces and other elements. New ideology in TCS design has come after appearance of new element – heat pipe(s) which is a super heat conductive thermal conductor with constant or variable thermal properties

Potential effects of optical solar sail degredation on trajectory design

The optical properties of the thin metalized polymer films that are projected for solar sails are assumed to be affected by the erosive effects of the space environment. Their degradation behavior in the real space environment, however, is to a considerable degree indefinite, because initial ground test results are controversial and relevant inspace tests have not been made so far. The standard optical solar sail models that are currently used for trajectory design do not take optical degradation into account, hence its potential effects on trajectory design have not been investigated so far. Nevertheless, optical degradation is important for high-fidelity solar sail mission design, because it decreases both the magnitude of the solar radiation pressure force acting on the sail and also the sail control authority. Therefore, we propose a simple parametric optical solar sail degradation model that describes the variation of the sail film's optical coefficients with time, depending on the sail film's environmental history, i.e., the radiation dose. The primary intention of our model is not to describe the exact behavior of specific film-coating combinations in the real space environment, but to provide a more general parametric framework for describing the general optical degradation behavior of solar sails. Using our model, the effects of different optical degradation behaviors on trajectory design are investigated for various exemplary missions

In-situ Magnesium Diboride Superconducting Thin Films grown by Pulsed Laser Deposition

Superconducting thin films of MgB2 were deposited by Pulsed Laser Deposition on magnesium oxide and sapphire substrates. Samples grown at 450C in an argon buffer pressure of about 10-2 mbar by using a magnesium enriched target resulted to be superconducting with a transition temperature of about 25 K. Film deposited from a MgB2 sintered pellet target in ultra high vacuum conditions showed poor metallic or weak semiconducting behavior and they became superconducting only after an ex-situ annealing in Mg vapor atmosphere. Up to now, no difference in the superconducting properties of the films obtained by these two procedures has been evidenced.Comment: 10 pages, 4 figure

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

Planetary Defense Ground Zero: MASCOT's View on the Rocks - an Update between First Images and Sample Return

At 01:57:20 UTC on October 3rd, 2018, after 3½ years of cruise aboard the JAXA spacecraft HAYABUSA2 and about 3 months in the vicinity of its target, the MASCOT lander was separated successfully by from an altitude of 41 m. After a free-fall of only ~5m51s MASCOT made first contact with C-type near-Earth and potentially hazardous asteroid (162173) Ryugu, by hitting a big boulder. MASCOT then bounced for ~11m3s, in the process already gathering valuable information on mechanical properties of the surface before it came to rest. It was able to perform science measurements at 3 different locations on the surface of Ryugu and took many images of its spectacular pitch-black landscape. MASCOT’s payload suite was designed to investigate the fine-scale structure, multispectral reflectance, thermal characteristics and magnetic properties of the surface. Somewhat unexpectedly, MASCOT encountered very rugged terrain littered with large surface boulders. Observing in-situ, it confirmed the absence of fine particles and dust as already implied by the remote sensing instruments aboard the HAYABUSA2 spacecraft. After some 17h of operations, MASCOT‘s mission ended with the last communication contact as it followed Ryugu’s rotation beyond the horizon as seen from HAYABUSA2. Soon after, its primary battery was depleted. We present a broad overview of the recent scientific results of the MASCOT mission from separation through descent, landing and in-situ investigations on Ryugu until the end of its operation and relate them to the needs of planetary defense interactions with asteroids. We also recall the agile, responsive and sometimes serendipitous creation of MASCOT, the two-year rush of building and delivering it to JAXA’s HAYABUSA2 spacecraft in time for launch, and the four years of in-flight operations and on-ground testing to make the most of the brief on-surface mission

Study of longitudinal heat transfer in low temperature heat pipes with axial grooves and discrete metal fibers for space thermal control systems

Development of space thermal control (TC) technique is tightly connected with achievements of heat pipe (HP) science and technology. The paper presents the experimental results of heat transfer for the following perspective HP designs with round cross section in wide range of operation parameters: saturation vapor temperature -70…+60 C, transferred heat power 0.5-100 W, axial heat flux density 10-250 W/cm2: a Copper – methanol HP with cooper discrete metal fiber capillary structure, external diameter 0.006 m, type – variable thermal conductance b Stainless steel – methanol HP with cooper discrete metal fiber capillary structure, external diameter 0.008 m, type- variable thermal conductance c Aluminum –ammonia HP with axial grooved capillary structure, external diameter 0.012 m, type- constant thermal conductance As characteristic parameter for comparison the ratio R=(T1-T2)/Q is used, where Q- transferred heat power, T1, T2 – averaged temperature of heat input and output zones of heat pipe. HPs “a” have shown essential impact of T2 of R, which varies from 60 K/W to 0.5 K/W. Physical basis of such behavior deals with changing of vapor flow regime in inner space and with the change of intensity of heat transfer by saturation vapor pressure. For HP “b” parameter R has variation from 90 K/W to 4 K/W. Physical basis of such behavior deals with creation of vapor –noncondensible gas boundary in coldest part of heat pipe and as the sequence, with variation of intensity of heat transfer. For HP “c” parameter R varies from 0.06 K/W to 0.045 K/W. Change of T2 has no essential impact of HP thermal resistance R. Comparison of longitudinal heat transfer characteristics of different types of heat pipes, obtained on representative quantity of tested units, allowed to define their functional abilities and duties in TC for different space applications

REGULATIVE CHARACTERISTICS OF THERMAL CONTROL SYSTEMS ON THE BASE OF VCHP AT VARIABLE HEAT GENERATION AND EXTERNAL HEAT EXCHANGE CONDITIONS

The paper is devoted to decision of the scientific-applied problem of creation of the high-efficiency thermal control systems (TCS) of scientific space apparatus, operating in non hermetical compartments, with the use of heat pipes (HP) of variable conductance (VCHP), executing the functions of transport of heat, control of the temperature of devices by maintenance of thermal balance in the system «mounting place of device in a spacecraft (SC) – device – space environment» on the required temperature level. Heat balance equations are the basis for definition of thermal parameters of TCS elements: heat pipe, radiator, thermal insulation, flexible elements, low conductance supports, cables, contact resistances, providing function of passive TCS. Proposed conceptions of TCS do not foreseen subsidiary electric power consumption of spacecraft and they are based on the use of own heat generation of scientific devices, or heat of subsidiary sources such as sun, providing stabilization of device temperature at the level of 290 K. Such approach is experimentally tested for the groups of scientific devices: for separate electronic cards (with mass of 0,3 kg), autonomous electronic unit (mass to 5 kg) and device panel of SC compartment (mass – 60 kg) at the change of own heat generation ration 1:10, temperatures of SC 253….323 K and illumination by external sources (sun, planet) to 270 W/m2