소형 달착륙선과 로버의 달 밤기간 생존을 위한 비용 효율적 열제어 시스템

학위논문 (박사) -- 서울대학교 대학원 : 공과대학 기계항공공학부, 2020. 8. 김규홍.A lunar day is equivalent to a month on the Earth. As a result, the daylight period of a lunar day is extremely hot due to two weeks of sunlight, and lunar nights are extremely cold due to the lack of sunlight for an equally long period. During lunar nights, lunar probes have no energy access from sunlight, and are required to withstand two weeks under extremely cold conditions. Various devices have been proposed as solutions to ensure lunar landers and rovers are able to withstand two-week-long lunar nights. Certain methods utilize radioisotope heating units (RHU) and radioisotope thermoelectric generators (RTG), the heat source device and electric generator of lunar probes, respectively. In addition, radiator lids, which are thermal switches, have been applied for this purpose. Recently suggested ideas include thermal wadis, which are thermal storage devices, and devices that store energy from under the lunar surface during the lunar daytime. As such thermal stability devices are accompanied with problems such as high costs, safety problems, or limitations in terms of weight or design, small lunar probes opt to forsake survival during the two-week-long lunar night and are instead designed for missions during the two-week daytime period.달의 하루는 지구의 한달에 해당하며, 태양이 비치는 달의 2주동안은 매우 뜨거운 상태의 달의 낮기간이고, 태양이 비치지 않는 달의 2주동안은 매우 차가운 달의 밤기간이다. 달 탐사선은 달의 밤기간동안 태양으로부터의 에너지 공급도 없고, 매우 추운 주변 환경 가운데서 2주간의 긴 기간을 버텨야 한다. 달 탐사를 위한 착륙선과 로버에는 2주동안의 달의 밤기간동안 생존하기 위해 그동안 여러가지 장치가 제안되고 시도되어 왔다. 달의 밤기간동안 달 탐사선의 열 공급장치인 RHU(Radioisotope Heating Unit), 혹은 전기 발생장치인 RTG(Radioisotope Thermoelectric Generator)등이 사용되어 왔고, 열 개폐장치인 Radiator Lid등이 사용되었다. 최근에는 열 저장 장치인 Thermal Wadi나 달의 지표면 아래에 달의 낮기간 동안 에너지를 저장하는 장치 등의 아이디어도 제시되었다. 이와 같은 온도 유지장치들은 가격적인 문제나 안전문제 혹은 무게나 구현의 어려움 등의 문제점을 동반하기 때문에, 소형 달탐사선의 경우에는 달 밤 기간 동안의 생존을 포기하고 달의 낮 기간인 2주동안만을 임무기간으로 설계되어 제작되기도 하였다. 본 논문에서는 소형 달착륙선과 로버도 2주동안의 달의 밤기간동안 생존할 수 있는 아이디어를 제시하였는데, 그것은 MLI (Multi-Layer Insulation) curtain system이다. MLI curtain system은 달 착륙선에 Lid와 RHU등의 장치를 적용하여 달의 밤 기간 동안 생존할 수 있도록 하고, 달 탐사 로버는 달 착륙선의 MLI curtain 안의 shelter로 이동하여 생존할 수 있도록 설계된 시스템이다. MLI curtain system은 소형 로버의 달 밤 기간 생존을 위하여 그 아이디어가 간단하고 적은 비용으로도 구현할 수 있는 장점을 갖는다. 본 논문에서는 MLI curtain system의 실현 가능성을 증명하기 위해 일반적인 우주비행체 열설계 열해석 과정을 따라 열설계와 열해석을 수행하였고, 열설계 요구조건에 부합하는 결과물을 얻었다. 뿐만 아니라 MLI curtain system feasibility를 향상시키기 위해 열제어 뿐만 아니라 구조, 전력, 소프트웨어 측면에서 고려할 점도 논하였으며, MLI curtain system의 안정성을 향상시키기 위한 아이디어도 제시하였다. 그 결과 MLI curtain system은 소형 로버의 긴 임무기간 동안의 생존을 위해 실현 가능하고 적합한 아이디어임을 증명할 수 있었다.CHAPTER 1. INTRODUCTION 1 1.1 History of Lunar Exploration Programs 1 1.2 Specifications of Unmanned Lunar Rovers 4 1.3 Thesis Objective and Proposal 9 CHAPTER 2. THERMAL DESIGN OF LUNAR EXPLORATION VEHICLES 11 2.1 General Spacecraft Thermal Design Process 11 2.2 Thermal Environment of the Moon. 16 2.3 Thermal Hardware of Lunar Exploration Vehicles . 20 2.4 Thermal Design of Lunar Landers and Rovers . 31 CHAPTER 3. METHODS OF LUNAR NIGHT SURVIVAL 33 3.1 Nuclear Energy 33 3.2 Lid System 34 3.3 Feasible Proposals. 36 3.4 A Proposal of MLI Curtain System. 39 3.5 Cost Effectiveness of MLI Curtain System. 42 CHAPTER 4. IMPLEMENTATION OF THE MLI CURTAIN SYSTEM. 44 4.1 Technical Challenges of the MLI Curtain System 44 4.2 The Shape of MLI Curtain 51 4.3 Suggestion of the Rover without Solar Cells 52 CHAPTER 5. THERMAL MODELING. 53 5.1 Governing Equation 53 5.2 Thermal Modeling Method of Lunar Surface . 58 5.3 Thermal Modeling of Lunar Regolith and Verification. 63 5.4 Thermal Modeling of the Lunar Lander and the Rover 69 CHAPTER 6. THERMAL ANALYSIS . 75 6.1 Thermal Design Requirements of the Lander and Rover 75 6.2 Thermal Analysis Case 76 6.3 Analysis Result of the Radiator Area and a Number of RHU. 81 6.4 Temperature Trends of the Lunar Lander and the Rover 83 6.5 Analysis of the Influence of the RHUs and Lid 94 6.6 Regolith Temperature Analysis according to the Effect of the MLI Curtain 96 6.7 Thermal Analysis considering the Fluff Layer being Blown Away 98 6.8 Summary of Thermal Analysis Results. 102 CHAPTER 7. CONCLUSIONS 103 REFERENCES. 105 초 록 112Docto

A Systematic Process for Assessing Human Spacecraft Designs in Terms of Relative Safety and Operational Characteristics

The research efforts in this dissertation are focused on reducing uncertainty in the conceptual design phase through a process of establishing a minimum functionality baseline before trading Safety and Operability in proposed spacecraft configurations. The challenge in human spacecraft development is how to combine the parts into a working design that complies with many requirements for top level mission objectives, safety, and mission success. The design methodologies presented here provides designers and decision makers with additional methods that provide an overall view of candidate design concepts. This work establishes a definition for a minimum functional design and is the first to group the fundamental mass parameters of a human spacecraft in the categories of Physics, Physiology, Safety, and Operability. The minimum functional baseline configuration described in this work is different from previous approaches because it eliminates the bias toward a minimum set of requirements. The amount of Safety in the spacecraft is the mass dedicated to safety through similar or dissimilar redundancy, safety components, margins, and dispersions. The amount of Operability in the spacecraft is the mass used to perform mission objectives and make functions easier or efficient. Because human spacecraft are highly coupled systems, the introduction of mass in one subsystem has downstream effects on other subsystems that are not easily recognized by designers and the use of rapidly reconfigurable prototypes allows designers and multidisciplinary teams to utilize Boundary Objects as a means of communication for maturing designs. The mass addition process coupled with the minimum functionality approach creates a tradespace of spacecraft configurations and provides designers with an overall view of how various levels of Safety or Operability will affect the overall spacecraft mass. The decisions made in the conceptual design phase are critical to the success of the program and uncertainty can lead to unnecessary redesign in later phases. The previous methods can be combined into a conceptual design process that couples easily with typical industry approaches to human spacecraft development. The use of minimum functionality as a precursor to more conventional approaches allows the spacecraft configuration to take shape before detailed CAD and higher fidelity analyses