An Investigation into Thermal Design for Asteroid Landers

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2010:046

M A S T E R ' S T H E S I S

An Investigation into Thermal Design for Asteroid Landers

John Pegg Campbell

Luleå University of Technology Master Thesis, Continuation Courses

Space Science and Technology Department of Space Science, Kiruna

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CRANFIELD UNIVERSITY

CAMPBELL PEGG

AN INVESTIGATION INTO THERMAL DESIGN FOR ASTEROID LANDERS

SCHOOL OF ENGINEERING

Masters of Science in Astronautics and Space Engineering (‘SpaceMaster’)

MSc THESIS Academic year: 2009-2010

Supervisor: Dr. J. Kingston June 2010

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CRANFIELD UNIVERSITY

SCHOOL OF ENGINEERING

Masters of Science in Astronautics and Space Engineering (‘SpaceMaster’) MSc THESIS

Academic year: 2009-2010

CAMPBELL PEGG

An Investigation into Thermal Design for Asteroid Landers

Supervisor: Dr. J. Kingston

June 2010

This thesis is submitted in partial (45%) fulfilment of the requirements for the degree of Master of Science

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ABSTRACT

In-situ sampling of Near-Earth Asteroids (NEAs) has the potential to enable a con- siderably greater understanding of the formation of the Solar System. A significantly difficult aspect of a sample return mission to an asteroid is the variable thermal environment that the spacecraft has to endure. For a spacecraft to survive, appropriate radiator placement is necessary for internal heat rejection. Since the craft is also required to land on the asteroid, which radiates intense IR fluxes, the placement of the radiator on an asteroid lander is not straightforward.

An analysis was performed into potential areas on an asteroid lander that may be useful for the location of a radiator. It was found that an asteroid-facing radiator would be ineffective, whilst radiators located on the sides and top may be plausible under specific conditions. An analysis of how the radiator temperature would vary when a rendezvous with an asteroid occurs was also performed. It was found that if a spacecraft lands on an asteroid for less than ten seconds, then due to the landing process on the NEA’s surface there would be a minimum of 2-3 radiator temperature change from the original starting position and that the longer the landing time the higher this temperature change becomes.

New relationships were derived between the radiator temperatures of a craft situ- ated on the asteroid’s surface and the thermal characteristics of the asteroid. An equation for the average power required by a spacecraft located on a rotating asteroid was developed, enabling the thermal engineer to swiftly estimate basic asteroid lander characteristics. Each of the theoretical analyses performed were successfully validated using ESATAN-TMS. The guidelines and equations derived are of additional value to the thermal asteroid lander engineer as they can also be utilised for more detailed thermal design if parameters for specific missions are employed.

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ACKNOWLEDGEMENTS

I would like to give a special thanks to Dr. Jenny Kingston who was my supervisor throughout the duration of my thesis, who provided me with the means and knowledge to conduct this report. I would also like to thank the three industry experts, Steven Price and Simon Barraclough of Astrium UK and Philippe Poinas from the European Space Agency for providing me with advice and valuable feed- back throughout this thesis.

I would like to acknowledge the ‘Erasmus Mundus’ programme for providing a schol- arship for the Spacemaster programme and enabling me to conduct this course and thesis.

Many thanks to my parents who proofread a lot of my work and had a crash course in the thermal control system of an asteroid lander and my beautiful girl friend, Katherine Bennell, who dedicated a significant amount of time, effort and encour- agement for me to complete this dissertation.

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LIST OF FIGURES

1.1 Asteroids in orbit [Brown, 2010] . . . 3 1.2 The propellant to mass ratio for potential asteroid targets, where

Propel represents the propellant to mass fraction and DV is the Delta V to reach the asteroid, CP demonstrates if chemical propulsion is used and SEP is solar electric propulsion. [Agnolon, 2007]. . . 7 1.3 Temperature profile of an ideal asteroid where the sub-solar point is

the north pole. . . 8 1.4 Thermal inertia approximation for variously sized asteroids [Delbo

et al., 2007] . . . 9 1.5 The surface temperature of 1999 JU3 vs longitude versus longitutde

[Koschny, 2009] . . . 10 1.6 The subsoil temperature of the regolith of 1999 JU3 [Delbo and

Michel, 2008] . . . 11 1.7 The solar flux is dependent on the distance from the Sun [Gilmore,

2002]. . . 12 1.8 How the albedo flux varies with distance from the asteroid 1999 JU3 13 1.9 Variation of IR flux with distance from asteroid 1999 JU3 . . . 14 1.10 The IR radiation of an ideal asteroid . . . 14 1.11 Demonstration of absorption, transmission and reflection, look right,

sorry I can’t highlight references [Mecart, 2010] . . . 23 1.12 Typical composition of a MLI blanket [Gilmore, 2002] . . . 25 1.13 A diagram of how a heat pipe works [Gilmore, 2002]. . . 27

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LIST OF FIGURES LIST OF FIGURES

1.14 A demonstration of louvers on a spacecraft [Orbital] . . . 29

2.1 Standard mission trajectory [Agnolon, 2007] . . . 31

2.2 The landing phase profile . . . 34

2.3 The view factor of the side and bottom panels to the asteroid. . . 36

2.4 Standard Asteroid Lander Configuration . . . 41

2.5 Where the side, top and bottom panels are located on the Asteroid Lander in relation to the Sun’s and Asteroid’s Radiation. . . 47

2.6 The thermal inputs and outputs for a panel analysis. . . 48

2.7 The radiation flux experienced by a complete craft when landing on an asteroid for three hours with a surface temperature of 400 Kelvin. 51 2.8 The layout of the MATLAB code to calculate the radiator Tempera- ture Profile with the Maximum Radiator Temperature at 30. This outputs both the radiator area and temperature profile for a given starting altitude and landing altitude profile. . . 52

2.9 Incident radiation fluxes exposed to the craft during a asteroid stay mission . . . 52

2.10 ESATAN model of standard asteroid lander. . . 55

2.11 ESTAN model of the inside of the standard asteroid lander. . . 55

2.12 ESATAN model of standard asteroid lander with a plate representing the asteroid . . . 56

2.13 ESATAN radiative and thermal case that allows for a steady-state analysis of the spacecraft. The asteroid and craft were continuously facing the Sun to get complete solar flux. . . 57

2.14 ESATAN radiative and thermal case that allows for the calculation of the landing phase of a mission . . . 59

2.15 The incident fluxes on the spacecraft when conducting a the descent simulation in ESATAN. The black line represents the altitude profile of the lander. . . 60

2.16 ESATAN radiative and thermal case that allows for the calculation of an asteroid stay mission . . . 61

2.17 The asteroid’s surface temperature through the ESATAN simulation . 62 3.1 IR flux variation from the Surface of an asteroid versus the spacecraft’s altitude. . . 64

3.2 IR flux variation onto an spacecraft near the surface of an asteroid. . 65

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3.3 The sink temperatures of the panel facing the asteroid (bottom panel). 66 3.4 The minimum altitude at which the radiator would fail if placed on

the bottom panel. . . 67 3.5 The sink temperatures of the side panel of the standard asteroid lander. 68 3.6 The sink temperatures of the craft’s side panel versus the altitude of

the craft from the NEA’s surface. . . 69 3.7 The sink temperatures of the top panel versus the distance from the

Sun. . . 70 3.8 The ESATAN and theory comparison for the sink temperature of the

side and bottom radiator temperatures. . . 71 3.9 Cooling capacity of a bottom panel radiator if the spacecraft was

designed to reach thermal equilibrium at a certain altitude and temperature , where α = 0.25 and = 0.8. . . 72 3.10 Maximum cooling capacities if the spacecraft was design to reach

thermal equilibrium at a certain altitude and temperature for the side panel. . . 74 3.11 Maximum cooling capacities if the spacecraft was design to reach

thermal equilibrium at a certain altitude and temperature for the top panel. . . 75 3.12 The amount of internal power needed for a radiator to stay inside its

thermal limits. . . 75 3.13 Extra power required on spacecraft for the worst cold case thermal

scenario for a specific radiator size and internal component power from 0 to 300 Watts. . . 76 3.14 Extra power required on the standard spacecraft for the worst case

thermal scenario (cold case 1b) versus the radiator size at specific values of internal component power. Top panel. . . 77 3.15 The temperature profile of the side radiator for a 10 second landing

and internal component power output of 150 Watts over a variety of thermal capacitances . . . 79 3.16 The temperature profile of the side radiator for a 10 second landing

and internal component power output of 50 Watts over a variety of thermal capacitances. . . 80 3.17 The temperature change of a side radiator over the landing phase

with a 10 min landing versus the thermal capacitance of the Panel. . 81

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3.18 The temperature change of a side radiator over the landing phase with a 3 hour landing (day landing) versus the thermal capacitance of the panel. . . 81 3.19 Minimum thermal capacitance for a side panel for a specific landing

time and internal power. Note results below a thermal capacitance of 10,000 J/K were not generated because it is assumed to be too low. 83 3.20 Minimum side radiator area versus the thermal capacitance of the

panel for various landing times. Internal component Power is 100 Watts. . . 84 3.21 Minimum side radiator area versus the internal component power for

various landing times. Thermal capacitance is 100,000 J/K. . . 84 3.22 The radiator and electrical component temperatures during a landing

profile. Also a comparison of how these temperatures vary if the conductance between these two items is large or small. No internal power was added to the system . . . 86 3.23 The radiator and electrical component temperatures during a land-

ing profile. Also a comparison how these temperatures vary if the conductance between these two items is large or small. 100 Watts of internal power was added to the system . . . 87 3.24 Absorbed radiation through the side radiator. . . 94 3.25 Total absorbed power by side radiator. . . 94 3.26 The absorbed solar power input into the top panel. Radiator area is

1m². . . 97 3.27 The absorbed IR power input into the bottom panel. . . 98 3.28 A comparison between the theory and ESATAN results of the radiator

temperature during a asteroid stay mission. No internal power is added to the system. . . 100 3.29 A comparison between the theory and ESATAN results of the radiator

temperature during a asteroid stay mission. 100 Watts of internal power is added to the system. . . 101 3.30 Comparison between the radiator temperature and internal compo-

nent temperature for different conductance values. Internal power 100 Watts . . . 102 4.1 Comparison between the radiator temperature and internal compo-

nent temperature for different conductance values for both the math- ematical analysis and in ESATAN. The internal power is set at 100 Watts . . . 107

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C.1 The temperature profile of a side radiator for various thermal capacitances . . . 125 C.2 The temperature profile of a side radiator for various thermal capac-

itances . . . 126 C.3 The temperature profile of a side radiator for various thermal capac-

itances . . . 129 C.10 The temperature change between the equilibrium and maximum ra-

diator temperature for an asteroid rendezvous . . . 130 C.11 The temperature change between the equilibrium and maximum ra-

diator temperature for an asteroid rendezvous . . . 130 C.12 The temperature change between the equilibrium and maximum ra-

diator temperature for an asteroid rendezvous . . . 131 C.13 The minimum radiator area for a side radiator during a rendezvous

with an asteroid . . . 131 C.14 The minimum radiator area for a side radiator during a rendezvous

with an asteroid . . . 132 C.15 The minimum radiator area for a side radiator during a rendezvous

with an asteroid . . . 132

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LIST OF TABLES

1.1 Delta-V comparison of an asteroid mission to Mars and the Moon [Perozzi et al., 2001] . . . 6 1.2 Potential targets for a sample return mission to an asteroid [Agnolon,

2007], [Mueller et al., 2007], [Campins et al., 1995]. . . 6 1.3 The range of surface temperatures of 1999 JU3 for different thermal

inertias [Koschny, 2009]. . . 10 1.4 Past asteroid/planetary rendezvous missions . . . 15 1.5 Future and concept missions to NEO . . . 16 1.6 Typical thermal operating temperatures [Wertz and Larson, 2003] . . 17 1.7 Landing timeline, where T is the time of the landing on the surface

[Agnolon, 2007]. . . 22 2.1 A typical trajectory timeline where Y is the starting year [Agnolon,

2007] . . . 32 2.2 Worst case thermal environment when in deep space, where HC 1,

CC 1a and CC 1b are the hot case, eclipse cold case and deep space cold case respectively. . . 33 2.3 Worst Case thermal environment when orbiting the asteroid, where

HC 2 and CC 2 represent the hot case and cold case during the orbit of the craft around the NEA. . . 33 2.4 Worst case thermal environment when descending to the asteroid’s

surface. . . 35

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LIST OF TABLES LIST OF TABLES

2.5 View factors (assuming a cubic spacecraft) between the radiation from the Sun to panels of the spacecraft. . . 36 2.6 Worst case thermal environment for a hovering mission, where HC

Ho means ‘Hot Case Hovering’. . . 37 2.7 Worst case thermal environment for touch & go mission, HC TG

means ‘Hot Case Touch & Go’ . . . 37 2.8 Worst case thermal environment for day landing mission, HC DL

means ‘Hot Case Day Landing’. . . 38 2.9 Worst case thermal environment for asteroid stay mission, where HC

AS and CC AS mean Hot Case Asteroid Stay and Cold Case Asteroid Stay respectively. . . 38 2.10 Summary of thermal worst case scenarios. All fluxes have the units

of W/m² . . . 39 2.11 Standard asteroid lander characteristics. . . 40 2.12 Internal components of a standard asteroid lander . . . 44 2.13 The parameters that were varied and held constant during the landing

phase transient analysis . . . 51 2.14 The parameters that were varied During the landing phase transient

analysis . . . 53 2.15 The ESATAN model parameters for the standard asteroid lander. . . 54 2.16 Asteroid characteristics and orbital parameters for the ESATAN model

to calculate the sink temperatures . . . 58 2.17 Asteroid plate material characteristics and orbital parameters for the

ESATAN model to calculate the radiator temperatures for the landing profile . . . 59 2.18 Asteroid plate material characteristics and orbital parameters for the

ESATAN model to calculate radiator temperature for the asteroid stay mission. . . 61 3.1 Summary table of comparisons between the side and bottom radiators. 69 3.2 The mean incident fluxes into the radiator. Asteroid surface tem-

perature = 400 Kelvin, α/ = 0.31, AU = 0.96, Q = 0 Watts, side panel. . . 95 3.3 Side panel results for asteroid stay analysis. Asteroid maximum sur-

face temp = 400 K, Q = 0 Watts for Min X₁ and M inCp/A_rad calculations. . . 96

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3.4 The mean incident fluxes into the top radiator. Asteroid surface temperature = 400 Kelvin, α/ = 0.31, AU = 0.96, Q = 0 Watts, top

panel . . . 97

3.5 Top panel results for asteroid stay analysis. Asteroid maximum surface temp = 400 K, Q = 0 Watts for Min X₁ and M inCp/A_rad calculations. . . 97

3.6 The mean incident fluxes into the radiator. Asteroid surface temperature = 400 Kelvin, α/ = 0.31, AU = 0.96, Q = 0 Watts, bottom panel. . . 98

3.7 Bottom panel results for asteroid stay analysis. Asteroid maximum surface temp = 400 K, Q = 0 for Min X₁ and M inCp/A_rad calculations. 99 3.8 Comparison results for asteroid stay analysis. Asteroid maximum surface temp = 400 K, Q = 0 Watts for Min X₁ and M inCp/A_rad calculations. . . 99

4.1 Summary table of limitation between the side, bottom and top radiators for the sink temperature analysis. . . 104

4.2 Maximum cooling capacity of a radiator located either on the side, top or bottom of a spacecraft. It is assumed that the spacecraft is 1 AU from the Sun. . . 104

4.3 Power needed by a panel to keep the internal components inside its operational limits., where the emissivity is assumed to be 0.8. . . 104

4.4 Side radiator temperature change from thermal equilibrium at 2000 metres from the asteroid when conducting a rendezvous with the NEA. The assumed panel thermal capacitance was 150,000 J/K with a heat output (Q) of 100 Watts. . . 105

4.5 Minimum side radiator area for a craft that has to conduct a rendezvous with and asteroid. Assumed panel thermal capacitance of 150,000 J/K and heat output (Q) of 100 Watts. . . 105

A.1 1999 JU3 orbital properties . . . 115

A.2 Parameters for IR flux calculation . . . 116

A.3 Parameters for eclipse time calculations . . . 116

A.4 Cold Case Gravity Assist Summary . . . 117

A.5 Maximum and minimum distance the spacecraft will be from the sun 117 A.6 Cold Case 1b flux Summary . . . 117

A.7 Hot Case 1 flux Summary . . . 118

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A.8 IR flux calculation results . . . 118

A.9 Cold Case 2 flux Summary . . . 119

A.10 Hot Case 2 flux Summary . . . 119

A.11 Hover Case flux Summary . . . 119

A.12 Touch & Go flux Summary . . . 120

A.13 Hot Case AS flux Summary . . . 120

A.14 IR flux calculation results . . . 121

A.15 Cold Case AS flux Summary . . . 121

A.16 The eclipse time calculation. . . 122

B.1 Constraints for orbital parameter calculations. . . 123

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NOTATIONS

Short notation Description

AU Astronomical Unit

AOCS Attitude and Orbital Control system

BOL Beginning of Life

CC Cold Case

Cp Thermal Capacitance

EM Electromagnetic

EOL End of Life

equil equilibrium

HC Hot Case

IR Infra Red

IS Solar Incident flux

IA Albedo Incident flux

IP IR Incident flux

ISRU In situ Resource Utilization

JAXA Japan Aerospace Exploration Agency

LEO Low Earth Orbit

LLO Low Lunar Orbit

MTO Mars Transfer Orbit

NEO Near Earth Object

OBDH On Board Data Handling

OSR Optical Solar Reflector

Q Internal Heat

RHU Radioisotope Heating Unit

s/c Spacecraft

TT&C Telemetry, Tracking, and Commanding TWTA Traveling Wave Tube Amplifier

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CHAPTER 1 LITERATURE REVIEW

1.1 Introduction

Since humans started to look up at the stars they have been trying to explore and understand the solar system. In recent times this has led to the production of a large number of Earth-based telescopes for remote sensing. However, remote sensing incurs a multitude of limitations when conducted on Earth. Hence, humans have pushed forward to escape the bounds of the Earth’s gravity and to look more closely at these heavenly bodies, opening a new frontier. Space exploration needs to become cheaper and more efficient with the same scientific output, so humans have been looking at different celestial bodies that adhere to these requirements.

One type of these bodies which researchers are focusing on are asteroids. This is due to their proximity to Earth and the excellent scientific information that can be collected from their surfaces, enabling greater understanding of the history of the solar system and planetary formation. To achieve these scientific goals and enhance human knowledge, a better understanding of these bodies is needed.

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Thesis Aims and Objectives Literature Review

1.2 Thesis Aims and Objectives

The overall aim of this thesis is to investigate the design of the system level thermal control for an asteroid lander exposed to different asteroid environments. Specifically this pertains to:

Characterising the thermal constraints for an asteroid lander,

Providing a standardised theoretical model of an asteroid lander,

Proposing recommendations for thermal control systems for different asteroid environments and

Deriving guidelines and typical values for the thermal control system for the pre/phase A studies.

This thesis presents the meeting of these objectives, with the Literature Review and Methodology incorporating the first two aims and the Results and Discussion assessing the second two. The achievement of these aims is significant as such a standardised tool specific to the thermal design of asteroid landers has not been found in the literature.

1.3 Asteroids

An asteroid is a small body that orbits the Sun and is comprised of the building blocks of the Solar System [Peebles, 2000]. Most of the asteroids are found in the Asteroid Belt of the Solar System located between Mars and Jupiter and lying at a distance of 2-3.5 AU [Peebles, 2000]. There are two other main locations where asteroids are present: one being on either side of Jupiter’s orbit at the Sun-Jupiter L4 and L5 points which are called ‘Trojans’ [Chamberlin, 2010]; the others are rotating closer to Earth and more commonly known as Near Earth Asteroids or Objects (NEAs or NEOs).

NEAs are categorized into four classes [Yeomans, 2010]:

Apollo

Apohele

Aten and

Amor

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Asteroids Literature Review

An Apohele is a NEA whose orbit is completely inside the orbit of the Earth whilst an Amor has an orbit which is always outside the Earth’s orbit but inside that of Mars. An Aten is defined as an asteroid with a semi-major axis smaller than the Earth’s whilst an Apollo has a semi-major axis larger than Earth’s axis. However, due to the asteroid’s orbital eccentricity both types cross Earth’s orbit, leading to a potential collision risk [Yeomans, 2010]. A depiction of the asteroids in relation to the asteroid types are shown in Figure 1.1.

Figure 1.1: Asteroids in orbit [Brown, 2010]

Since there are a large variety of asteroids in our Solar System, numerous classification systems have been used to categorise them. The most common classification systems are the Tholen and SMASS Classifications [Binzel, 2010],[Tholen, 1989].

For the purpose of this document, the Tholen classification will be used because it is the most commonly employed. . Within this classification system, the majority of all bodies are placed into three categories:

C-group

S-group, and

X-group

where each of these groups can be broken down even further into different types.

The C-group asteroids are the most common type comprising approximately 75% of

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Motivation for Sample Return Asteroid

Mission Literature Review

known asteroids [Tholen, 1989]. These asteroids have a spectrum that is consistent with carbonaceous material and are believed to contain the primitive material of the Solar System, making them of great scientific interest [Koschny, 2009]. These asteroids generally have very low albedos (typically in the range of 0.03 to 0.1), thus making them more difficult to investigate from Earth. Another material found on C-group asteroids is water, hence it can give off a reddish tinge when observed in the optical wavelength [Tholen, 1989].

The S-group asteroids are the next largest group, comprising of 17% of the known asteroids [Tholen, 1989] and generally have a siliceous composition. In comparision to the C-group the albedo of these asteroids is higher (in the range of 0.1 to 0.22) [Tholen, 1989]. The last type of asteroid in the Tholen Classification system is the X-group which encompasses the remaining asteroids that do not fit into the C- and S-types. This is the smallest group and once again can be broken up into further subgroups. The most interesting type of X-group is the M-type where the majority are composed of nickel-iron in either a pure form or with small amounts of silicates [Koschny, 2009]. The albedo of these asteroids ranges between 0.1 and 0.2 [Tholen, 1989]. These types are also relatively flat and look reddish due to their metallic nature . Other types of asteroids are also present in this system but due to the low number of objects that are classified under these types, further knowledge of them is less relevant to this study.

1.4 Motivation for Sample Return Asteroid Mission

1.4.1 Scientific and Engineering Motivation

The most commonly accepted theory of planetary formation was proposed by a Russian named Viktor Safronov who hypothesised the Planetesimal Theory [Cessna, 2009]. This theory explains that a planet is formed from a rotating disk of gas called the protoplanetary disk with materials from the solar nebulae. This material slowly comes together under the force of gravity, initially in small parts, until it forms a planet, which is defined as a body in hydrostatic equilibrium [Mamajek, 2009].

Asteroids are believed to be remnants of this process that were not able to form a complete planet. As such the material in asteroids (and comets) is believed to hold the secrets to the original composition of the Solar System [Agnolon, 2007].

Planets also hold this secret (including Earth) but they have been altered over time by different chemical and heating processes, thus destroying the evidence. Asteroids and small bodies have also been a major factor in contributing water and organic material (essential to life on Earth) to planets, with 10-15% of the Solar System’s asteroid population colliding with terrestrial planets over the planetary formation

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period [Koschny, 2009]. Further understanding of the composition of these small bodies would allow humans to develop a knowledge of the formation of the Solar System as well as the origins of life on Earth. Moreover, NEAs have the potential of impacting Earth, causing catastrophic damage similar to how the current theories perceive that dinosaurs became extinct [Chapman, 2004]. A better understanding of these potential hazards may help to avoid such a fate.

Another benefit of increasing our knowledge of the composition of these small celestial bodies is the use of in situ resource utilization (ISRU). As described in Section 1.3 some asteroids are largely composed of a pure form of iron or nickel. These asteroids would have a market value of approximately one trillion US dollars on earth, hence creating a potential market for space industries. The material, if mined, could also be used for space applications so that the raw minerals would not have to be launched from Earth; instead future spacecraft could be manufactured or refueled and purely supplied from space resources. Therefore, understanding the composition of asteroids may have a financial advantage in the future [Kowal, 1988].

There are multiple ways of acquiring information on NEOs, including remote sensing from Earth or satellites or directly collecting a sample. The most popular method utilised to date has been remote sensing from Earth, especially with regard to assessing potential targets for possible asteroid missions. Unfortunately, there are limitations of using these techniques, including the resolution of the imagery, together with inaccuracies due to the NEA’s proximity. If an asteroid regolith sample is assessed directly these inaccuracies can be effectively minimised and would enable the analysis of trace elements in the sample [Agnolon, 2007]. To allow for an even higher precision and sensitivity of an asteroid composition the sample would need to be returned to Earth to allow for the performance of ground base tests [Koschny, 2009]. This is because high precision and sensitivity would be relatively difficult to perform in situ due to the mass, size, power and many other limitations of a spacecraft [Koschny, 2009].

Due to the reasons stated above, various space agencies have submitted mission proposals to collect a sample from an asteroid and bring it back to Earth [Koschny, 2009]. This includes JAXA that has already completed a mission to a NEA with the Hayabusa exploration vehicle [Kawaguchi et al., 1996]. Presently this vehicle is on its return trip and the success of this mission is ambiguous due to many system failures on course. Only after the potential sample has landed on Earth will it be determined whether it has achieved all its mission goals.

1.4.2 Cost Motivation

There are scientific benefits for analysing asteroid mineral composition but there are also numerous cost benefits for landing on asteroids as compared to landing on planets. The close proximity of these bodies to Earth along with their reduced grav-

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ity fields means that the delta-V budget required to land on NEA’s is significantly less than the equivalent budget for planets or for the Moon [Perozzi et al., 2001] . A summary of these budgets is detailed in Table 1.1.

Table 1.1: Delta-V comparison of an asteroid mission to Mars and the Moon [Perozzi et al., 2001]

Moon Mars Asteroid

Earths Surface to Earth Surface to LEO 9.5 Earth Surface to

LEO 9.5 LEO 9.5

LEO to LLO 4.1 LEO to C3 3.2 LEO to C3 3.2

LLO to Moon’s

surface 1.8 C3 to MTO 0.6 C3 to asteroid 1

MTO to Mars Capture

Orbit 0.9

MTO to LMO 1.4

LMO to Mar’s surface 4.1

LEO to moon 5.9 LEO to Mars 10.2 LEO to asteroid 4.2

The delta-V budget for a mission to an asteroid would vary dependent on the asteroid under investigation. There are a number of potential targets that would satisfy the scientific requirements of a sample return mission and suitable dates for the launch window. These NEA’s have already been summarised by various space agencies, as detailed in Table 1.2, along with the associated delta-Vs and mass fractions in Figure 1.2. Asteroid 1999 JU3 is the most likely asteroid to be visited, due to the requisite low mass fraction (propellant to total mass ratio) of the propulsion system needed to reach the object. As such, the orbital characteristics of this NEA will be used as the standard asteroid on which this analysis will be conducted on for a typical mission and spacecraft. A summary of likely targets for a sample return mission is shown in Table 1.2.

Table 1.2: Potential targets for a sample return mission to an asteroid [Agnolon, 2007], [Mueller et al., 2007], [Campins et al., 1995].

Parameters 1999JU3 1989UQ 2001 SK162 2001 SG286 Wilson- Harrington

a (AU) 1.19 0.92 1.93 1.36 2.64

e 0.19 0.26 0.47 0.35 2.64

inclination (deg) 5.88 1.29 1.68 7.75 2.79

Perigee (AU) 0.96 0.67 1.01 0.89 0.99

Apogee (AU) 1.42 1.16 2.84 1.83 4.28

Rotational Period (Hr) 7.67 7.73 68 tbd 6.1

Diameter (Km) 922 700 1520 350 1300

Albedo 0.063 0.06 Tbd Tbd 0.1

Max Surface Temp (K) 360 440 340 375

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Asteroid Thermal Environment Literature Review

Figure 1.2: The propellant to mass ratio for potential asteroid targets, where Propel represents the propellant to mass fraction and DV is the Delta V to reach the asteroid, CP demonstrates if chemical propulsion is used and SEP is solar electric propulsion. [Agnolon, 2007].

1.5 Asteroid Thermal Environment

Reaching an asteroid is typically much simpler and cheaper than reaching a planet (reference Table 1.1 and Figure 1.2) but once on the asteroid’s surface the key dif- ficulties start. One of the major issues with an asteroid is the thermal environment that the spacecraft has to survive during the landing and sample collection phase which is highly dependent on the NEO’s surface temperature. Therefore, this parameter needs to be understood and calculated.

1.5.1 Asteroid Surface Temperature

Theoretical models can be used to predict the surface temperature of an asteroid.

Even though theoretical models do have limitations on the accuracy of the results, they provide a good approximation to help with initial estimates. According to the theoretical model in ‘A Thermal Model for Near-Earth Asteroids’ [Harris, 1998], the temperature of the Asteroid on the side that is exposed to the Sun is defined by equation 1.1 .

T (ω) = T_{M ax}cos^1/4ω (1.1)

Where T is the temperature, ω is the angular distance to the sub-solar point (the

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point on the asteroid where the Sun is perceived to be directly over-head) and T_{M ax} is the maximum temperature of the asteroid located at the sub-solar point. TM ax is approximated by:

T_{M ax} = (1 − A) S ησ

1/4

(1.2) Where A is the asteroids albedo, S is the solar flux, is the emissivity of the asteroid, η is the correction factor (assumed close to one [Harris, 1998]) and σ is the Stefan- Boltzmann constant.

Figure 1.3: Temperature profile of an ideal asteroid where the sub-solar point is the north pole.

This calculation is a good approximation for the sunlit side of the asteroid [Harris, 1998], but to calculate the night side temperatures it is necessary to know the thermal inertia and rotational velocity of the NEA.

Thermal Inertia

The thermal inertia is a very similar concept to inertia for momentum, in which it parameterises the ability for the asteroid to lose and gain heat. The thermal inertia and the rotational velocity of the NEA are the two factors that are needed to calculate the surface temperature. The rotational velocity affects surface temperature by directly exposing half of the object to the solar radiation whilst the other half is in shadow. Therefore, the sunlit side will heat up due to the increase of solar flux while the night side will cool down due to the Infrared (IR) emission. If the asteroid rotates quickly then the solar energy is more evenly spread out across the

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surface. Conversely, if it rotates slowly then larger thermal extremes will appear on the surfaces [Barraclough and Trenkel, 2009]. How quickly this radiation is absorbed by the asteroid is dependent on the thermal inertia and is described as such:

Γ =√

k.r.c (1.3)

where k is the thermal conductivity, r is the density of the soil at the outer surface of the asteroid and c is the specific heat capacity of the surface material [Delbo et al., 2007]. For example, when a NEA has a dusty surface, it acts like an insulation layer over the asteroid, not allowing the heat to be absorbed into the NEA. This means that when this dusty surface is directly exposed to solar radiation it will heat up quickly because the heat cannot diffuse through the asteroid [Barraclough and Trenkel, 2009]. The opposite effect is seen when the surface is no longer exposed to the radiation. This dusty surface asteroid is known to have a low thermal inertia. For high thermal inertia objects the surface is rocky and dense allowing for the heat to be absorbed more quickly and thus have a larger lag time on the surface temperature [Barraclough and Trenkel, 2009]. Since the thermal inertia is relatively difficult to calculate because of some unknown quantities in the equation, approximations can be drawn from Figure 1.4.

Figure 1.4: Thermal inertia approximation for variously sized asteroids [Delbo et al., 2007]

Using Figure 1.4 and knowing the diameter of 1999 JU3 is 922 metres, a good approximation for the thermal inertia would be in the range of 100 to 500 J m⁻²s^−0.5K⁻¹.

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In another reference [Barraclough and Trenkel, 2009], the thermal inertia is assumed to be 116 J m⁻²s^−0.5K⁻¹, whilst [Delbo and Michel, 2008] has assumed the thermal inertia of 1999 JU3 is 300 J m⁻²s^−0.5K⁻¹, both of which are inside the estimated range. Knowing these two factors concerning thermal inertia, namely rotational ve- locities and thermal models, one can give an estimation of the surface temperature for the NEO, as shown in Figure 1.5.

Figure 1.5: The surface temperature of 1999 JU3 vs longitude versus longitutde [Koschny, 2009]

When observing this figure, one can see that the thermal inertia can greatly vary the surface temperature of the asteroid at different orbital positions. The range of surface temperatures versus the thermal inertia are shown in Table 1.3

Table 1.3: The range of surface temperatures of 1999 JU3 for different thermal inertias [Koschny, 2009].

Thermal inertia Perigee Surface Temp (K) Apogee Surface Temp (K)

(J/s^0.5/m²/K) Max Min Max Min

50 160 390 140 320

300 220 370

700 250 350 220 270

2500 280 300 240 260

Since the exact values would not be found until the craft reached the asteroid, the design of the lander would have to factor in these uncertainties.

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Asteroid Subsurface Temperature

Another important factor that needs to be considered when trying to simulate the thermal environment of the asteroid is the change in temperature in relation to the depth of the regolith. How this temperature varies is highly dependent on a multitude of factors including the distance from the Sun and the thermal inertia of the asteroid, but Figure 1.6 stands as a good approximation for 1999 JU3.

Figure 1.6: The subsoil temperature of the regolith of 1999 JU3 [Delbo and Michel, 2008]

It is evident that the asteroid has a constant temperature 0.15 metres into the regolith and this temperature is independent of which side is facing the Sun. This value will be used to approximate a good thermal model for the asteroid in this investigation.

1.5.2 Radiation Fluxes

The surface temperature of an asteroid is an important factor when considering the thermal environment that the spacecraft would experience, yet some of the most important influences on the temperature of the spacecraft are the radiation sources that are present. These sources include [Wertz and Larson, 2003]:

Solar Radiation

Albedo Flux

IR Radiation

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Solar Radiation

The solar radiation is the radiation directly from the Sun with an EM spectrum similar to that of a blackbody. The intensity of the radiation will be highly dependent on the distance the NEA is from the Sun, as is shown in Figure 1.7

Figure 1.7: The solar flux is dependent on the distance from the Sun [Gilmore, 2002].

The intensity of the solar flux as shown in Figure 1.7 is characterized by the equation [Gilmore, 2002]:

q_solar = 1367.5

AU² W/m² (1.4)

where AU called the Astronomical Unit and it is the distance from the sun to the earth. Since the orbital characteristics of the potential targets are known, the Solar flux at the asteroid can range between 3050 and 75W/m² depending on which asteroid is chosen. For 1993 JU3 the range is 1500 to 680W/m² [Koschny, 2009].

Albedo Flux

The albedo flux is the solar radiation reflected from the surface of the NEA at the same ratio as the albedo. This reflected radiation would have a very similar spectrum to the solar radiation. The albedo’s intensity is highly dependent on the solar flux and the surface texture of the NEA. To calculate the albedo flux one would use the following equation [Gilmore, 2002]:

q_α = Aq_solar (1.5)

where A is the NEA’s albedo and qα is the albedo flux. For the relevant asteroids that will be studied, the majority have a very low albedo, thus the bulk of the solar radiation will be absorbed by the asteroid.

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View Factor Since the albedo and IR flux are either reflected or radiated out of the asteroid, the distance and the relative angle of the two objects (spacecraft and asteroid) from each other becomes an important factor for radiation intensities incurred by the spacecraft. The view factor is the percentage of radiation that is emitted from one surface and falls upon another. There are many ways to calculate the view factor from the asteroid to the spacecraft but a good approximation for a flat plate facing the asteroid is to use the following equation [Brown, 2002].

F_a−b = R²_{N EA}

(H + R_{N EA})² (1.6)

where R_{N EA}is the radius of the asteroid (metres), H is the spacecraft’s distance from the asteroid’s surface (metres) and F_a−b is the view factor from a to b. This method assumes that the asteroid is an idealised perfect sphere (which is most likely not the case) and that the flux radiates out normally to the surface of the asteroid. These are not the most accurate assumptions but they will suffice for a good approximation that may be adapted to specific asteroids once a specific mission is defined.

Figure 1.8: How the albedo flux varies with distance from the asteroid 1999 JU3

IR Flux

When the asteroid absorbs the solar radiation, it has to be re-emitted as IR radiation (mainly comprised of Electromagnetic (EM) waves in the IR frequencies). This radiation is mainly dependent on the surface temperature and may be calculated by the radiation equation [Gilmore, 2002]:

q_IR= σAT⁴ (1.7)

where T is the surface temperature, is the emissivity, σ is the Stefan-Boltzmann Constant and A is the area that the fluxes are emitted from (this is normally assumed

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to be 1 for units of W/m²). It is evident that the surface temperature of the asteroid must be known to calculate the emitted IR flux, which is dependent on many unknown characteristics of the asteroid. Using the maximum and minimum temperatures that are most likely to be present on 1999 JU3 (150 - 400 Kelvin) [Delbo and Michel, 2008], the IR flux was calculated and the variation over the distance from the asteroid is shown in Figure 1.9. This is a valid approach, as knowing the extremes of temperature enables a knowledge of the radiation bounds the spacecraft must operate within.

Figure 1.9: Variation of IR flux with distance from asteroid 1999 JU3

Figure 1.10: The IR radiation of an ideal asteroid

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Spacecraft Properties and Constraints Literature Review

1.6 Spacecraft Properties and Constraints

The preceding sections have focused on the environment necessary for a lander to survive. In this section the discussion will mainly concentrate on the properties and constraints of the spacecraft. This will enable an understanding of the previous missions that have flown to NEAs, the normal thermal constraints of a spacecraft, the heat sources of a lander and the types of mission profiles that are experienced throughout such a mission.

1.6.1 Past, Present and Future NEA Lander Missions

In the recent history of space there has been only a few examples of flights to NEAs.

Such missions include Rosetta [Hechler, 1997], Hayabusa [Fujiwara et al., 2004] and NEAR [Dunham et al., 2002] (shown in Table 1.4) as well as a large range of concept missions which are shown in Table 1.5. These concept or future missions studies have focused heavily on craft that have landed or are aimed to interact with the surface of a NEA.

Table 1.4: Past asteroid/planetary rendezvous missions

Mission Agency Launch Yr Mission Description Reference NEAR NASA 1996 Fly-by Mathilde + rendezvous [Dunham et al., 2002]

to Eros

Hayabusa JAXA 2003 Sample Return mission to [Fujiwara et al., 2004]

Itokawa

Rosetta ESA 2004 Landing on Churyumov [Hechler, 1997]

- Gerasimenk (comet)

By observation of Table 1.4 and 1.5, it is clear that there have been a large number of studies on NEO landers, yet there have only been three missions that have either been completed or are presently in flight. In fact, NEAR was never designed to land on an asteroid; it was merely an added bonus as there was reserve fuel present on the spacecraft. Thus, there are opportunities available for future agencies/companies to gain new science and engineering knowledge by going to these NEOs.

1.6.2 Thermal Constraints

Different asteroid landers have different thermal constraints dependent on the equipment and payloads that the spacecraft is carrying. However, the majority of the components in any spacecraft are common and so there exists a typical range of temperature constraints. These typical thermal ranges are shown in Table 1.6.

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Table 1.5: Future and concept missions to NEO

Mission Agency Mission Description Reference

Marco Polo ESA Sample return mission to a NEA [Koschny, 2009]

Phobos-Grunt Roscosmos Sample return from Phobos [Marov et al., 2004]

Aladdin NASA Visit the Martian Moons Phobos and Deimos

[Agnolon, 2007]

Champollion - Deep Space 4

NASA Sample return from a comet [Kerridge et al., 1997]

Comet Nucleus Sample Return

NASA Comet sample return [Agnolon, 2007]

Deimos Sample Return

ESA Sample return mission from the Mars moon Deimos

[Agnolon, 2007]

Eve (European Venus Explorer)

ESA Rendezvous mission with Venus [Sivac and Schirmann, 2004]

Gulliver NASA - JPL Sample return from the Deimos [Agnolon, 2007]

Hayabusa II JAXA Hayabusa improved, NEA sample return

[Kawaguchi et al., 2008]

Hera Arkansas

Uni, USA

NEA multiple sample return [Agnolon, 2007]

Ishtar Astrium In-situ multiple asteroid mission [D’Arrigo]

Leanard ESA In-situ asteroid mission concept [Agnolon, 2007]

OSIRIS NASA NEA sample return [Agnolon, 2007]

Simnone UK, Italy,

ESA

Microsat rendezvous mission to NEAs

[Agnolon, 2007]

Astex DLR In-situ exploration to two NEAs [Agnolon, 2007]

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Table 1.6: Typical thermal operating temperatures [Wertz and Larson, 2003]

Components Min () Max ()

Overall Internal Components 0 50 Payload

Optical Sensors 15 25

Infrared Module -40 30

Radiometric Units -10 50

On board Computer -10 50

Electrical Power

Batteries 0 25

Solar Arrays -105 110

Power control Unit -20 55

Attitude Control

Sun and Earth Sensors -30 50

Magnetometer -80 80

Electronic Units -10 55

Gyro package 0 50

Propulsion

Tank and lines 7 55

Thrusters 7 65

Harness

Internal -15 55

External -100 100

Thermal Control

Multi-Layer Insulation (MLI) -160 250

Heaters, Heat pipes -35 60

Radiators -95 60

Structures

Nonalignment Critical -45 65

Alignment Critical 18 22

Mechanisms

Electric Motors -45 80

Continued on next page

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Table 1.6 – continued from previous page

Components Min () Max ()

Antennas

Parabolic Reflector -160 95

TT&C -65 95

It is evident that the temperature ranges can vary depending on the type of equipment present inside the spacecraft but the general limiting factors that affect the internal temperature are the electrical equipment and propellant that force the internal temperature to be between approximately 0 and 50.

1.6.3 Heat Generation

There are many sources of heat that can be used by the spacecraft to raise its internal temperature. The first category is the external sources that naturally exist (such as solar radiation, IR radiation and albedo flux), which were discussed in the previous Sections. The next major category is produced from the electrical and mechanical components of the spacecraft. For an average spacecraft the major heat sources include:

Batteries

Payload

Wave tube amplifiers

Processors and

Mechanisms

These sources provide a heat input into the system that results in an overall heat increase of the spacecraft. The amount of heat these systems put into the craft depends upon the individual components and when they are used.

Electrical Components

All electrical components in a spacecraft will produce a certain amount of heat when in use, as none are perfectly efficient. Most of this heat production is caused

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by the internal resistance of the components. This can vary greatly depending on the component, the power needed for it to run and how much of this power is converted into heat. It is assumed for this report that 100% of the power needed for these components to run will result in thermal dissipation thus simplifying the calculations. [Gilmore, 2002]

Battery

The battery is one of the most critical components of the spacecraft, as there is a limited temperature range any battery can operate within and also because of the varying heat output it produces. The heat generated from a battery will vary greatly with the type of battery used, if it is discharging or charging and the amount of power it holds. The heat from this source comes mainly from the internal resistance of the battery when charging and discharging. It is assumed that when the charge/discharge efficiency for a battery is given then the losses that are converted into heat can be estimated [Gilmore, 2002].

Mechanisms

The next source of heat in a spacecraft is from mechanical components and their associated friction. Once again the number and types of mechanical components present will dictate the amount of thermal energy produced. Fortunately on an asteroid lander there are only a few moving components present, which include;

Attitude and Orbit Control Subsystem (AOCS) package,

Solar array deployment and orientation mechanisms,

Antenna deployment and orientation mechanisms,

Sampling mechanisms and

Landing mechanisms.

In this paper it is assumed that most of these frictional heat sources are zero, mainly due to the sporadic and one off uses of these devices, except for the gyros which are assumed to be frictionless.

Propulsion

The propulsion subsystem has the potential to produce a lot of thermal input depending on which propulsion system is utilised. It is assumed that the main unit can be either electrical or chemical, as both of these devices have been proven to work

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for NEA missions. For the AOCS propulsion units, typically a chemical system is employed using the mono-propellant of hydrazine [Koschny, 2009]. For the complete propulsion system, the normal procedures on how to remove the heat production from these units is by radiation out into space, thus ensuring none of the heat generation goes into the spacecraft systems. If the propulsion system has recently been fired then there is higher chance that a percentage of the energy radiated out will be exposed to the system but for this analysis it is assumed that none of this heat influences the energy balance for the spacecraft.

Other Heat Sources

There is an additional heat source that is able to affect the temperature of a spacecraft, namely a re-radiator flux from one surface to another. A good example is the solar arrays, which when exposed to a solar flux, heat up. Once again this solar array will emit IR radiation to cool down and some of this radiation will be absorbed by another part of the craft. This is a phenomenon that needs to be carefully monitored in any thermal analysis.

1.6.4 Spacecraft Mission

Section 1.5.2 discussed the environment in which the spacecraft must operate. One of the key ways to control the thermal environment on the spacecraft is by control- ling the spacecraft’s orientation and mission profile. During a multitude of studies conducted on asteroid missions, some investigation has been made into some key scenarios concerning sampling strategies. Four techniques are: [Agnolon, 2007]

Hovering,

Touch and go,

Short term stay and

Long term stay.

Hovering

A Hovering technique is where the spacecraft moves closer to the asteroid and will hover above it while collecting a sample. The length of time and the height of the hover will be dependent on the type of sampling equipment used. Sampling techniques can include: firing projectiles for the collection of extracted dust, a surface penetrator which is fired and attached by a tether to the lander or even using a long mechanical collecting arm. The altitude from the surface of the asteroid can

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vary from 1 to 10 metres, whilst being within the time frame of approximately ten minutes [Agnolon, 2007], depending on which sampling technique is used. The main advantage of the hovering technique is the weight reduction because landing legs are not needed. Disadvantages include a higher degree of complexity and a need for a very good control system to ensure a stable hover.

Touch & Go

The Touch & Go technique involves the spacecraft landing on the surface and then taking off again within a small time frame (approximately on the order of seconds) [Agnolon, 2007]. This is then a very quick transfer of the lander on the surface with the bulk part of the spacecraft being less than a meter away from the asteroid (dependent on the size of the lander legs). Due to the short dwell time, the lander would not reach thermal equilibrium while on the surface, thus there would be a reduction in the requirement for the spacecraft to withstand the extreme thermal environment of the NEA. This method also allows the spacecraft to ignore the need for an anchoring device which would be required if the lander was to remain on the surface of the asteroid for a longer period. The Touch & Go technique, then, reduces the mass of the lander and the complexity of the mission. If this sampling strategy was conducted it would have the potential to be simply repeated in different sampling areas, allowing the spacecraft to pick up a wide variety of samples.

Short and Long stay

The Short and Long Stay mission (also known as Day Landing and Asteroid Stay missions, respectively) differ by the length of time the spacecraft remains on the asteroid’s surface. For a Short Stay mission this would entitle the lander to stay on the NEA for no longer than approximately half of an asteroid’s rotational period so the spacecraft is only subjected to a narrow band of surface temperatures [Agnolon, 2007]. As the lander would reach thermal equilibrium, significant thermal design would be necessary compared to the ’Touch & Go’ mission. For the Long Stay mission, the lander would experience the full range of surface temperatures during the day and night of the asteroid. Significant design challenges would be incurred by the thermal system for this mission due to the large temperature variations on the asteroid. Both of these mission types would require an anchoring system which might include an actual anchor or merely a reverse thrust to keep the lander on the surface. Both of these systems are plausible for the Short missions, but for Long Stay missions, an anchor would be preferable due to the amount of propellant that would be used during a reverse thrust [Katzkowski, 2009].

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Thermal Control Techniques Literature Review

Rendezvous Timeline

For the different sampling strategies that were mentioned above, the typical timeline for the rendezvous of the craft with the asteroid is shown in Table 1.7. This gives a good approximation on how long the craft would be exposed to the asteroid environment and compares the different mission profiles.

Table 1.7: Landing timeline, where T is the time of the landing on the surface [Agnolon, 2007].

Altitude (m) Hover Touch Day Landing Asteroid

& Go Landing

Parking Orbit 2000 T - 4h T - 4h T - 4h T - 4h

Lost Comms 500 T - 1h T - 1h T - 1h T - 1h

Landing 0-10 T T T T

Takeoff 0-10 T + 20mins T + 10s T + 2h T + 7h

Regain Comms 500 T + 1:20h T + 1h T + 3h T + 8h

Parking Orbit 2000 T + 12h T + 12h T + 12h T + 21h

1.7 Thermal Control Techniques

In the previous section, the majority of the constraints of a spacecraft were addressed in relation to the thermal subsystem. In this section the different techniques used to keep the components between the thermal constraints for typical NEO landers will be discussed. Key techniques include:

Surfaces and finishes,

Insulation,

Radiators,

Louvre,

Heat pipes,

Heaters and

Radioactive Heating Units (RHU).

Most of these techniques are passive control methods (methods that do not use any control techniques), with the exception of louvres, heaters and RHU, which are some of the more common or useful devices to implement. The use of active methods causes the mass of the lander to increase significantly, therefore it is preferable to maximise use of the passive methods. These techniques are described further in Section 1.7.1.

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1.7.1 Surfaces and Finishes

When EM waves are exposed to a surface, certain percentages are reflected, absorbed or trans-mitted through the surface as shown in Figure 1.11. These percentages are highly dependent on the material properties being exposed to the radiation as well as the wavelength of the EM waves that are being radiated. In most cases the EM waves are broken down into two groups that are most relevant for thermal control: the visible and the IR bands. The percentages of reflectivity, absorptivity and transitivity have the potential to be completely different for the IR and visible wavelengths. For the visible spectrum the percentage of radiation that is absorbed is called the absorbance (α) while for the IR spectrum the percentage absorbed is called the emittance (). The reason why it is called the emittance is because a material will also emit radiation at the IR frequency depending on the surface temperature, through the exact opposite process to absorbance [Gilmore, 2002]. These two values (absorbance and emittance) are given numbers between zero and one where one represents complete absorbance while zero represents complete reflection of the radiation.

Figure 1.11: Demonstration of absorption, transmission and reflection, look right, sorry I can’t highlight references [Mecart, 2010]

As different surfaces have different α and properties, one can reduce or increase the overall temperature of a spacecraft by varying these properties. This is not an active method of thermal control but it has the potential to dampen the overall oscillations of the heat fluxes experienced by the spacecraft.

Common Surfaces and Finishes

Thermal control surfaces are categorized into four common groups:

solar reflector,

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solar absorber,

flat reflector and

flat absorber.

These thermal control surfaces essentially exercise the four different options available for managing the relative absorbance and emittance. For example, the flat reflector will reflect both the IR and visible spectrum (α ≈ 0.1, ≈ 0.1) while the flat absorber will do the exact opposite and absorb everything (α ≈ 0.9, ≈ 0.9). The solar reflector will reflect the majority of the visible spectrum whilst absorbing the IR waves (α ≈ 0.1, ≈ 0.9) and the exact opposite will happen for the solar absorber.

For further information please refer to the “Thermal Control Handbook” [Gilmore, 2002].

Surface Degradation

The optical characteristics of a thermal control surface can change during a mission as a result of charged particles, ultraviolet (UV) radiation and high vacuum and contaminate films that can deposit on the outside of the spacecraft surfaces [Gilmore, 2002]. Most of these surface degradation effects result in an increase of absorptivity with minimal effects on the IR emittance. An increased absorptivity can lead to an increase in the solar flux that is absorbed by the material, thus increasing the temperature of the thermal control surface and potentially the spacecraft’s internal components. This is obviously a problem for the design of the thermal control sub-systems. Therefore, an understanding of these problems are important to help maintain the internal temperature of a spacecraft.

Contaminants This is mainly due to the particles or compounds that are out- gassed or emitted from a thrust plume forming on the surface of the craft and therefore darkening the surface over time. The majority of this contamination is minimised by choosing minimal outgassing material, building the majority of the components in clean rooms and protecting surfaces from exhaust plumes. [Gilmore, 2002]

UV radiation and Charged Particles UV radiation can initiate chemical reactions in materials that form other compounds on the surface. Similarly to contam- inates, this would change the optical properties of the material. Charged particles also have the ability to change the surface characteristics of a craft by causing chemical reactions in the material. Once again this would normally lead to an increase in the solar absorbance of the surface. These effects can be reduced by choosing materials that have a high resistance to these reactions .[Gilmore, 2002]

An Investigation into Thermal Design for Asteroid Landers

M A S T E R ' S T H E S I S

An Investigation into Thermal Design for Asteroid Landers

John Pegg Campbell

AN INVESTIGATION INTO THERMAL DESIGN FOR ASTEROID LANDERS

An Investigation into Thermal Design for Asteroid Landers

ABSTRACT

ACKNOWLEDGEMENTS

CONTENTS

LIST OF FIGURES

LIST OF TABLES

NOTATIONS

CHAPTER 1

LITERATURE REVIEW

1.1 Introduction

1.2 Thesis Aims and Objectives

1.3 Asteroids

1.4 Motivation for Sample Return Asteroid Mission

1.4.1 Scientific and Engineering Motivation

1.4.2 Cost Motivation

1.5 Asteroid Thermal Environment

1.5.1 Asteroid Surface Temperature

1.5.2 Radiation Fluxes

1.6 Spacecraft Properties and Constraints

1.6.1 Past, Present and Future NEA Lander Missions

1.6.2 Thermal Constraints

1.6.3 Heat Generation

1.6.4 Spacecraft Mission

1.7 Thermal Control Techniques

1.7.1 Surfaces and Finishes