Solar Array Wing Research Articles

Automated monitoring of flatness of solar array wings

This article discusses capabilities of a new special-purpose software-measuring system for monitoring flatness of solar array (SA) wings. The performance data and specifics of work are presented. The specifics of measurement of SA wings flatness are described. The allowable deviations from flatness of SA wings of the science power module (SPM) as part of the International Space station are given. The composition of the measuring complex is developed. The visual approach of making plans for testing methods and measurement techniques based on visual components is considered in detail. The elements from which the measurement plans are built are considered. Sample plans for verification procedures, methods of measuring flatness of the front and back wing surfaces are provided. Basic requirements for processing of measurement results which are stored for compatibility with the old manual approach of measurements are specified. Basic measurement processing algorithms for calculation of general and local wing out-of-flatness are provided. Key words: SA wing, solar array wing, meter, component model, overview, general out-of-flatness, local out-of-flatness.

Natural Vibrations of a Spacecraft’s Solar Array Wing

In the process of deploying the wings of solar arrays of a spacecraft, natural transverse vibrations arise in them. The study of these vibrations is an urgent problem, since in some cases they are capable of disrupting the commutation of photoelectric converters. The transverse vibrations of the wing originate from the connection of its sections when locking (coupling) hinge devices. In this case, the kinetic energy of the sections is converted into the potential energy of natural vibrations of the solar array wing. This problem is traditionally solved by using a finite element model with subsequent calculations in the ANSYS program. The purpose of this study is to develop an approximate analytical method for mathematical modeling of the natural elastic vibrations of elements of the solar array wing arising in the process of its deploying.

GaInP/GaAs/Ge 3중 접합 태양전지 배열기의 정지궤도에서 전력 성능 평가

The satellite on geostationary orbit accommodates multiple payloads into a single spacecraft platform and launched in June 26, 2010. The electrical power required to the satellite during sunlight is generated by a solar array wing. The solar cells are the GaInP/GaAs/Ge Triple Junction cells named Gaget2 cells from RWE Space, which were based on a Spectrolab epitaxy. This paper evaluates solar array power performance at end of design life based on the trend analysis results for the flight data on geostationary orbit. The estimated solar array power performance at end of design life compares with the power performance provided by solar array manufacturer. The solar cells show nominal behavior without significant degradation through the trend analysis results.

정지궤도위성 전력계 성능 평가

정지궤도위성은 다수의 탑재체를 하나의 위성체 플랫폼에 탑재하고 2010년 6월 26일에 발사되었다. 전력계는 태양 및 식 기간 상태에서 완전 조절 <TEX>$50V_{DC}$</TEX> 전력 버스를 제공한다. 위성에서 요구되는 전력은 태양전지 배열기 윙에서 생성되며, 에너지는 192.5Ah 용량의 리튬-이온 배터리에 저장된다. 본 논문은 전력계의 성능 평가를 향후 정지궤도위성 설계에 활용하기 위해 전력계의 중요 설계 변수들을 선정하고, 지상에서 시험 결과와 궤도상에서 운영 결과를 비교 분석하였다. 설계로부터 궤도상에서 운영 결과까지의 성능 평가를 통해 전력계는 중요한 성능감소 없이 정상적으로 동작되고 있음을 입증하였다. The satellite on geostationary orbit accommodates multiple payloads into a single spacecraft platform and launched in June 26, 2010. The Electrical Power Subsystem provides a fully regulated power bus at <TEX>$50V_{DC}$</TEX> in sunlight and eclipse conditions. The electrical power required to the satellite is generated by a solar array wing and the energy is stored by a Li-Ion battery with a capacity of 192.5Ah. This paper selects the main design parameters, compares and analyzes with the results at ground test and in orbit operation to apply this performance evaluation of the Electrical Power Subsystem to next satellite design on geostationary orbit. The Electrical Power Subsystem is demonstrated nominal behavior without significant degradation through the performance evaluation from design to in orbit operation.

통신해양기상위성의 열평형 시험 모델 및 예비 예측

COMS (Communication, Ocean and Meteorological Satellite) is a geostationary satellite and developed by KARl for communication, ocean and meteorological observations. It will be tested under vacuum and very low temperature conditions in order to verify thermal design of COMS. The test will be performed by using KARI large thermal vacuum chamber, which was developed by KARI, and the COMS will be the first flight satellite tested in this chamber. The purposes of thermal balance test are to correlate analytical model used for design evaluation and predicting temperatures, and to verify and adjust thermal control concept. KARI has plan to use heating plates to simulate space hot condition especially for radiator panels of satellite such as north and south panels. They will be controlled from 90 K to 273 K by circulating GN2 and LN2 alternatively according to the test phases, while the main shroud of the vacuum chamber will be under constant temperature, 90 K, during all thermal balance test. This paper presents thermal modelling including test chamber, heating plates and the satellite without solar array wing and Ka-band reflectors and discusses temperature prediction during thermal balance test.

Human and robotic repair of a solar array wing during ISS assembly mission 10A

With the installation of a new module and the relocation of three other modules, including multiple hand-offs from the station arm (SSRMS) to the shuttle arm (SRMS), International Space Station (ISS) assembly mission 10A/STS-120 was anticipated to be one of the most complicated ISS assembly missions ever attempted. The assembly operations became even more complex when a solar array wing (SAW) on the relocated Port-6 (P6) truss segment ripped while being extended. Repairing the torn SAW became the single most important objective for the remainder of STS-120, with future ISS assembly missions threatened by reduced power generation capacity if the SAW could not be repaired. Precise coordination between the space shuttle and ISS robotics teams led to an operational concept that combined the capabilities of the SRMS and SSRMS robotic systems in ways far beyond their original design capacities. Benefits of consistent standards for ISS robotic interfaces have been previously identified, but the advantages of having two such versatile and compatible robotic systems have never been quite so spectacular. This paper describes the role of robotics in the emergency SAW repair and highlights how versatility within space robotics systems can allow operations far beyond the intended design scenarios.

The stretched lens array (SLA) [spacecraft solar power

At IECEC 2001, this team presented a paper on the new stretched lens array (SLA), including its evolution from the successful SCARLET array on the NASA/JPL Deep Space 1 spacecraft. Since that conference, the SLA team has made significant advances in SLA technology, including component-level improvements, array-level optimization, space environment exposure testing, and prototype hardware fabrication and evaluation. This paper describes the evolved version of the SLA, highlighting recent improvements in the lens, solar cell, photovoltaic receiver, rigid panel structure, and complete solar array wing. In addition to excellent durability in the space environment, the near-term SLA will provide outstanding wing-level performance parameters: 180 W/kg specific power; 300 W/m <sup xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">2</sup> power density; 300 V operational voltage; 85% savings in cell area (cm <sup xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">2</sup> /W) and cell-related cost ($/W) compared to planar arrays; 9 kW/m <sup xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">3</sup> stowed power at launch.

The chemistry of micrometeoroid and space debris remnants captured on hubble space telescope solar cells

Prior to the retrieval in 1993 from low Earth orbit (LEO), the “—V2” Solar Array wing of the Hubble Space Telescope was exposed to hypervelocity impacts (micrometre to millimetre scale) from both micrometeoroids and space debris. The initial survey of the damage (100–3500μm diameter sized craters) identified that micrometeoroid remnants dominated the flux in the 100–1000μm size regime, with debris dominating >1000μm. These residues were composed of remnants of silicate minerals, calcite, metal sulfides and metals that often appeared as complex poly-mineralic melts within melt pits. A further survey of 10–100μm diameter craters identified that the most common chemistry was space debris with the crossover from meteoroids to debris being at around 30μm D CO. Residues include remnants of specialised steels and paint fragments but the dominant type is aluminium and aluminium oxide, which are almost certainly remnants of solid rocket motor operations. It is found that the relative contribution of debris as a function of size, agrees remarkably with a prediction derived using flux data from Long Duration Exposure Facility and a meteoroid model.

Lightweight inflatable solar array

A lightweight deployable solar array wing in the 200-1000-W range has been developed, on the Inflatable Torus Solar Array Technology Demonstration (ITSAT Demo) Project. The power density of a flight unit could be as high as 93 W/kg for a 200 W-class wing, including structure and deployment mechanisms. In Phase I, a proof-of-concept torus and array were constructed and deployed in the laboratory. During Phase II, a revised torus and array were constructed and tested at L'Garde and the Naval Research Laboratory. The qualification tests included random vibration, deployment in a thermal vacuum chamber, natural frequency determination, and thermal cycling. The flight design uses 2-milthick crystalline silicon cells on an atomic oxygen-protected flexible Kapton film substrate folded accordion style for stowage. The support structure is a rectangular frame comprised of two cylinders (which are inflated then rigidized), the array stowage box, and its cover. The cylinders, flattened, folded and stored for launch, are deployed by inflating with N2 and rigidized by straining the cylinder laminate material controllably beyond the elastic limit. The engineering protoflight array was designed for optimum power density but, because of availability, some of the components came from excess production runs. Because of this, the actual power density of the test article was 59 W/kg, or 36% lower than the baseline flight array. However, using components as designed, the projected 93 W/kg can be achieved. Because of a simple deployment mechanism, the cost of an ITSAT-type solar array is about one-half to two-thirds that of competing systems.