Evaluation of Potential Propulsion Systems for a Commercial Micro Moon Lander

Konstantinos Papavramidis

Master of Science Thesis KTH School of Industrial Engineering and Management Energy Technology TRITA-ITM-EX 2019-231 Division of Heat and Power Technology SE-100 44 STOCKHOLM

Master of Science Thesis

TRITA-ITM-EX 2019:231

Evaluation of Potential Propulsion Systems for a Commercial Micro Moon Lander

Konstantinos Papavramidis

Approved Examiner Supervisor

Date Björn Laumert Nenad Glodic Commissioner Contact person

-2-

Authors

Konstantinos Papavramidis Aeronautical and Vehicle Engineering KTH Royal Institute of Technology

Place for Project

Stockholm, Sweden

Examiner

Björn Laumert Professor in Energy Technology Division of Heat and Power Technology KTH Royal Institute of Technology

Supervisor

Nenad Glodic THRUST Programme Director Division of Heat and Power Technology KTH Royal Institute of Technology Abstract

In the advent of Space 4.0 era with the commercialization and increased accessi- bility of space, a requirement analysis, trade-off options, development status and critical areas of a propulsion system for a Commercial Micro Moon Lander is carried out. An investigation of a suitable system for the current mission is examined in the frame of the ASTRI project of OHB System AG and Blue Horizon. Main trajectory strategies are being investigated and simulations are performed to extract the ∆V requirements. Top-level requirements are extracted which give the first input for the propulsion design. An evaluation of the propulsion requirements is implemented which outlines the factors that are more important and drive the propulsion design. The evaluation implements a dual comparison of the requirements where weighting factors are extracted, resulting the main drivers of the propulsion system design. A trade-off analysis is performed for various types of propulsion systems and a preliminary selection of a propulsion system suitable for the mission is described. A first-iteration architecture of the propulsion, ADCS and GNC subsystems are also presented as well as a component list. A first approach of the landing phase is described and an estimation of the required thrust is calculated. A unified Bipropellant propulsion system is proposed which fills out most of the mission requirements. However, the analysis shows that the total mass of the lander, including all the margins, exceeds a bit the mass limitations but no the volume limitations. The results shows that a decrease in payload capacity or the implementation of a different trajectory strategy can lower the mass below the limit. In addition, further iterations in the lander concept which will give a more detailed design, resulting to no extra margins, can drive the mass below the limit. Finally, a discussion on the results is done, addressing the limitations and the important factors that need to be considered for the mission. The viability of the mission due to its commercial aspect is being questioned and further investigation is suggested to be carried out on the ”micro” lander concept.

ii Sammanfattning

I tillkomsten av Space 4.0 era med kommersialisering och ökad tillgänglighet av rymden, en kravanalys, avvägningsalternativ, utvecklingsstatus och kritiska om- råden av ett framdrivningssystem för en kommersiell mikro månlandare bärs ut. En undersökning av ett lämpligt system för det aktuella uppdraget genomförs inom ramen för ASTRI-projektet för OHB System AG och Blue Horizon. Olika strategier för banor undersöks och simuleringar utförs för att extrahera ΔV-kraven. Topp-nivå krav definieras och ger den första inputen för designen av framdrivningssystemet. En utvärdering av framdrivningskraven implementeras och belyser de viktigaste faktorer som driver design av framdrivningssystemet. En avvägn- ingsanalys utförs för olika typer av framdrivningssystem och ett preliminärt urval av ett framdrivningssystem som är lämpligt för uppdraget beskrivs. En arkitek- tur för framdrivningen, ADCS och GNC-delsystem presenteras såväl som en kom- ponentlista. Ett första tillvägagångssätt av landningsfasen beskrivs och en upp- skattning av den nödvändiga dragkraften beräknas. Ett enhetligt Bi-propellant framdrivningssystem föreslås som uppfyller ut de flesta uppdragskraven. Anal- ysen visar dock att summan av månlandarens massa, inklusive alla marginaler, överstiger massbegränsningarna men inte de volymbegränsningarna uppsatta i projektet. Resultaten visar att en minskning av nyttolastkapaciteten eller genom- förandet av en annan banstrategi kan minska den totala massan då den inom gränsvärdena. Dessutom, ytterligare iterationer i månlandarens koncept som kom- mer att ge en mer detaljerad design, vilket resulterar i inga extra marginaler, kan leda till att den uppskattade massan minskar ytterligare. Slutligen förs en diskus- sion om resultaten, med hänsyn till de begränsningarna och de viktigaste faktor- erna som måste beaktas för uppdraget. Lönsamheten hos uppdraget på grund av sin kommersiella aspekt är ifrågasatt och vidare utredning föreslås utförs på ”mikro” månlandare konceptet.

iii Keywords

Propulsion, Moon Lander, Trajectory, Space, Commercialization of space, Space 4.0, OHB System, Blue Horizon, ASTRI program, Space exploration, Service-oriented space program

iv Acknowledgement

I would like to thank Nicolas Faber at Blue Horizon Sárl & OHB System AG for his continuous support during the ASTRI-OHB commercial lunar lander project. I would also like to thank Nenad Glodic, my supervisor, and Björn Laumert, my examiner, for agreeing to participate in this thesis project with me. Finally, I would like to thank Gunnar Tibert for coordinating the participation of KTH Royal Insti- tute of Technology in the ASTRI-program, which made this thesis project possible. Last but not least, I would like to thank my lovely partner, Tonya, for support- ing me in all aspects through all these years. Also my family and my closest rela- tives, who they supported me and helped me during my studies. Without all these people around me, obtaining not only my first MSc degree but also this second MSc degree wouldn’t be possible.

v Contents

1 Introduction 1 1.1 Background ...... 1 1.2 Purpose and Goal ...... 3 1.3 Outline ...... 6

2 Literature Study 7 2.1 Ways to the Moon ...... 7 2.2 Overview of the Space Propulsion Technology ...... 14 2.3 Review of past missions ...... 30

3 CMML Case 35 3.1 Approach and Method ...... 35 3.2 Designing Philosophy of the PS ...... 42 3.3 Preliminary selection of suitable PS ...... 56

4 Conclusions 77

vi List of Figures

1 ∆V requirements on different trajectories. [9] ...... 8 2 Hohmann Transfer. [19] ...... 9 3 Phasing loop Transfer. [13] ...... 10 4 Bi-elliptic Transfer. [14] ...... 11 5 WSB Transfer. [4] ...... 12 6 Low-Thrust Transfer. [13] ...... 13 7 Classification of Propulsion systems. [26] ...... 17 8 Schematic of a Solid fuel Motor. [21] ...... 18 9 Schematic of a Hybrid Propulsion Rocket. [9] ...... 19 10 Schematic of a Cold Gas System. [21] ...... 20 11 Schematic of a Mono-propellant System. [27] ...... 22 12 Schematic of a Bi-propellant System. [4] ...... 23 13 Electric propulsion configurations. [25] ...... 24 14 Thrust versus Specific Impulse on different PS. [25] ...... 28 15 ∆V performance of PS concepts. [21] ...... 28 16 GMAT simulation of Hohmann trajectory to the Moon. With red color is the transfer trajectory and in light blue is the TCM maneuver. In grey color is the Moon orbit around Earth...... 38 17 GMAT simulation of Moon arrival. With red color is the arriving circular lunar orbit and the landing phase. In light blue is the TCM maneuver while arriving on the Moon...... 39 18 Bi-elliptic trajectory...... 40 19 Requirements weighting...... 44 20 Dry mass in [kg] of the lander on Moon surface for different PS configurations...... 51 21 Total fuel in [kg] required for the mission for different PS configurations...... 52 22 Total mass in [kg] of the lander for different PS configurations. . . . 52 23 Lander’s wet mass for different payload mass...... 55

vii 24 Proposed engine configuration on the lander. Left: Bottom view of the lander. Main engine and assist thrusters are shown with red font. The ADCS thrusters are shown as black boxes from this view. Right: Isometric view of the lander. The ADCS thrusters configuration is shown with a representation of control on all 3-axes. The red arrows represent the thrust direction during firing with results in a control of an axis...... 57 25 Functional block diagram of the selected PS...... 60 26 Preliminary architecture layout of the ADCS & GNC subsystems. . . 64 27 Global landing reference system, [45]...... 70 28 Altitude as a function of time during the landing phase...... 73 29 Altitude as a function of the surface distance covered throughout the landing phase...... 73 30 Thrust angle, Thrust and Lander mass as a function of time during landing phase...... 74 31 Horizontal and Vertical velocities as a function of time during landing phase...... 75 32 Altitude vs Time.Zoom-in the last 2 phases of the landing...... 87 33 Altitude vs Downrange. Zoom-in the last 2 phases of the landing. . 87 34 Horizontal and Vertical velocities as a function of time. Zoom-in the last 2 phases of the landing...... 88

viii List of Tables

1 Summary table of PS characteristics. [21] ...... 30 2 Mission characteristics of Lunar Landers...... 31 3 Mission characteristics of Lunar Orbiters...... 34 4 ∆V Budget of Hohmann transfer...... 38 5 ∆V Budget of Bi-elliptic trajectory...... 39 6 Weight factors description ...... 43 7 PS requirement matrix...... 44 8 Lander mass budget (Including equipment level margins of each subsystem)...... 46 9 Required fuel in [kg] for every maneuver and for different chemical PS configurations...... 50 10 Required fuel in [kg] for every maneuver and for different electric- chemical PS configurations...... 51 11 Comparison of different configurations considering other important factors...... 53 12 List of preliminary components of the selected PS...... 61

13 KPI of Propulsion subsystem. Thrust, power and Isp values are extracted from [37] ...... 61 14 Mass and Volume of the PS. Main engine mass is extracted from [37]. The rest of the values are based on the equations from chapter 3.2. The pressurant tank volume is estimated 10% of the propellant tank volume based on [21]...... 62 15 Mass and Volume for each component of the PS...... 62 16 KPI of ADCS & GNC subsystems...... 65 17 Volume and Accuracy factors for each component of ADCS & GNC subsystems...... 65 18 Data required for the estimation of the average landing thrust. . . . 68 19 Final results of the simulation...... 72

ix Abbreviations

ADCS - Attitude Determination and Control System

ADN - Ammonium Dinitramide

ASAP - Ariane Structure for Auxilary Payload

CGS - Cold Gas System

CMML - Commercial Micro Moon Lander

CML - Concept Maturity Level

CoG - Center of Gravity

DMG - Design Maturity Gate

DSN - Deep Space Network

DOI - Descent Orbit Insertion

EDL - Entry, Descent and Landing

ESA - European Space Agency

FEEP - Field Emission Electrostatic Propulsion

GNC - Guidance, Navigation and Control

GS - Ground Station

GTO - Geosynchronous Transfer Orbit

HAN - Hydroxylammonium Nitrate

HET - Hall Effect Thruster

IMU - Inertial Measuerement Unit

KPI - Key Performance Indicator

LEO - Low Earth Orbit

LiDAR - Light Detection And Ranging

x LLO - Low Lunar Orbit

LV - Launch Vehicle

LOI - Lunar Orbit Insertion

LRO - Lunar Reconnaissance Orbiter

LTO - Lunar Transfer Orbit

MEMS - Micro-Electromechanical System

MMH - Monomethylhydrazine

MON - Mixed Oxides of Nitrogen

NASA - National Aeronautics and Space Administration

PPT - Pulse Plasma Thruster

PSLV - Polar Satellite Launch Vehicle

PS - Propulsion System

SMAD - Space Mission Analysis and Design

SOI - Sphere of Influence

TCM - Trajectory Correction Maneuver

TLI - Trans-Lunar Injection

TRL - Technology Readiness Level

UDMH - Unsymmetrical Dimethylhydrazine

ULA - United Launch Alliance

UPS - Unified Propulsion System

WSB - Weak Stability Boundary

xi Nomenclature

∆V - change in velocity, required impulse to perform a maneuver

T - thrust

Isp - specific Impulse

Issp - system-specific Impulse

Itot - total Impulse

ve - effective exhaust velocity

m˙ - mass flow rate

Pe - nozzle exit pressure

Pa - ambient pressure

Ae - nozzle exit area

mi - initial mass

mf - final mass

mPS - mass of the propulsion system

g0 - gravitational acceleration of Earth

gMoon - gravitational acceleration of Moon

N2 - nitrogen

Ag - argon

NH3 ammonia

N2H4 - hydrazine

H2O2 - hydrogen Peroxide

N2O - nitrous Oxide

C - fill ratio, Vp/Vt

xii mp - propellant mass

Vp - propellant volume

Vt - tank volume mT - tank mass mpr1 - pressurant mass in propellant tank mpr2 - pressurant mass in pressurant tank mTpr - pressurant tank mass mADCSpropellant - ADCS propellant mass mresidual - residual mass mcase - solid motor case mass

Vcase - solid motor case volume

Pc - solid motor chamber pressure

Pop - operational pressure

ρ - density

M - molecular weight

Kp - tank performance factor

Kpr - pressurant tank performance factor zp - gas compresibility factor propellant zpr - gas compresibilty factor pressurant

T - temperature

H - periselene, start of descent

D - downrange

G - gravitational constant

xiii mMoon - mass of Moon

RMoon - radius of Moon mlander - mass of lander at start of descent

Favg - average required thrust

Wnet - total work done

Wgravity - work done by gravity

Wthrust - work done by thrusters uorbit - orbital velocity

Korbit - kinetic energy in orbit

∆y - altitude difference

xiv 1 Introduction

1.1 Background

Few decades ago, the lunar exploration started again to get the interest from space community, but the major players are still unchanged. Government-driven entities are the key players in deep space missions where USA and Russia have the most of them since their start about six decades ago. Europe, Japan, India and China are the new entries in the lunar exploration competition that has restarted. Main reason that only governmental entities have conducted such missions until now is the prohibitive cost involved, costing hundreds of millions, in combination with the low interest of lunar missions after the first successful human landing on the Moon in 1969 with Apollo mission. After this long stagnation in Moon projects, it was in early 1990s where the start of interest in Moon took place, again with government funded missions. Though, the mission goals were low-key compared to the huge efforts during the space race in the 1960s. The primary cost in such projects remains the launch cost which is the reason of the exclusivity of space by the national agencies. The most competitive launch price is offered by SpaceX at $62 million while a lot of companies with smaller launchers, and thus cheaper, start to pop out. Still, the launch offered by United Launch Alliance (ULA), which is a more established company, remains high at $225 million. The high cost of the missions is compromised by the efforts to ensure that whatever goes to space, it will work when it gets there; use of high quality and high assurance components, implementation of redundancy, extensive testing and fastidious project planning are some of the elements that increase the overall mission cost. [1][2][3][4] Commercialization is the way to a more affordable and accessible space. After 2000s, there were a big activity in Moon projects where new players, both national and commercial, came into play. Missions such as the European SMART-1, the Chinese Chang’E-1 and Chang’E-2, the Japanese SELENE-Kaguya, the Indian Chandrayaan-1, and the Lunar Reconnaissance Orbiter (LRO) from the U.S were the start of the new era of lunar exploration. The recent lunar exploration programs are focusing on the rapid development of low cost and high-level technologies. Companies are seeking to make profit in space industry and they are will- ing to take more risks, push for innovation and competitive prices compared to

1 the government agencies. Reasons on that are the slow growth in space industry being reflected on administrations and bureaucracy related to the governmental agencies while new start-ups usually tend to take more chances aiming to achieve better results faster. [3][4][5][6] A decade ago, the Google Lunar X-Prize, an international competition among private enterprise teams gave the opportunity to a new age of commercial access to space. The goal of the competition was the development of a Lunar Lander, able to make a soft landing, carrying a rover which is capable of traveling 500m, taking pictures and send data back to Earth. As it was aforementioned, push for innovation and competitive prices are some of the main assets that companies can provide, however they can also produce some technological breakthroughs in the space industry. For instance, SpaceX has shaken up the launcher suppli- ers with their reusable launch vehicles, causing a dramatic cost reduction. The commercial services will not encompass just asteroid mining or advanced world- wide communication, but also platforms carrying payloads to space. After the Lunar X-Prize remained unclaimed, resulted several teams to continue develop- ing their lander, offering a commercial transportation platform. These service- oriented transportation platforms can spread out the costs (e.g. development, launch, etc.) in different missions and different customers. Attention to concepts such as reusability and standardization can keep the overall cost down. [1][5][6] [7] There are several reasons on the restart of lunar exploration. First, programs on space exploration are usually accompanied with new technology demonstrations and produce new science. National endeavors can contribute to the advance- ment of technology and science of a particular nation, but commercial endeavors can have a global impact on these fields as well. Second, the Moon presents a scientific interest related to the solar system and Earth. There have been conducted a lot of studies about the Moon (second most known celestial body after Earth) which is linked with the origin of Earth and how the Earth-Moon system created. Third, the Moon offers resources that could be useful in the future in the vision of human settlements. For example, water resources on the Moon surface could be used as a fuel for future launches or as a life-support resource for humans living and working there. [5][6]

2 From this discussion and for all these reasons, Moon has taken the attention of private entities with the vision of making the Earth and Moon one system and expanding the humans’ living sphere into space [6]. Thus, providing a low-cost platform and finding the market gap on the different competitors, it is an important factor in having a competitive advantage in this new era of space exploration.

1.2 Purpose and Goal

The new space race in Moon exploration and its commercialized nature are the basis of the kick-off of the current project. The present thesis is part of this project, titled “Feasibility study to develop a Commercial Micro Moon Lander (CMML)” which is conducted by OHB System AG and Blue Horizon in the frame of the AS- TRI program . The main goal of this project is to provide a platform that can carry payloads to the Lunar surface to potential customers who want access to the Moon and making profit in a sustainable way. The purpose of the present work is to provide a solution regarding the propulsion sub-system. What makes the project different is its commercial application, compared to the ones run by governmental agencies. This will drive the propulsion system design to implement a different mindset where low cost and size compatibility to a micro lander will play the most important roles. Having said that, the ASTRI project organization needs to be outlined, linking it with the present thesis since this will give a clear picture of the present work. The project is following the standardised NASA system lifecycle, extending from Pre-Phase A (or Phase 0) to Phase F. It will use the concept of CMLs (Concept Ma- turity Level) and DMGs (Design Maturity Gate) to measure progress throughout the project lifecycle (more information on the CML concept can be found in [8]). In the ASTRI frame, the project will last 18 months with the goal of completing the Phase A, ideally reaching the Phase B (CML 7) by October 2019. These phases that are intended to be covered are known as the “Formulation” stage. This covers the development of the project, designing concepts, creating and evaluating designs. The team consists of 8 people where each of them are assigned to different subsystems of the lander as well as the mission architecture. The 18 months are split in 2 parts where the first 6 months of the project the present thesis will be carried out and the remaining 12 months will be the continuation of the project in the

3 company as graduates. Thus, during the first 6 months the project is at the Pre- Phase A (or Phase 0) which gives the frame of the present work that is required to be accomplished. In particular, during this phase the mission starts from zero and a preliminary concept of the mission needs to be established in order to build a reference mission. This first iteration will give the necessary input for the conceptual design of the propulsion system as part of the mission. As the mission will be developed, the subsystems design (including the propulsion subsystem) will go into a more detailed design which will take place in the next phases.

Thesis objectives

The objectives of the thesis are outlined below, providing the frame of the present work. The purpose of the thesis is to provide the basic guidelines and solutions regarding the propulsion sub-system and produce all the necessary requirements which fulfill the mission requirements ensuring the nominal performance of the spacecraft. The commercial aspect is the main trait of the mission which makes it different and is addressed in the present thesis. Thus, the thesis will focus mainly on the top-level design of a propulsion system, investigating different concepts and proposing a preliminary design. This means that the thesis will provide an overview and will keep the design on a system level, not intending to go into a detailed design in this premature stage of the project, as mentioned above. Below, the main characteristics of a mission will be presented, including the mission statement, the needs-goals-objectives and the top-level requirements derived on the first iteration of the mission concept. Based on these requirements, it will be deduced the requirements of the propulsion system.

Mission statement

To provide continuous and affordable access to the entire lunar surface for rapid and iterative technology demonstrations, commercial endeavours, and to promote the development for a future permanent lunar settlement.

4 Needs, Goals, Objectives

N: To have affordable access to the lunar orbit and surface in order to enhance exploration of the Moon. G1: To transport a broad range of payloads to the lunar orbit and surface for as large a customer base as possible. G2: To encourage the development of technology for permanent human settlement. O1: To develop a commercially viable micro moon lander capable of a high frequency launch rate at a low cost. O2: To be capable of transporting a wide variety of payloads following a standard safely to the Moon.

Top-level requirements

ID Requirement The lander shall carry a payload with a total mass of around 10 kg, total M_R-01 volume of around 10 U and with a total power consumption of around 10 W. The lander shall perform a soft landing on the lunar surface. M_R-02 The maximum landing speed shall be around 2 m/s. The lander shall have a high-frequency launch rate of at least one launch M_R-03 per month. M_R-04 The lander shall survive at least for an entire lunar day (14 earth days). M_R-05 The lander shall be able to deploy payloads on the lunar surface. The lander shall be able to deploy payloads and CubeSats in lunar orbit M_R-06 whether other payloads are going to the lunar surface or not. The maximum launch mass of the lander shall be 400 kg and maximum volume of M_R-07 the lander shall scale within a cylinder of 1500mm diameter and 1500mm height. The ∆V budget required to reach the Moon surface shall be around 3200 m/s M_R-08 (excluding safety margins). The lander design shall include modular subsystems following a procedure M_R-09 such as the Cubesat model.

5 Requirement M_R-08 is based on Figure 1 [9]. It acts as a first-guess estimation which gives the initial input. In chapter 3.1, a trajectory simulation is carried out which results in a better estimation of the ∆V required. The requirement states that the value shall be around 3200 m/s which is validated by the results produced in chapter 3.1. Requirement M_R-07 is based on the ASAP configuration [10] and derived from the selected launch scenario (Chapter 3.1). Requirement M_R-02 is an average speed for a soft landing based on previous missions (e.g. Chang’e 3 [11]). The rest requirements are based on other commercial and mission aspects where the analysis and details are out of the scope of the present work. In general, all the top-levels requirements provide the input for the design of each subsystem of the lander, including the propulsion system.

1.3 Outline

After the introduction, the thesis moves on in Chapter 2 describing the past lunar missions, orbiters and landers, focusing in the selected trajectory and the choice in the propulsion system. In addition, a discussion on the different approaches to get to the Moon surface will be added as a part of this chapter. An overview of the basics on the propulsion principles and the state of the art of the space propulsion systems is also given in this chapter, with a discussion on the main advantages and disadvantages of each type of system and the main applications of each system to be included. In addition, an overview of available technology (”off-the-shelf” components) in order to keep a low cost will be included. Then, the main work of the thesis will be included on Chapter 3 and it will consist of 2 parts: first part will start with some basic trajectory simulations, creating the ∆V budget and the maneuver envelope. The second part will present the trade-off study between different propulsion options, the requirements analysis of the propulsion system for the current mission, the selection of a suitable system and its mass budget. Also, the interfaces of the propulsion system with the ADCS (Attitude Determination and Control System) and GNC (Guidance, Navigation and Control) systems will be introduced in this chapter. A description as well as a simulation of the most critical phase of the mission, the landing phase, is presented afterwards. Finally, Chapter 4 will include the conclusions of the present work and discuss the further studies that needs to be carried on the mission.

6 2 Literature Study

In this chapter, the discussion on 3 main topics from literature is presented. First, the different approaches on getting to the Moon are described and the main im- plications are discussed. Second, a description on the various types of propulsion systems to date is carried out and their different applications are stated. Third, an overview of the past lunar missions will be presented, giving a better picture on the trajectory approach and the selection of the propulsion system.

2.1 Ways to the Moon

There are different approaches to get from Earth to the Moon. A typical diagram of ∆V requirements for Earth and deep space trajectories is shown in Figure 1. As it can be seen from the graph, an approximate ∆V budget to get to Lunar surface is 3.2 km/s. In general, the trajectory can be split in 3 major parts, [12]:

• Leaving the Earth

• Transfer from Earth to the proximity of the Moon

• Approaching the Moon

Regarding the ”Leaving the Earth” phase, there are 2 main ways to launch: the first is to launch directly to a trans-Lunar trajectory and the second is to go to a parking orbit around the Earth and then, after a coast time, to perform a maneuver that will put the spacecraft into a trans-Lunar trajectory. The second method is called ”Launch-Coast-Burn”. These 2 different approaches are dependent mainly from the capability of the launch vehicle. For example, there are launchers such as Falcon 9 and Ariane 5 that can put a spacecraft into a trans-Lunar trajectory compared to smaller launchers, like the Indian PSLV (Polar Satellite Launch Vehicle), which have the capability to launch in Low Earth Orbit (LEO) and Geostation- ary Transfer Orbit (GTO). The first missions during 1960s used the direct launch to Lunar Transfer Orbit (LTO) but soon the next missions switched to Launch- Coast-Burn method. The latter method was proved to be more efficient from a mission design point of view compared to the former method, since it was allowing to launch on any day of the month. [12][13][14]

7 Figure 1: ∆V requirements on different trajectories. [9]

Next, for the ”Transfer from Earth to the proximity of the Moon” phase, there are several types of transfers that can be utilized. The ones that have been used to date are, [3][4][12][13][14][15] Larson1999 [16][17][18]:

• Hohmann transfer, see Figure 2.

• Phasing loop, see Figure 3.

• Bi-elliptic, see Figure 4.

• Weak Stability Boundary (WSB), see Figure 5.

• Low-thrust, see Figure 6.

The main differences among these kind of transfers are mainly the direct and indirect ”nature” of the trajectory and the mission time it takes to reach the Moon. The Hohmann transfer is the most known and most used maneuver and it has been implemented in most of the missions due to its simplicity and reliability. This type of transfer is an ellipse with the perigee close to the Earth and the apogee close to the distance of the Moon. It is a direct and fast transfer but expensive in

8 terms of ∆V with a transfer time between 2 and 5 days. The lowest ∆V is required for a Hohmann transfer when the apogee of the trans-Lunar orbit is equal to the Earth-Moon distance. To reduce the transfer time (less than 5 days), the apogee is needed to be further increased with a small penalty in the ∆V. The conditions to perform this kind of transfer are valid when the declination of the Moon is smaller than the inclination of the parking orbit (which corresponds to the latitude of the launch site). The worst case inclination of the Moon’s orbit is between -28.6◦ and

28.6◦. So, the preferred launches should be performed from a region with this latitude, such as Cape Canaveral (28.6◦ latitude), where there are 2 launch opportunities per day. All the Luna and Apollo missions during 1960s to 80s used the ”classical” Hohmann transfer. [3][4][12][13][14][15] Larson1999 [16][17] [18]

Figure 2: Hohmann Transfer. [19]

A different strategy to the Hohmann transfer is the so called ”Phasing loop” transfer that simply uses the classical Hohmann transfer but including more steps. In this type of transfer, the orbit is gradually raised by the parking orbit till to reach the distance Earth-Moon. In particular, additional burns in the perigee of

9 the parking orbit are performed, raising the apogee to the desired distance. Typ- ically, a phase loop transfer of 2.5 or 3.5 is used and the time to reach the target is increased to 2-3 weeks compared to the direct Hohmann transfer. With this strategy, the launch window can be extended, giving the time for checking the operational conditions of the spacecraft and performing possible orbit corrections or other minor fixes. Also, it can minimize the losses due to the non-impulse nature of the burns. [3][4][12][13][18]

Figure 3: Phasing loop Transfer. [13]

An indirect method of transfer is the ”bi-elliptic” and it is used when the ge- ometry of the parking orbit, such as GTO, is not compatible with the target orbit. In a situation like this, a plane change maneuver is required in order to achieve the desired orbit. In general, plane change maneuvers are very costly in terms of ∆V, leading to several kilometers per second. However, the cost in ∆V can be reduced if the velocity of the spacecraft is low. The velocity of the spacecraft is low in apogee. The bi-elliptic scheme is consisted of two half elliptic orbits. In the first half, a transfer as far away as possible from the Earth, typical apogee of 1 million kilometers, is performed. This is the TLI (Trans-Lunar Injection) maneuver, resulting in an apogee with a low velocity. The second half connects the end of the

10 first half (apogee) with the Moon (perigee). In this second half, a plane change and orientation maneuver is performed at the apogee in a such a way that at the perigee of the orbit the Moon is encountered. This is the ”mid-course” correction maneuver, where a small penalty in plane change is achieved in the expense of an extended mission time, typical 50-60 days. It is important to note that with this method the ∆V needed for the LOI (Lunar Orbit Insertion) maneuver is compa- rably lower than a direct transfer since the spacecraft arrives at the perigee of the trans-Lunar orbit where the velocity is higher and the relative velocity with respect to the Moon is lower. Bi-elliptic transfers from a parking orbit, such as GTO, are more complex but are less dependent on appropriate launch windows. [13][14] [16][17]

Figure 4: Bi-elliptic Transfer. [14]

The fourth method discussed here, is the so called ”Weak Stability Boundary” or WBS which can reduce even more the ∆V. The WSB method aims to reduce the total ∆V of the transfer instead of lowering the ∆V of just the TLI maneuver. The way of doing that is based on ”stealing” orbital energy from other bodies, such as the Sun and the Moon. This can be achieved by raising the apogee of the parking orbit, for instance the GTO, to a distance of approximately 1.4-1.5 million kilometers away from Earth to a region that is called Weak Stability Boundary where the gravity of the Sun is in the same order with the gravity of the Earth. These re- gions are typically found close to the Lagrangian points. A small maneuver in that

11 region can lead to significant variation in the Moon arrival conditions. Arriving on that region, the strong perturbations of the Sun can drive a spacecraft to the Moon orbit without the need of a thrust burn as it was for bi-elliptic transfer. This is achieved with the raising of the perigee to a radius equal to Earth-Moon distance. Upon the transfer to the Moon, the arrival to the Earth-Moon WSB region can be used to tune the arrival condition and to acquire a ballistic capture at the Moon by taking advantage the Earth’s gravity. In such a way, it can be attained a ∆V saving of 25% in the expense of a greater TLI thrust burn maneuver compared to bi-elliptic. The result is that a lower total ∆V than bi-elliptic can be achieved, however, raising greatly the mission time with typical values of 2-4 months. In addition, this type of transfer is complex and it requires a lot of control due to its unstable trajectory nature. [12][13][14][16]

Figure 5: WSB Transfer. [4]

Finally, the ”low-thrust” transfer is a spiral type trajectory due to the very low thrust levels and it is accompanied by the use of electric propulsion. This method is more complex compared to the previous ones and the mission time to reach the Moon, the highest. Unlike chemical propulsion where the orbital maneuvers are executed by short bursts, low-thrust transfer requires long burns. In general, this type of transfer is typically consisted of a part with continuous thrust, a part of

12 thrust arcs and a set of lunar gravity assist. The first part, the continuous thrust, is performed to mainly raise the perigee of the orbit as well as the apogee in a second place. Next, the thrust is performed in part of the orbit, around the perigee region, raising the apogee while spiraling out to reach the Moon distance. Also, during this non-continuous thrust part results in raising the perigee distance too. After finishing this phase, a ”resonant orbit” is achieved, meaning that the spacecraft encounters the Moon in the same place after a specific number of orbits. Last, the gravity assists are used to take advantage of the gravitational pull of the Moon in order to set-up and tune the trajectory and the arrival conditions, for instance performing plane changes and perigee raising such that the spacecraft meets the Moon in the final orbit with a close to zero C3 (characteristic energy), allowing for a low-thrust orbit insertion maneuver. [3][4][13][15]

Figure 6: Low-Thrust Transfer. [13]

The last part, the ”Approaching the Moon” phase, is consisted of 2 different methods that usually are utilized. The first is a direct Lunar descent where the trajectory is such that it aims for a direct lunar landing. Usually these trajectories are impact trajectories meaning that the spacecraft possesses a big velocity when it encounter the Moon and it requires a huge amount of ∆V in order to brake. The

13 first missions to the Moon where they didn’t require a soft landing but rather they went to an impact (or hard landing) trajectory, they used this kind of transfer. The second method, which is most used over the last lunar missions, consists of the LOI maneuver close to the periselene (closest point of the orbit with respect to the Moon) where the spacecraft is captured by the gravity of the Moon in an elliptical orbit (a Lunar parking orbit). After coasting for some orbits or circularizing the orbit (this is also called phasing loops) by performing successive burns to lower the aposelene (farthest point of the orbit with respect to the Moon), a Descent Orbit Insertion (DOI) maneuver is performed to lower the periselene and start descending to the lunar surface. [3][4][12][13][14]

2.2 Overview of the Space Propulsion Technology

In this chapter the available and various space propulsion technologies are described. Propulsion systems are crucial for any space mission and usually occupy a big part of the whole spacecraft. They are used to provide the required thrust and to perform all the necessary maneuvers during a space mission. These maneuvers are usually consisted of major maneuvers such as orbital transfers and minor maneuvers such as attitude control and orbital control, [20][21].

2.2.1 Propulsion principles

Propulsion is based on the third Newton’s law which is described as: ”for every action there is an equal and opposite reaction”. A propulsion system utilizes an expelled mass or fuel principle, where the fuel is ejected at some velocity in one direction, creating a force (thrust) in the opposite direction. For chemical type of systems, this mass ejection involves the production of high temperature and high pressure products due to the combustion and their acceleration to supersonic speeds in a convergent-divergent nozzle. In case of ion propulsion, the thrust is produced by the acceleration of charged particles (plasma) through an electric field. [9][15][20][22][23] The 4 main performance factors that characterizes the propulsion systems are: thrust (T), specific impulse (Isp), effective exhaust velocity (or exhaust velocity)

(ve) and Delta-V (∆V). These factors are essential when design a propulsion sys-

14 tem. The thrust, as shown in Equation 1, is dependent of the mass flow (m˙ ) and the exhaust velocity (ve) as well as from the exit pressure (Pe), ambient pressure

(Pa) and the nozzle exit area, (Ae). In the vacuum of space the ambient pressure is close to zero and in combination with typical small exit pressure the last term is in general close to zero and the thrust is given mainly from the first term of the Equa- tion 1. In Equation 2, the specific impulse is dependent of the thrust (T) and the mass flow (m˙ ) and it gives a measure of the efficiency of the engine. It is defined as the impulse (integral of thrust over time) delivered per unit mass of fuel. The exhaust velocity is the velocity in the exit region of the engine and is defined as the product of the specific impulse and gravity acceleration, as shown in Equation 3. Equation 4 is the well-known ”Rocket equation” by Konstantin Tsiolkovsky and relates the exhaust velocity with the initial (mi) and final mass (mf ) of the spacecraft. The equations that describes the basic performance factors are presented below, [9] Larson1999 [20][23][24][25]:

T =˙mv +(P P )A mv˙ (1) e e − a e ≈ e

T Isp = (2) mg˙ 0

ve = g0Isp (3)

mi ∆V = veln (4) mf One additional factor that describes the performance of the entire spacecraft propulsion system and helps in the evaluation of the propulsion on system level is the so called ”System-specific impulse (Issp)”. This factor takes into account all the different components of a propulsion system, such as tanks, power supply and power processing systems in case of electric propulsion, that can form a major

”dead” dry mass, compromising the overall performance of the system. Thus, Issp can act as a complementary aid to the Isp factor in the choice and sizing of the

15 propulsion system. The Issp factor is described by the equation:

Itot Ns Issp = [ ] (5) mPS kg where the Itot is the total impulse delivered by a certain quantity of fuel and mPS is the mass of the entire propulsion system (PS). [21]

2.2.2 Types of propulsion

The various types of propulsion systems can be categorized in terms of fuel type, size, basic function, type of energy source or method of thrust production, [22]. A typical classification of the propulsion systems is illustrated in Figure 7. The present work is focused on the state-of-art space propulsion systems that typically used on previous and current missions. With this in mind, the main types of propulsion considered are 3:

• Solid fuel propulsion

• Liquid propulsion

• Electric and Ion propulsion

In the next paragraphs, each of the category is described and emphasis is given on their typical characteristics, applications and advantages/disadvantages.

Solid fuel propulsion

A propulsion system using solid fuel is consisted of a cylindrical motor case with an empty central core , a nozzle and an igniter. The main characteristic of a solid rocket propulsion system is that both fuel and oxidizer are stored in the thrust chamber. There are 2 types of fuels, [21]:

• homogeneous fuel, which are fuels that contain a sufficient amount of oxy- gen to sustain the combustion process.

• composite fuel, which are a mixture of powdered metal (fuel), crystalline oxidizer and a polymer binder.

16 Figure 7: Classification of Propulsion systems. [26]

Common combination of fuel/oxidizer is synthetic polymers and ammonium perchlorate. The basic principle of this system is the start of the combustion of the fuel induced by the igniter and the production of high temperature and high pressure gas that is guided through the empty central core to the nozzle. Then, the gas is expanded through the nozzle and produces the thrust to move the spacecraft. For redundancy reasons, more than one igniters are employed. The thrust capability of these systems is high, with typical thrust value of 106N. However, they offer a moderate specific impulse, with typical values of 250-280s. A disadvantage of such systems is that they are ”one-shot” devices, meaning that once starts to burn it will stop until the whole fuel is used and thus they are lacking of controllability and restarting capability. A solution to mitigate this problem is, for instance, the use of a constellation of hundreds of Solid fuel micro-thrusters that can be fired simultaneously or in sequential mode, providing every time the required thrust. These technique uses MEMS (Micro-Electromechanical System) technology to provide a miniaturized system. On the other hand, the design of the solid fuel rockets is simple and compact, the fuel/oxidizer is premixed and since it is in solid state there is no sloshing, the motor cases are usually lightweight and they also offer great packaging capability and long-term fuel storage, [9][15][21] [22][23][24][26]. A typical graph of a solid fuel motor is shown in Figure 8.

17 Figure 8: Schematic of a Solid fuel Motor. [21]

The solid propulsion systems have a high thrust capability with a moderate specific impulse level and deliver the total impulse in one firing, making them suitable for major orbital maneuvers. Characteristic applications of a solid rocket system is as a launcher’s booster, an apogee or perigee kick motor, an upper stage of a launcher that put the spacecraft in the desired orbit. In addition, they have been used as a separable stage of a spacecraft for braking maneuvers, for instance in a LOI maneuver, [9][15][21][22][26]. Hybrid propulsion is an alternative modification to solid propulsion, giving a solution to some issues of these systems. The main principle of the system is the combination of solid fuel with an oxidizer in liquid state (usually oxidizers that are used in bi-propellant systems). The oxidizer is fed into the combustion chamber of the solid rocket where it reacts with the solid fuel and expanding the gas products through a nozzle. These systems offer the design simplicity and robustness of a solid motor with an improvement in performance and controllability that is found in bi-propellant propulsion. In addition, this system provides a restarting capability and an increase in thrust in the expense of increased complexity. Fi- nally, these systems have not been used in spacecraft propulsion so far, [9][24] [26]. In Figure 9, a typical schematic of a hybrid system is shown.

18 Figure 9: Schematic of a Hybrid Propulsion Rocket. [9]

Liquid fuel propulsion

The liquid propulsion systems can be categorized in 3 main groups, [21][22]:

• cold or warm gas systems

• mono-propellant systems

• bi-propellant systems

The description of the main characteristics of each system is discussed below. Cold gas systems is comprised of a high pressure gas that is expanded through a de Laval nozzle to produce the thrust. Common fuels that are used is Nitrogen

(N2), Argon (Ag), hydrocarbons like ammonia (NH3) and propane among other. A typical system consists of a high pressure tank (200-300 bar), nozzles, valves, pressure regulators, pressure transducers and filter. This kind of system is the simplest, low size and reliable (due to low storage pressure) but is limited in specific impulse (10-80s) and delivered ∆V. Low storage pressure enhances the reliability of the system as well as the use of non-toxic fuel and decreases the design cost. Use of liquid fuels reduces the storage volume but they impact the stability of the spacecraft with the effect of sloshing. To mitigate this problem, anti-sloshing baffles technology is used in modern systems. One of the main disadvantages is the decrease of the thrust and specific impulse over time because of the decay of pressure inside the tank (due to use of fuel) and usually an extra pressure regu- lator is used to provide constant pressure and avoid this issue. Thus, this type of system is limited by the pressure inside the tank and typical thrust values ranges from 0.02 to 10N. Typical applications of cold gas systems are for attitude and orbit control due to the capability of small impulse bits and large number of impulse

19 cycles which can offer precise attitude control maneuvers. [15][21][22][24][25] [26]

Figure 10: Schematic of a Cold Gas System. [21]

In warm gas systems, the main difference is that the gas is additionally heated before expansion by an external source, increasing the specific impulse and thrust compared to cold gas. External source can be an electric resistant or the use of solar energy which heats the fuel directly, [15][21][22][24][25][26]. A typical cold gas configuration is shown in Figure 10. Mono-propellant propulsion systems utilize a single fuel (no use of oxidizer) to create thrust. Most used fuel on these kind of systems is Hydrazine

(N2H4) where is decomposed through a catalyst, producing a high temperature and pressure gas which is then expanded through a nozzle. Common examples of catalyst is platinum and solid manganese dioxide. Usually, nitrogen gas is used to expel the fuel from the tank to the catalyst. Other propellants used are hydrogen peroxide (H2O2) and nitrous oxide (N2O). Nowadays, ”green” fuel (non-toxix),

20 alternatives to hydrazine, are employed to spacecraft propulsion with similar performance characteristics while reliability and lower cost is increased. Typical examples are the Ammonium Dinitramide (ADN) and Hydroxylammonium nitrate (HAN). Major disadvantage is the required higher preheat temperature, leading to a higher required power. Concerning small spacecrafts, unstable and energetic fuel can have a negative effect on launch opportunities as secondary payload due to safety concerns, making the development of ”green” propellants a necessity. Mono-propellant systems offer moderate thrust levels (up to 22N) and specific impulse up to 220s. They are relative simple systems (compared to bi-propellant) and reliable. Other aspects of mono-propellant systems is their low cost and the capability of producing low thrust (>0.5N), offering more precise impulse bits. Thus, they are most used in attitude and orbit control, [9][15] Larson1999 [21] [22][24][25][26]. A typical schematic is shown in Figure 11. Bi-propellant propulsion systems are comprised of a fuel and oxidizer stored in separate tanks. The propellants are only mixed in the combustion chamber where they react spontaneously (hypergolic), producing a gas with high temperature and pressure and expanding through the nozzle. These systems give higher performance in terms of specific impulse (up to 320s) and thrust (typical 500N for space propulsion) in the expense of higher system complexity. Typical fuel combination are Monomethylhydrazine (MMH) or Unsymmetrical dimethylhydrazine(UDMH) with Dinitrogen tetroxide (N2O4, MON) or a less toxic combination of kerosene with Hydrogen peroxide (H2O2). Also, in bi-propellant systems ”green” propellants start to be employed for the same reasons explained in the previous paragraph. Main components of such a system are fuel tanks (with propellant management devices), gas-pressurant tank, fuel pipes, thrusters, check valves, relief valves, pressure regulators, filters, fill and drain valves, pyro- or latch valves and pressure transducers justifying the increased complexity of the system. A schematic of a bi-propellant systems is shown in Figure 12. Another implication of these systems is the increase of cost due to the 2 different fluidic feed systems that add on complexity. However, the great performance characteristics compared to the other systems make them suitable for missions with required lower mass and increased payload capabilities. They are a suitable choice for mis-

21 Figure 11: Schematic of a Mono-propellant System. [27] sions involving major orbit changes such as interplanetary transfer and they are also used in telecommunication satellites around Earth. In addition, a significant merit of bi-propellant systems is that can offer a Unified Propulsion System (UPS). The UPS shares a common fuel tank and provide the capability of producing multiple thrusts, covering different requirements such as orbital transfers, attitude control and orbit control. Bi-propellant thrusters for attitude control can produce thrust, starting from 10-20N, offering control strategies but lacking of high precision compared to other systems such as mono-propellant and cold gas. [9][15] Larson1999 [21][22][24][25][26]

22 Figure 12: Schematic of a Bi-propellant System. [4]

Electric and Ion propulsion

In this type of propulsion, the main difference compared to chemical propulsion is the external supplied energy to the fuel. Thus, there is a variation with respect to their performance limitations. While chemical propulsion systems are characterized as energy limited systems due to the limit in the energy released from the propellants, electric propulsion systems are power limited since the fuel kinetic energy that can be produced is dependent on the mass and efficiency of the energy conversion equipment that transforms the external source of energy (solar or nu- clear) to electrical energy. The major categories on electric propulsion are: [21] [23][24][25]

23 • Electrothermal systems

• Electromagnetic systems

• Electrostatic systems

Typical configurations of these systems are illustrated in Figure 13.

Figure 13: Electric propulsion configurations. [25]

In electrothermal systems, the fuel is heated electrically and then is accelerated by expansion through a nozzle. These systems are mainly divided in 2 categories: resistojet and arcjet. [9][23][24][25] In Resistojets, the fuel is heated by an electric heating element (e.g. resistance), raising the temperature beyond the stagnation temperature and leading to an increase of the specific impulse. The basic principle of resistojets is the same as a CGS (Cold Gas System) with the only difference the fuel heating before nozzle expansion. In terms of thrust, they offer small values of thrust within a range

24 of 0.005-0.5N and are more suitable for attitude control where they have been usually used in previous missions. The specific impulse is greater than CGS with typical values of 300s. Common propellants that are used with this system are

Hydrazine (N2H4) and Ammonia (NH3). As in any other liquid fuel system, they experience issues with sloshing. [9] Larson1999 [24][25] On the other hand, Arcjets are capable of exceeding the heating temperature of the typical heating elements that limits the resistojets, producing higher exhaust velocity. An constant arc discharge is created in the throat of the nozzle by a cathode and the fuel is heated while it passes through the arc and expanded through the nozzle. Higher performance comes in the expense of higher complexity for arcjets due to the difficulties on sustaining a steady arc discharge and higher power demand. However, they provide a high specific impulse compared to chemical propulsion, typical value of 800s, and moderate thrust levels within round 1N.

Similar to resistojets, Hydrazine (N2H4) and Ammonia (NH3) are common used propellants and sloshing effects are experienced in this type of system. Again, main application for arcjets is for attitude control due to their limited thrust. [9] Larson1999 [24][25] In Electromagnetic systems, a discharge arc is applied to a gas, heating it to high temperatures and convert it to neutral plasma. The plasma then is accelerated to high velocities via the interaction between the induced magnetic field and the discharge current, producing the thrust (Lorentz force) to the spacecraft. [23] [24][25] Magneto-Plasma-Dynamic (MPD) thruster is a typical application of an electromagnetic system and combines the magnetogasdynamic with the electrothermal arc jet technologies. The plasma is accelerated by the arc heating and the electrodynamic forces. However, the dominant acceleration mechanism is the induced magnetic field. These type of systems are very power-demanding and thus the propulsive efficiency is low ( 10%) at low power levels (1kW). At higher power ∼ levels (MW range) the efficiency is increasing significantly ( 40%). Cathode ero- ∼ sion, as in electrothermal systems, is a major life limitation characteristic. Never- theless, they offer a thrust capability of a few Newtons with a high specific impulse ( 1500s). [9][24][25] ≥ A variation of an MPD thruster is the so called Pulse Plasma Thruster (PPT).

25 These thrusters use a solid fuel (usually Teflon) which is heated, again, under a discharge arc, the material ablates and is accelerated due to the interaction of the electric field and the self-induced magnetic field. Simplicity due to solid fuel, small power demands (100-200W) and small impulse bits are some of the main characteristics. On the other hand, they show very low efficiency ( 10%) due to heat ≤ transfer losses of the fuel. [9] Larson1999 [23][24][25] In Electrostatic systems, a gaseous fuel is first ionized and then is accelerated by a strong static electric field. The exhaust gas is actually a stream of positively charged particles that needs to be neutralized to avoid spacecraft charging. This is achieved with the use of a cathode electron source which emits electrons in the exhaust beam of the thruster. Different applications of these systems have been developed and are divided mainly into Ion engines, Hall thrusters and Field Emission Electrostatic Propulsion (FEEP). In general, these systems provide high specific impulse and low thrust, commonly used in deep space missions and are suitable in employing the spiral, low-thrust trajectory strategies. [9] Larson1999 [24][25] Ion engines are based on ions acceleration between 2 grids under the effect of an electric field. Xenon gas is the most widely used fuel. The fuel is ionized by electron bombardment that are produced in a central cathode. The ionized products are filtered by an extraction grid which operates in a slight negative potential. Then they are accelerated through the second grid which is set in a large negative potential. In the exhaust beam, again, electrons are emitted to neutral- ize the plume. Major problem on these engines is the cathode erosion from the ions. However, they posses high values of specific impulse (3000-4500s) and low thrust values (10-200mN). They have been used in missions for attitude control and station-keeping operations. [9] Larson1999 [24][25] Hall Effect Thrusters (HET) implement an annular ionization where a neutral gas is ejected close to an upstream anode and electrons are ejected down- stream through a cathode. Inner and outer magnetic coils generate a rotating magnetic field that put the electrons in a rotating motion in the chamber, making them to stay longer and increasing the probability of collisions with the neutral gas and thus ionizing it. An electric field is also created between the anode and the cathode. The ions are heavier and experience a weaker acceleration by the

26 magnetic force and they accelerate primarily due to the electrostatic field with a high velocity, producing thrust. They provide lower specific impulse ( 1500) than ∼ ion engines but produce higher thrust close to one order of magnitude higher. In addition, they do not posses an acceleration grid which are components with a life limit due to erosion. [9] Larson1999 [24][25] FEEP thrusters have similar characteristics with the previous systems but they differ mainly in the way the charged particles are produced. They usually employ liquified metal fuel such as Cesium, Indium and Gallium which are used to extract charged droplets from the fluid surface. A strong electric field is set up at the tip of 2 electrodes that are closed to each other, creating a ”Taylor cone” at the surface of the fluid. This ”Taylor cone” is a result of the balance between the electrostatic pull and the surface tension of the liquid fuel and in this way the local field strength is enhanced even more at the apex. The fluid cone creates a capillary feed system for the fuel close to the tip and leads to a spray of positive charged particles. Then, these particles are accelerated through the negative acceleration grid, creating the thrust. Again, a neutralizer is used to emit electrons in the beam to avoid potential charging build-up on the spacecraft. They provide high values of specific impulse in the expense of low thrust levels due to power limitations. [24][25]

2.2.3 Advantages, Disadvantages & Applications

A summary of the main characteristics and merits of each type of propulsion is given and a discussion in the view of a lander application is followed. In the pre- ceding paragraphs, it can be seen that electric propulsion offers significantly high values on specific impulse, making them fuel-efficient systems in the expense of higher power demand that adds on propulsion system mass. However, the thrust levels that they can provide are notably lower than chemical propulsion and in a lander context this type of propulsion is not suitable. On the other hand, chemical propulsion provides high thrust and small to moderate specific impulse, considering it as a favorable option for a lander. This is clearly illustrated in Figure 14.

As it is discussed in chapter 2.2.1, the Issp allows a more accurate determination of the propulsive performance and describes those propulsion system design parameters which influence system mass in relation to delivered impulse. This is shown in Figure 15, where the ∆V performance of typical PS is compared.

27 Figure 14: Thrust versus Specific Impulse on different PS. [25]

Figure 15: ∆V performance of PS concepts. [21]

28 Finally, a summary table with advantages, disadvantages and main performance characteristics is given in Table 1.

Specific Thrust Type Impulse Advantages Disadvantages [N] [s] Simple, Low Isp & Issp, reliable, Cold Gas 0.02-10 50-80 low Thrust very low cost Simple, No controllabil- reliable, ity & restarta- Solid Mo- 28000- compact 280-300 bility, very high tor 47000 size, rel- Thrust ative low cost

Moderate Isp & Simple, Mono- Issp, Moderate 0.5-22 200-230 reliable, propellant Thrust low cost

High High cost, High Bi- Thrust, 4-500 290-320 complexity propellant High Isp &

Issp High I & 3 sp 5 10− - Low Thrust Resistojet · 130-500 Issp, Less 0.5 fuel High I & High power, 2 sp 5 10− - Arcjet · 400-1500 Issp, Less complicated 5 fuel interfaces

29 High power, High I & 6 sp 5 10− - low thrust, high PPT · 1500 I , Less 3 ssp 5 10− complexity · fuel

High power, High I & 2 sp 10− - 1500- low thrust, high HET Issp, Less 0.5 2500 complexity fuel

Very high Very High power, low 3 10− - Ion 3000 Isp & Issp, thrust, high 0.2 ≥ Less fuel complexity

Table 1: Summary table of PS characteristics. [21]

2.3 Review of past missions

In this chapter, a discussion on the past lunar missions is held and a review of the trajectory strategies and choice of propulsion system is described. In the discussion is included both lunar landers and lunar orbiters that have been flown to date. [3][4][14] The main missions during the 1960s and 70s were the Apollo and Surveyor program from US and the Luna program from Russia. Luna-9 was the first spacecraft which achieved a soft landing on Lunar surface. Apollo 11 mission was the first manned mission which placed the first people to walk on the Moon. All the missions during that period used bi-propellant propulsion systems for all the required maneuvers. The LV (Launch Vehicle) was dedicated to put the spacecraft into LTO. These missions were characterized by high complexity and high cost. Safety (especially for manned missions) and high performance were vital. The Surveyor spacecrafts used bi-propellant thursters combined with a solid motor which was ejected during the lunar descent phase. The thrusters were used for

30 orbit corrections, attitude control and descent. After these programs ended, there was a huge time gap for the next lunar mission, and especially one involving a lunar lander. In 2013, the Chinese mission Chang’e 3 made China the third nation to land on the Moon after Apollo and Luna programs. Again, the propulsion system of this mission was a bi-propellant system, using a main throttable engine assisted by a combination of different thrusters for the descent phase, the hover- ing phase, the LOI maneuver, the correction maneuvers and the attitude control. Again, as for the previous missions, the LV put the spacecraft in LTO. In table 2, a summary of the propulsion systems and the trajectory of the lander missions to date is shown. [3][4][14]

Mission Nation TLI approach On-board PS Remarks Throttable main engine Apollo pro- Direct Injection US Bi-propellant (4.7-44 kN) & gram by LV 445 N attitude thrusters (x16) Luna Direct Injection Throttable main Russia Bi-propellant Program by LV engine Vernier engines Surveyor Direct Injection Bi-propellant & US (x3) & Ejectable program by LV Solid motor solid motor Throttable main engine (1.5-7.5 kN), Direct Injection Chang’e 3 China Bi-propellant 150 N thrusters by LV (x16), 10 N attitude thrusters (x12)

Table 2: Mission characteristics of Lunar Landers.

The TLI approach has a major impact in the selection of the PS. As it will be shown in the revision of lunar orbiter missions that took place after 1990, the TLI was the most important factor in the selection of the PS. In total 9 missions of lunar orbiters have been flown up to date. Table 3 shows the basic characteristics of their missions. On the one hand, it can be seen that from the 9 missions, the 8 used chemical propulsion while only 1 mission used electric propulsion as the main propulsion system. Mono-propellant, bi-propellant or a combination of them among the chemical propulsion systems were used. On the other hand,

31 different strategies were implemented for the TLI approach. These are the direct injection, the phasing loop, the WSB and the spiral (low-thrust) approach. The most preferred approaches were the phasing loop and the direct injection where 7 out of 9 missions used these strategies. One mission used the WSB approach and the other one chose the spiral method. The spiral approach is associated with the implementation of an electric propulsion system as it is the most suitable one. Chemical propulsion is linked with the rest of the strategies where different com- binations of chemical propulsion systems with TLI approaches can be selected. [3][4][14] In Table 3, some design and performance details of the PS of the lunar orbiters are presented. In particular, the different types of PS as well as different combi- nations among them are shown. In most of these missions, chemical propulsion is implemented, where bi-propellant systems were used due to high performance and mono-propellant systems for simplicity. Some of the missions chose to combine the merits of the two systems with a penalty in an increased complexity and cost. Electric propulsion was only used in one of these missions, resulting in a very low fuel consumption. In addition, the associated TLI strategies are evident in the table too. It is worth to mention that the spiral approach is accompanied by the use of electric propulsion. Finally, it is outlined the type and the number of engines used for each mission as well as the type of propellants used, covering all the required maneuvers. [3][4][14] Other important factors that can effect the choice of the PS is the specific mission objectives, such as low-cost philosophy or observations of more than one celestial bodies. Also, the orbiter mass is an important factor while selecting a suitable PS for a mission. It can be seen from Table 3 that lightweight spacecrafts are more flexible in the selection of the PS and they prefer the mono-propellant solution due to lower cost, higher reliability and simplicity (when not considering CLEMENTINE which had specific mission objectives). On the other hand, heavier spacecrafts prefer the bi-propellant solution or a dual mode system, such as the Indian and Chinese missions, due to the high performance and the lower amount of fuel needed compared to mono-propellant. In case of electric propulsion, the significant low thrust values and the high specific impulse make them suitable for lightweight spacecrafts. However, factors such as the mission time and the target

32 lunar orbit seem to not affect the selection of the PS. [3][4][14]

Mission Nation TLI approach On-board PS Remarks Hiten Japan WSB Mono- Hydrazine propellant fuel, 23N thrusters (x8), 3N thrusters (x4) CLEMENTINE US Phasing loop Solid motor Solid motor, & Dual mode 490N en- (mono- & gine (x1) with bi-propellant) MMH/MON propellants, 22N thrusters (x2), 4N thrusters (x10) Lunar Prospector US Direct injection Mono- Hydrazine fuel, propellant 22N thrusters (x6), solid motor for TLI SMART-1 Europe Spiral Electric propul- 68mN Hall sion effect thruster (x1) SELENE Japan Phasing loop Dual mode 500N en- (mono- & gine (x1) with

bi-propellant) N2H4/MON-3, 20N thrusters (x12), 1N thrusters (x8)

33 Chang’e 1 China Phasing loop Bi-propellant MMH/MON propellants, 490N engine (x1), 10N thrusters (x12) Chandrayaan-1 India Phasing loop Bi-propellant MMH/MON propellants, 440N engine (x1), 22N thrusters (x8) LRO US Direct injection Mono- Hydrazine fuel, propellant 88N thrusters (2x2), 20N thrusters (2x4) Chang’e 2 China Direct injection Bi-propellant MMH/MON propellants, 490N engine (x1), 10N thrusters (x12)

Table 3: Mission characteristics of Lunar Orbiters.

34 3 CMML Case

3.1 Approach and Method

The proposed Commercial Micro Moon Lander (CMML) case is described in this chapter. The key points of the CMML study are highlighted below:

• Mission-Trajectory analysis

• Launch scenario

• Mission main drivers

A basic trajectory analysis is putting forward, giving the preliminary ∆V budgets of the mission using MatLab and GMAT software. The description of the trajectory, where each maneuver taking place is mentioned and the comparison between 2 different trajectory approaches is included. Then, the propulsion system design based on the CMML case is analyzed. First, the identification of top-level requirements is established. Then, a requirement analysis using weighting factors is done where a dualistic comparison to each requirement with the other ones is used. With this method, a basic comparison of the requirements is achieved and a hierarchy of the requirements is produced which drives the propulsion design. The comparison is summarized in a matrix where the final weighting factors are shown as percentages. In addition, a matrix with arguments on each requirement comparison can be found in Appendix. Then, the designing philosophy of the PS is described where the margin philosophy and the step-by-step procedure of PS dimensioning are included. One important aspect of the mission is the Entry, Descent and Landing (EDL) phase where the approach and the breakdown of the maneuver sequence are described and the average thrust required is calculated for this case. After setting-up the procedure for the PS calculation, a trade-off analysis is taking place in order to have a clear picture on the different options for the mission. Combining the trade-off analysis with the technology available and other important factors that needs to be considered and were derived from the mission particularities. Basic analytical equations are used to derive results.

35 The results of the trade space analysis are presented and discussed as a following step. This leads to the selection of a suitable PS for the mission and the main characteristics of the design are described. A top-level architecture lay-out with some of the key performance indicators (KPI) are shown to complete the presen- tation of the final selection, providing the proper argumentation.

Mission-Trajectory Analysis

The main objectives of the mission, as it has already been stated, are the establishment of a delivery platform that can put payloads on the Moon surface and create a high frequency launch rate. As a consequence, the mission strategy requires to select an orbit that can offer rockets going regularly there and in the same time the lander design shall be feasible to go to the Moon from that orbit. The option of GTO as the parking orbit can fulfill these objectives. A plethora of communication satellites use GTO as their parking orbit and they use their on-board PS to go to their final orbit which is the GEO orbit. Thus, there are a lot of launchers that go frequently in GTO, for instance Falcon 9 and Ariane 5 rockets. This offers a lot of ride-shares opportunities as a secondary payload. On the other hand, GTO could offer a feasible lander design because of the lower ∆V demand and thus lower fuel demand compared to LEO orbit, since LEO orbits can offer the high frequency requirement too. Having said that, the mission strategy can be described as following: After de- taching from the launching vehicle in GTO, the lander will start to use its own PS for all the maneuvers until touchdown of Moon surface. In GTO orbit, the lander will make its first burn in perigee to raise the apogee, going to an LTO orbit. The apogee of the LTO elliptic orbit will be equal to the distance between Earth and Moon. While coasting in LTO, the lander will make another burn to correct its orbit and to have the proper arrival condition on Moon. In particular, this maneuver, which is called TCM (Trajectory Correction Maneuver), changes the inclination of the LTO in order to meet the Moon at the right place and time. Also, it should be noticed that the maneuver starts before reaching the Moon’s SOI (Sphere of In- fluence) in order to avoid the effect of Moon’s gravity. This strategy has been used and described in previous Moon missions. While in the proximity of the Moon, the lander has to lower its velocity in order to be captured from the Moon’s gravity

36 and match its velocity. This will be achieved with the LOI maneuver where the PS needs to make a big burn in the opposite direction to slow down the lander and enter a LLO (Low Lunar Orbit) around the Moon. After orbiting the Moon, the lander will make the last major maneuver which is the landing maneuver. This maneuver consists of different steps and they will be described later where the EDL strategy will be analyzed. To sum up, the trajectory phases and the ∆V budget are shown in Table 4.A simulation of the trajectory strategy using GMAT software was derived, giving a graphical representation of the mission and the ∆V required as the outputs. It is worth to mention that the calculations on GMAT are based, among others, on the vis-viva equation, orbital parameters, launch dates, etc. GMAT is a user-friendly software, which is open-source and designed by NASA, for trajectory design where the user is able to propagate the spacecraft and extract the ∆V values. The demonstration of the trajectory design on GMAT is not presented in this document since it is out of the scope and a plethora of tutorials are available online. The simulation is built based on [28], where a script of the software as well as the simulation logic can be found. In Figure 16, the graphic output of the simulation is shown, where in red color shows the ”Leaving the Earth” phase and in light blue color the application of the TCM with relative inclination change. The grey line indicates the Moon orbit around the Earth. Graphical artefacts of the trajectory are due to the application of both forward-propagation and a back-propagation (details can be found on [28]) used to compute the TCM maneuver at the border of the Moon SOI. The ”Arrival to the Moon” phase is shown in Figure 17. It can be noticed that in Table 4 the last row includes a margin in the total ∆V budget. This is due to the ESA (European Space Agency) margin policy implementation that is used while in the preliminary design of a mission and it will be discussed later in the PS design. It can be seen from the results that the total ∆V (excluding the margin) is fairly close to the top-level requirement established at the beginning.

37 Maneuver ∆V [m/s] TLI 676 TCM 137 LOI 831 Landing 1726 Total ∆V 3370 Total ∆V with 5% margin 3538

Table 4: ∆V Budget of Hohmann transfer.

Figure 16: GMAT simulation of Hohmann trajectory to the Moon. With red color is the transfer trajectory and in light blue is the TCM maneuver. In grey color is the Moon orbit around Earth.

A second trajectory strategy that could be implemented as a back-up solution is the implementation of the bi-elliptic trajectory. The graphical representation of this transfer is shown in Figure 18. As described earlier, the bi-elliptic trajectory is a 3-burn trajectory and can lead to a reduction of the total ∆V budget. This is shown in Table 5. The ∆V budget and the graphical representation of this type of trajectory was calculated by analytical calculations (through the vis-viva equation)

38 Figure 17: GMAT simulation of Moon arrival. With red color is the arriving circular lunar orbit and the landing phase. In light blue is the TCM maneuver while arriving on the Moon. using MatLab. The main equations can be found in Appendix A. Also, in this case the total ∆V (excluding the margin) is close enough to the top-level requirement.

Maneuver ∆V [m/s] Transfer maneuver 1 756.20 Transfer maneuver 2 354.53 LOI 226.92 Landing 1726 Total ∆V 3063.65 Total ∆V with 5% margin 3216.83

Table 5: ∆V Budget of Bi-elliptic trajectory.

The main differences of these 2 trajectories are the ∆V budget and the mission time. Using a bi-elliptic trajectory the savings in ∆V is around 322 m/s which is translating in fuel savings of about 42 kg. However, the mission time between these 2 trajectories differ significantly. The Hohmann transfer will take around

39 Figure 18: Bi-elliptic trajectory.

5 days to reach Moon while the bi-elliptic trajectory will take around 63 days. It is worth to mention that the bi-elliptic transfer needs less total propulsive energy than a direct transfer when the ratio between the radius of final and initial orbits is larger than 12 [14]. From a scientific point of view, the bi-elliptic choice could be the case but for a commercial mission a fast transfer could attract more customers which allow to have a platform that can provide a fast delivery service with high repeatability factor. Radiation which is related with mission time can also act as a drawback on selecting a bi-elliptic trajectory. On the other hand, implementation of the bi-elliptic transfer makes the mission to be less dependent on appropriate launch windows. Since the case of this study is a commercial mission, the Hohmann transfer is the selected trajectory with the bi-elliptic transfer being able to act as a back-up trajectory strategy.

40 Launch scenario

The launch scenario is one of the most important parameters for a mission. It affects the trajectory strategy and the cost of the mission since it takes a huge part of the total cost. From a commercial point of view, sharing the launch vehicle and take advantage of a launcher with a high launch rate can reduce the mission cost, generate more income and create a niche compared to the other competitors. For this reasons, a piggy-back rideshare and a high launch rate is implemented. There are 2 European launchers, the Ariane 5 and the Arianespace Soyuz which they use the ASAP (Ariane Structure for Auxiliary Payload) interface for secondary payloads. The mission focuses on a micro-lander and thus these 2 launchers are the most suitable in this concept where the launch cost as secondary payload using the ASAP is around $1.200.000 [29] and the launch rate could be up to 6 launches per year. More details and requirements on the ASAP configuration can be found in [10]. Some of the most important characteristics of the ASAP configuration are:

• Mass limit of 300 kg for Ariane 5 and 400 kg for Arianespace Soyuz

• Volume limit of 1.5 m in diameter and 1.5 m in height

• Strict deviation limits of the (Center of Gravity)

• Need of compatibility with the ASAP adapter structure

• Resistance in high vibration loads

From the preliminary mass budget of the whole spacecraft, the 300 kg was deemed unfeasible and the Ariane 5 launcher option was discarded. This left the Arianespace Soyuz option which will set-up some of the requirements of the PS designing. Summing up, the selected launch scenario consists of:

• Launch as a secondary payload

• Provide a high launch rate

• Use of the Arianespace Soyuz launcher

• Launch from Kourou, New Guinea, the launch site of Arianespace

• Launch in GTO as the parking orbit

41 Mission main drivers

The present study is driven by the commercial character of a mission. The commercial concept is a very complex context and normally requires an extensive analysis, something that is out of the scope of the present thesis. Nevertheless, it is worth to mention the main drivers of this particular mission that affect the overall design of the PS. In contrast to scientific and public missions, commercial missions requires to create an income apart from providing the mission deliver- ables, which makes this factor as one of the most important. In addition, low cost on private missions is a main driver because it can distinguish a viable from a non-viable mission. COTS (Commercial Off-The-Shelf) components can cause a reduction on the cost and also in the development time which is an important driver too. Another main driver is flexibility due to the fact that different customers have different requirements and requests. This can change the subsystem design of the lander which drives the design to adapt a more universal configuration to cover as more cases as possible. Finally, the first mission on these kind of studies acts as a demonstration technology mission which will put the spacecraft as a reliable platform for the selected mission and will attract potential customers.

3.2 Designing Philosophy of the PS

In this section, all the parameters that have been considered for the designing of the PS are discussed. Having described the launch scenario and the trajectory strategy, the requirements of the PS for the mission can be derived. In Table 7, the requirement matrix of the PS is presented. At this stage of the project, only the top-level requirements were identified where initial estimations and assumptions were made based on previous designs and on mission requirements. As the project progresses, the breakdown of the top-level requirements into specific and more detailed requirements will be carried out. A comparison of the requirements is used and a rough ranking is implemented in order to have a better picture which requirements will be the major drivers when selecting a suitable PS for the lander. A dual comparison (or pairwise method), meaning that a comparison of each requirement with all the other requirements, is used to extract weighting factors which will show the importance of each require-

42 ment. It should be noticed that all the requirements are important and needs to be fulfilled but the weighting factors will assist the selection of the PS configuration. The weighting scores of the dual comparison are shown in Table 6.

Definition Score Less important 0 Equally important 0.5 More important 1

Table 6: Weight factors description

The requirement matrix of the PS is shown in Figure 7. This matrix contains some of the main functions that need to be delivered for this mission. In Figure 19, the comparison of each requirement is deducted and the weighting score is shown for each of them. In particular, in each row a requirement is compared with all the other and a total score is summed up. Then, every score of each requirement is summed up to a total score which is shown in the last row with name ”Sum”. The weighting factor for every requirement is the ratio between its individual score divided by the ”Sum”. The result shows that the main drivers for the designing of the PS is the thrust and ∆V required for the mission as well as the ASAP compatibility, mass and volume restrictions. The argumentation for the individual score of each requirement can be found in Appendix C. These results act as baseline for the PS design and the high score requirements affect any design decision in the PS selection and sizing process. Hence, the approach to the choice of the PS is based on the thrust and ∆V requirements as well as on the ASAP compatibility with the mass and volume restrictions.

43 ID Requirement PS_R-01 The PS shall respect the ASAP compatibility. PS_R-02 The PS shall deliver a soft landing. PS_R-03 The PS shall not compromise the limitations in terms of mass and volume. PS_R-04 The fuel of the PS shall be able to be stored. PS_R-05 The PS shall deliver the required ∆V for each maneuver. PS_R-06 The PS shall deliver the required thrust for each maneuver . PS_R-07 The PS shall have a high performance. PS_R-08 The PS shall be able of performing multiple ignitions during the mission. The PS design shall be able to take advantage of off-the-shelf components PS_R-09 use. PS_R-10 The PS shall be able to have thrust controllability. PS_R-11 The PS shall contain suitable fuel. The PS shall be able to deliver a constant operation pressure on the thruster PS_R-12 chambers and maintain constant pressure in the circuit. The PS shall be able to maintain a constant temperature on the PS_R-13 tanks, pipes and nozzle.

Table 7: PS requirement matrix.

Figure 19: Requirements weighting.

The sizing methodology of the propulsion system is discussed next. While designing a particular mission there is the need of dimensioning each subsystem following a common basis. In this study, it was agreed to follow the ESA margin

44 philosophy [30], where the following margins were considered:

• 5 % on the total ∆V

• 2 % on the required total fuel as residual

• 10 % extra volume for fuel and pressurant tanks

• Equipment margin on each component

• 10 % ADCS fuel ([31])

In table 8, the mass budget of the lander is shown. It is summarized the mass of each subsystem of the lander including the equipment level margins. These margins are described in the ESA margin policy [30] which is the policy that is followed in the project. In the nominal dry mass of each S/S except the Propul- sion S/S is applied an extra 20% system margin based again on the ESA margin policy. After the calculation of the dry mass of the Propulsion S/S as it will be presented in the following, the total dry mass and the total wet mass of the lander can be calculated. The mass budget of the lander is a preliminary estimation and the mass of each S/S is estimated based on previous studies (e.g. [32]) and preliminary calculations from each S/S engineer of the team. This acts as an input for the procedure followed in the estimation of the Propulsion S/S mass which is presented next.

45 Subsystem Mass [kg] ADCS 9.72 Communication 5.58 GNC 1.63 Mechanism 9.40 OBDH 0.77 Power 20.80 Structure 21.77 Thermal Control System 2.43 Payload 10 Harness 4.10 Nominal Dry Mass without Propulsion 86.20 Total Dry Mass without Propulsion (20% System Margin) 103.44 Propulsion 24.14 Total Dry Mass 127.58

Table 8: Lander mass budget (Including equipment level margins of each subsystem).

The ADCS fuel margin is a preliminary estimation based on the New SMAD (Space Mission Analysis and Design) book [31]. Further investigation of the required fuel for keeping the correct attitude of the spacecraft is needed for more precise estimation which will be carried out in the next phases of the project. The sizing of the propulsion system is based on the total wet mass of the spacecraft. However, the propulsion system is part of the dry mass of the spacecraft and the dry mass is required to calculate the fuel needed. This leads to an iterative process. Thus, the calculation steps of the propulsion system are described below:

1. Calculate the total dry mass of the spacecraft including the margins on subsystem level and the total margin on system level, without considering the propulsion dry mass which is unknown at this step.

2. Calculate the wet mass of the spacecraft using the rocket equation, where

the ΔV and Isp are given.

46 3. Calculate the required fuel from the wet and dry mass calculated above.

4. Based on the amount of fuel, calculate the mass of fuel and pressurant tanks and the amount of pressurant.

5. Select a suitable engine (preferably off-the-shelf component) and add its mass.

6. Add the masses of the tanks, pressurant and engine and calculate the dry mass of the propulsion system.

7. Add the dry mass of the propulsion system calculated to the dry mass of the spacecraft and start iterating.

8. Go to step 2 and follow the steps until step 7 until convergence.

9. After convergence, add equipment margins for each component on the calculated propulsion system and sum up the new dry mass of the propulsion system including the margins.

10. Follow step 8 until reaching a second convergence.

The equations that describes the calculations needed to follow the above steps can be summarized below [21]:

∆V m = m [1 e ve ] (6) p 0 · − m V = p (7) p ρ

Vp Vp C = Vt = (8) Vt ⇒ C P m = m op (9) T p · C ρ K · · p P (1 C) M m = m op · − · (10) pr1 p · C ρ z R T · · p · · 1.1 P M m = m · op · (11) pr2 p · ρ z R T · p · · 1.1 P z m = m · op · pr (12) Tpr p · ρ z K · p · pr

47 m =0.1 m (13) ADCSfuel · p m =0.02 m (14) residual · p P V m = c · case (15) case K

where mp is the fuel mass in kg, m0 is the initial (wet) mass of the lander in 3 kg, ve is the exhaust velocity, Vp is the volume of the fuel in m , ρ is the average 3 density of the fuel and oxidizer in kg/m , C is the tank filling ratio, mt is the mass of the tank in kg, mpr1 is the mass of pressurant in the fuel tank in kg, mpr2 is the of pressurant in the pressurant tank in kg, mTpr is the mass of the pressurant tank in kg, M is the molecular mass in kg/kmol, K is the tank performance factor 2 2 in m /s ,Pop is the operating pressure, R is the gas constant in kJ/K/kmol, zp is the gas compresibility factor in fuel tank, zpr is the gas compresibility factor in pressurant tank, mADCS is the mass of the fuel in kg for ADCS maneuvers which is 10% of the total fuel, mresidual is the mass of residual fuel in kg which is 2% of the total fuel, mcase is the mass of the solid motor case in kg, Vcase is the volume of 3 the solid motor in m and Pc is the motor chamber pressure in Pa. The procedure presented above is discussed next. The driving factor of the sizing of the PS as it can be seen in the equations is the propellant mass. For the lander design a bottom-up approach has been chosen among others. A bottom- up procedure is based on the estimation of the mass of all the S/Ss of the lander, summing up to extract the dry mass and then calculating the wet mass and the propellant required. This is the reason of the iterative procedure employed for the calculation of the PS due to the fact that the PS is based on the required propellant mass. By knowing the dry mass of each S/S except the PS, it is possible to calculate the wet mass and the propellant but this will not be the correct values since PS is missing. However, it acts as the first iteration which gives the possi- bility to evaluate the PS and then adding it to the initial dry mass. By continuing iterating this procedure, there will be a point where the values of the dry mass (including the PS), the wet mass and the propellant mass will result in almost the same value on the PS calculated in the previous iteration. This is the point where the procedure is finalized and the calculation of all required variables are accomplished. Due to the usage of the ESA margin policy, the equipment level margins are applied to the PS as well, resulting to the second iteration ”round” to achieve

48 the second convergence as mentioned in the procedure steps above. The main im- plications of this procedure are two. The first is that to size the PS there is always the need of knowing the dry mass of the spacecraft where in the early phases of a mission is hard to estimate. This leads to the second implication which is the accuracy of the method. Since the knowledge of the dry mass is hard to estimate and usually is based on top-level assumptions and simple calculations, it will also result to a not completely accurate calculation of the PS. In addition to this, the iterative procedure is always linked to errors which in this case are dependent on the amount of iterations used. In the case of the PS, the convergence stops to the second significant decimal point which for the purpose of a preliminary calculation is sufficient. All in all, the accuracy of the procedure is deemed to fit at this early phase of the mission as it is a usual practice for all Phase 0 mission studies. Alternative approaches, such as a top-down approach (which is the reversed approach implemented here) or others, could be studied and implemented on the next stages which could give a good trade-off. COTS components are used for the thrusters and their mass is fixed. The mass of the main engine is 3.6 kg [33] and the mass of each of the thrusters is 0.45 kg [34] regarding the Bi-propellant PS. The C factor takes 0.95 for bi-propellant systems, 0.75 for mono-propellant systems and 0.90 for solid systems. It should be noticed that the ∆V of low thrust from GTO to LLO is taken from the SMART-1 mission [35] and is equal to 3.9 km/s. Next, a trade-off analysis is carried out which will help on the decision of the PS suitable for the mission. The present study is focused on the state-of-art space propulsion systems that typically used on previous and current missions, which are more suitable for a commercial mission due to lower development time (high TRL - Technology Readiness Level) and enabling the choice of available “off-the-shelf” components and thus reducing the cost. With this in mind, the main types of propulsion under investigation are 3:

• Solid propulsion

• Liquid propulsion

• Electric propulsion

A trade-off study on different configurations was made in order to scope the feasibility concept. Two main different scenarios were considered: the first sce-

49 nario consists of using only one type of chemical PS for the mission while the second scenario considers a hybrid PS, meaning to combine 2 different PSs. For the second scenario, 2 cases of hybrid systems are considered. The first case is the implementation of electric propulsion from GTO to LLO and chemical propulsion for the landing phase. The second case is the implementation of 2 different chemical PSs, where in the scenarios of solid-mono-propellant and bi-propellant- mono-propellant configurations the mono-propellant system is used for the landing phase and the other stage is used from GTO to LLO. In the solid-bi-propellant configuration, the solid stage is used from GTO to LTO and the bi-propellant stage is used for the rest of the maneuvers. On Tables 9 and 10, the required fuel during each maneuver for different propulsion configurations is presented. It should be mentioned that the Isp used for each type of PS in order to calculate the fuel required based on the ∆Vs and the dry mass of the lander (where the dry mass of the lander is different for each type or combination of PSs, however is derived based on the same procedure described above) are:

• bi-propellant = 318 sec

• mono-propellant = 220 sec

• solid = 290 sec

• electric = 1640 sec

Bi- Solid/ Bi- Mono- Solid/ Bi- propellant/ Maneuver Solid Mono- propellant propellant propellant Mono- propellant propellant TLI 80.85 386.51 105.51 138.63 88.00 118.27 TCM 14.26 63.99 18.35 24.12 13.50 20.86 LOI 73.71 308.66 93.46 122.80 69.77 107.82 Landing 100.59 351.77 122.62 176.97 95.21 176.97

Table 9: Required fuel in [kg] for every maneuver and for different chemical PS configurations.

50 Electric/ Electric/ Electric/ Maneuver Mono- Bi- Solid propellant propellant Low-thrust 97.94 66.23 69.84 Landing 176.97 89.94 101.92

Table 10: Required fuel in [kg] for every maneuver and for different electric- chemical PS configurations.

In Figures 20, 21 and 22 show the results on dry mass, fuel and total mass for different propulsion configurations. Configurations that includes electric propulsion are favorable in terms of mass compared to the others. Bi-propellant system follows in terms of mass, giving the next smallest value. The highest values on mass are from mono-propellant system which requires lot of fuel due to small efficiency.

Figure 20: Dry mass in [kg] of the lander on Moon surface for different PS configurations.

Although electric propulsion offers significant fuel savings, other parameters are necessary to be considered when selecting a suitable PS for the mission under investigation. Such important factors are:

• Mission time

51 Figure 21: Total fuel in [kg] required for the mission for different PS configurations.

Figure 22: Total mass in [kg] of the lander for different PS configurations.

• Power demand

• Radiation

• Complexity (System and Trajectory)

These factors need to be considered in a commercial frame. In this context, mission time can be increased to 1 year with the implementation of electric propulsion and increase the radiation dose on the lander significantly. High radiation

52 will increase the mass of the spacecraft due to shielding. The implementation of chemical propulsion by means of Hohmann or Phasing loop trajectory can make the mission time to last around a week. Electric PS requires high power and for a Moon mission can be around 1kW which makes again the lander heavier due to bigger solar panels and the need of extra bulky components for energy conversion. This can cause volume issues too. Electric PS is a more complex system compared to chemical PSs which increase the complexity. However, the complexity becomes a lot higher due to the complex nature of low-thrust maneuvers which require a more sophisticated trajectory. Important to note that the lack of the landing capability of the electric PSs make these systems to be always combined with a chemical PS for the landing phase which put higher complexity in the design. In Table 11, the 4 best cases in terms of mass are compared taking into account other important mission factors.

Characteristic/ Electric/ Bi- Electric/ Solid/ Bi- Configura- Bi- propellant Solid propellant tion propellant Moderate Minimum Moderate Mass Low mass mass mass mass Power Low High High Low Mission time Low High High Low Radiation Low High High Low Complexity Lower High High Higher

Table 11: Comparison of different configurations considering other important factors.

All these parameters and factors that are considered lead to the selection of a suitable PS. Based on the above analysis, the PS that covers in the best way more of the factors is the Bi-propellant PS. Next chapter is describing the design of the selected PS. The wet mass of the lander using this propulsion configuration is 429.98 kg including all the margins. This value is above the ASAP mass limit. A better analysis of the mass budget and sharpening the margins could potentially reach the wet mass close to 400 kg. In order to scope the feasibility of the lander with

53 this type of PS, a trade-off analysis between the payload capability of the lander and the wet mass is carried out. In particular, the payload mass is varied by 1 kg from 1 to 10 kg and the dry and wet mass of the lander is calculated based on the procedure described earlier for each case. In Figure 23, the total mass of the spacecraft as a function of the payload mass is shown. It can be seen that for a payload in the range of 1-4 kg, the total mass is under the 400 kg limit of the ASAP configuration. In the range of 5-10 kg this threshold is violated. Also, it can be noticed that for every 1 kg of payload the wet mass of the lander is increased by around 5 kg. A further discussion on the results of the payload capability with respect to the lander wet mass follows. The initial payload mass selected as a baseline for the mission was 10 kg. In this range of payload mass, candidate payloads are small rovers, scientific instruments such as spectrometers and cameras, communication technologies and software payloads (e.g. for entertainment or scientific purposes). These payloads can weigh from a few kilos to tens of kilos. Thus, a single payload of about 10 kg or a combination of more than one lighter payloads could be possible to be accommodated on the lander. However, the results show that a lander with a payload capacity of over 5 kg exceeds the limits of the selected LV and launch scenario as a piggy-back. The payload capacity of the lander that can reduce the mass budget below the limit ranges between 1 and 4 kg. This capacity is rather small but still some of the proposed payloads could be possible to fit. For instance, cameras could be fitted in this range of mass, recording and mapping the surface of the Moon. In combination with spectrometers, which can also fit the mass range, can produce valuable scientific data for agencies and related corporations in terms of mapping and analyzing the soil of the lunar surface. An emerging technology in the space sector is the design of small rovers (standard- ized in the same way as CubeSats) which could assist the exploration of the Moon. Their lightweight design makes them suitable to this mass range too. In addition to this, small rovers could be used for entertainment purposes in a commercial mission. Finally, software payloads are an interesting candidate since they are weightless and they could generate profit via their services either for scientific or entertainment purposes. In this sense, a lander with a wet mass below 400 kg and a payload capacity 1-4 kg could be designed. However, the concern address-

54 ing this lander design is not the feasibility of the mission itself but the viability of the mission due to its commercial aspect. A commercial lander design shall generate a revenue that both covers the cost of producing the lander and makes a profit. Hence, a better consideration of the commercial aspects of the mission will give a better picture whether the realization of a mission is reasonable or not.

Figure 23: Lander’s wet mass for different payload mass.

As mentioned earlier, the requirements that drive the design of the PS are the thrust, ∆V, mass, volume and ASAP compatibility. Based on the analysis carried out above, a bi-propellant system is selected, driven by the aforementioned requirements. The thrust requirement drives the design due to the landing phase of the mission and the soft landing requirement which force the chosen PS to provide specific thrust levels. Hence, chemical propulsion with high thrust are more suitable than electric propulsion. The ∆V requirement drives the selection of the type of PS where high efficiency (e.g. bi-propellant over mono-propellant systems) is preferred in order to maintain the limits in mass and volume. In combination with the ASAP compatibility, mass and volume requirements make the selection of the PS to be towards to a monolithic system (to keep complexity low) with high performance, minimizing the fuel required to take advantage of the launch scenario chosen. This summarizes the choice of the PS based on the main drivers resulted by the ranking.

55 3.3 Preliminary selection of suitable PS

Based on the previous analysis, the propulsion system design is focused in utilizing a unified bi-propellant (MMH/MON) system. This system is taking advantage of the use of a common feeding system and sharing the same tanks between the main engine and the ADCS thrusters. Also, using one type of system keeps the complexity low. It consists of:

• 1 main engine of 420 N

• 16 thrusters of 22 N

• 4 spherical fuel tanks

• 1 spherical pressurant tank

A general rule is that bi-propellant systems can be heavier than mono-propellant systems at a specific range of impulses. At low impulses,below about 10,000 Ib- s, the system weight of the bi-propellant system due to additional hardware required outweighs the high efficiency gained with high specific impulse. Below total impulse levels of about 100,000 Ib-s, mono-propellant should be considered whereas on high impulses levels the bi-propellant system is the most suitable candidate [36]. The main phases required thrust maneuvers are the following:

• GTO to LTO

• LTO to LLO

• LLO to Moon surface

The 420 N engine acts as the main actuator for the spacecraft during the mission for all the maneuvers. Eight of the 22 N thrusters are facing downwards in the same direction as the main engine and act as assist to the main engine, especially during the landing where the main engine and the eight 22 N thrusters are fired to achieve the soft landing. The other eight 22 N thrusters act as ADCS thrusters, correcting the attitude of the spacecraft and also assisting the main engine while thrusting to maintain the proper thrust vector. The ADCS thrusters are

56 in a canted (tilted) position placed on each side of the lander and firing a combination of them can fix the attitude of the spacecraft. In Figure 24, a first representation of the thrusters configuration is presented. It can be seen that the main engine is placed centrally, surrounding by the 8 assist thrusters. Thus, in this configuration all the thrusters are engaged during the landing phase, providing the required thrust. For the rest of the maneuvers during transfer, only the main engine is used. In addition, on the right it is shown a representation of the 3-axes control of the lander with the proposed ADCS thrusters configuration. As already mentioned, the ADCS thrusters are placed on each side of the lander in a tilted position as it is shown in the graph. The thrusters are in 4 groups of 2 configuration. In the figure only 2 groups of thrusters are shown where the other 2 are located on opossite sides. The red arrows show the direction of the thrust and thus the control of each axis. The 3-axes control is possible with this configuration of the thrusters, where different firing combination will achieve the desired control. Also, it is needed to mention that the graph is not by scale and acts only as a description of the thrusters configuration concept.

Figure 24: Proposed engine configuration on the lander. Left: Bottom view of the lander. Main engine and assist thrusters are shown with red font. The ADCS thrusters are shown as black boxes from this view. Right: Isometric view of the lander. The ADCS thrusters configuration is shown with a representation of control on all 3-axes. The red arrows represent the thrust direction during firing with results in a control of an axis.

Two main attitude requirements are needed:

57 • Solar panels facing the Sun

• Antenna facing the Earth

The fuel tanks and the pressurant tank are made of titanium. Also, the tanks contain surface tension propellant management devices. As pressurant, Helium is used. The tanks are spherical since they give optimal accommodation in terms of volume and can fit the volume envelope of the ASAP. The ADCS thrusters, during landing phase, has a major role on adjusting the thrust vector of the main engine to the proper direction. Furthermore, the lander implements a 4-tank configuration in the same plane (horizontally). The 4-tank configuration employs a simultane- ous fuel depletion from all 4 tanks in order to maintain the CoG. The tanks are accommodated in a symmetrical way where the 2 fuel tanks are opposite to each other and the same applies for the 2 oxidizer tanks. This is the most used tank configuration among previous lunar lander missions (e.g. Apollo, Chang’e, SpaceIL) and future lunar landers (e.g. Astrobotics, PTScientints). However, active control is needed due to potential small deviations of the fuel amount depleted during burning. There are two modes of operation for the tanks: the blow-down mode and the pressure constant mode. The blow-down mode is usually used in mono-propellant configurations and fuel and pressurant are stored in the same tank separating them by a diaphragm. In this configuration the pressure in the tanks and thus the feeding pressure in the thrusters is not constant, leading to a not constant thrust. In the pressure constant mode the feeding pressure on the thrusters remains constant, and thus the thrust, by storing the pressurant in a high pressure external tank to the fuel tank. This keeps the thrust constant meaning that the performance of the system remains constant. Since the mission requires a high performance system and there are mass limitations, the selection of a constant pressure mode is desirable. A blow-down mode would require more fuel due to the decrease in performance. Considering the previous analysis that is done, the major identifiers needed for designing the PS are:

• Mass of all subsystems

• Launch scenario

58 • Trajectory strategy

• Type of maneuvers (finite burns)

The dimensioning procedure of the PS system is based on the mass of the other subsystems where changes in their mass affect directly the mass of the PS. The launch scenario affects the designing choice of the PS where different scenarios will give different ∆V budgets. The trajectory strategy can drive the PS design since different trajectories will give different option in the PS type and size. Fi- nally, type of maneuvers can influence the choice of a suitable PS where, for instance, low-thrust maneuver requires electric PS while chemical PS utilizes finite burns, and not impulsive, which take into account the associated losses. A functional block diagram is shown in Figure 25. In color grey boxes are the fuel and pressurant tanks. In the color blue boxes on the bottom are the main engine, the 8 thrusters and the 8 ADCS thrusters. In color yellow boxes are the pressure control unit which contains all the pressure sensors, the thermal control unit which contains all the heaters and thermal sensors and the electrical control unit which contains the electrical valves to operate the thrusters and tanks as well as all the cables connecting all the related components. The thin lines are all the pipes that connects each component of the PS to each other. In color green box is the OBC (On Board Computer) which controls and gives all the commands to operate the PS. In the color blue boxes on the right side of the graph are all the commands sequence that act as the input for the OBC depending on the mission phase. Mainly there are 3 different sequences which are: thrusted maneuvers sequence which contains all the commands required for changing orbits, the ADCS sequence which contains all the commands for the attitude correction of the lander and the landing sequence which contains all the commands required to execute during the landing phase.

59 Figure 25: Functional block diagram of the selected PS.

A list of a preliminary component list is presented in Table 12. The block diagram presented is a preliminary architecture of the system that is envisaged. How- ever, a more detailed architecture diagram including all the components will be produced in the next phases of the project.

60 Component Pressurant tank Fuel tanks Thrusters Pressure transducers Thermal sensors Fill valves Filters Pressure regulators Pyro-valves Thruster valves (solenoid)

Table 12: List of preliminary components of the selected PS.

The Key Performance Indicators of the selected PS components and their values are presented in Table 13.

KPI Value Mass See Table 14 Volume See Table 15 Power 35 W Thrust 420 N

Isp 318 sec

Issp 2858.5 Ns/kg ∆V 3538 m/s

Table 13: KPI of Propulsion subsystem. Thrust, power and Isp values are extracted from [37]

61 Parameter Value Fuel mass 302.38 kg PS dry mass 22.14 kg Total tank mass 12.89 kg Total tank volume 0.302 m3 Pressurant tank mass 4.92 kg Pressurant tank volume 0.0302 m3 Pressurant mass 0.77 kg Main engine mass 3.6 kg

Table 14: Mass and Volume of the PS. Main engine mass is extracted from [37]. The rest of the values are based on the equations from chapter 3.2. The pressurant tank volume is estimated 10% of the propellant tank volume based on [21].

Main Pressurant Fuel Pressurant KPI Fuel tank engine tank (MMH/MON) (Helium) Mass [kg] 3.6 3.22 4.92 302.38 0.77 Length: 0.503m Height: Diameter: Diameter: Volume 0.260 m3 0.024 m3 0.248m 0.84 m 0.38 m Width: 0.248m Mass: Mass: Equation Equation on chapter on chapter Equation 3.2, Di- 3.2, Di- Equations on Reference [37] on chapter ameter: ameter: chapter 3.2 3.2 Volume of Volume of a sphere a sphere equation equation

Table 15: Mass and Volume for each component of the PS.

62 ADCS & GNC interfaces

At this preliminary stage, the ADCS & GNC subsystems are sharing the same equipment due to coupled functions. The ADCS systems is needed to keep the right orientation of the spacecraft in space, for instance keeping the antenna pointing the Earth and solar panels pointing the Sun. The GNC system is ensuring that the spacecraft follows the right path. The ADCS system is required to provide the appropriate orientation in the following phases:

• The detumbling maneuver. After the separation of the LV, the ADCS is needed to stop the tumbling motion of the spacecraft.

• During all the thrusted maneuvers. During the burns of the main engine, the ADCS actuators are activated to assist the engine to keep the thrust vector in the correct orientation during all maneuvers. In this way, losses from the non-impulse nature of the engine’s burns or from a serpentine motion are avoided (or at least minimized). Dur- ing the EDL phase, the attitude thrusters adjust the thrust vector in the appropriate direction during the different stages of the EDL phase as it will be described later.

• Keeping the correct orientation of the antenna and the solar panels. The antenna shall point to Earth and the solar panels shall point to Sun.

The GNC system is required to give the position of the lander while in the proximity of the Moon. During the LTO and LLO phases, the position of the lander can be assessed by a Deep Space Network (DSN), using the appropriate group of GSs. On the other hand, the lander shall evaluate its position mainly during the EDL phase which is mission critical. A preliminary soft landing strategy is described below. The GNC system and its architecture combined with the ADCS system and also the propulsion system, is designed in order to realize this strategy. The following set of GNC components are chosen based on [38]:

• Inertial Measurement Unit - IMU (x1)

• Star tracker (x2)

63 • Sun sensors (x2)

• Radar altimeter (x1)

From the preliminary choice of the hardware for ADCS & GNC subsystems, the main interfaces are schematically described in Figure 26.

Figure 26: Preliminary architecture layout of the ADCS & GNC subsystems.

The lander retrieves information about its position applying the Position Sen- sors (GS, IMU and Radar Altimeter/LiDAR) and about its attitude applying the Attitude Sensors (Star tracker, IMU and Sun sensors). The Guidance Law block produces guidance commands based on the predesigned guidance law and the position information retrieved by the spacecraft. For the critical mission landing phase, the guidance commands are used for applying the Landing Control Logic that controls the Main Engine Actuator to perform the soft landing. As the Main Engine Thruster will need to maintain a correct thrust direction, the Landing Con- trol Logic is coupled with the Attitude Control Logic, by using also the information on the lander attitude and applying as actuators the attitude thrusters. Particu- larly the Guidance Law will be studied in detail implementing further investigation in order to minimize the fuel consumed during the landing “Powered Descent” phase. The KPI of the selected components of the ADCS & GNC subsystems and their values are presented in Table 16.

64 KPI Value Mass 11.35 kg Volume See Table 17 Power 5 - 12 W Accuracy See Table 17 Temperature 80o - 125o − Thrust 22 N

Isp 294 sec Computational complexity TBD Robustness TBD

Table 16: KPI of ADCS & GNC subsystems.

Star Radar KPI Thruster IMU Sun sensor tracker altimeter Mass [kg] 0.450 0.748 0.500 0.072 1.360 Length: Length: Length: Length: 0.186m Diameter: 0.1m 0.069m 0.142m Height: 0.089m Height: Height: Height: Volume 0.055m Height: 0.05m 0.052m 0.086m Width: 0.085m Width: Width: Width: 0.068m 0.1m 0.014m 0.086m Minimum Scale Impulse factor Ac- Accuracy 0.1o - 0.5o 0.03o 0.9144 m Bit: curacy: 0.0089N 300 ppm Reference [39] [40] [41] [42] [43]

Table 17: Volume and Accuracy factors for each component of ADCS & GNC subsystems.

65 EDL Approach

In the following, a description of the EDL approach is presented. At this stage of the project where the mission concept is being built and the outcomes of a Phase 0 study regarding the subsystems are a high level description of what the subsystem should do and which components could be suitable are sufficient. Thus, hereafter, a prelimnary approach of the EDL phase, discussing how this phase could be approached based on previous missions and how the system could look like, is introduced. In addition, first estimations of the thrust levels and the maximum thrust is produced by using two different calculations, justifying the engine configuration been selected. Regarding the ADCS configuration, as it is mentioned above that at this stage a description of the system functions is only required, the performance evaluation of the ADCS system is deemed unfeasible at this stage. Therefore, the ADCS configuration is only described at a level of what it should do and lool like. At later stages, an analysis of the system shall be performed. The EDL phase is one of the most crucial phases of the mission. Inspiration is taken from TeamIndus [44] strategy and its PS configuration (a proposed lunar mission) as well as the already flown mission Chang’e 3 [11] strategy for landing and they are adapted to our case. In general, while in the proximity of the Moon, a series of precise maneuvers are required in order to ensure a soft and safe landing. The major key points that have been followed in previous mission during this phase are:

• After the circularization orbit of 100 km radius is achieved, the entry phase is started.

• A braking maneuver to decrease the periapsis to around 15 km occurs.

• After lowering the periapsis, a major braking at that point takes place to “kill” most of the (horizontal) velocity.

• Quick adjusting of the thrust vector for a powered descent, reducing further the velocity (in both directions).

• Adjust the thrust vector in vertical position for landing.

• Powered landing starts at a range of couple hundreds of meters

66 • Few meters above the surface, the engine is cut-off and a free fall of the lander occurs.

The circularization of the orbit gives more flexibility and landing options. After spotting the landing site, the lowering of the periapsis by firing the main engine is followed. At the periapsis the EDL phase starts and it is the point where the main braking burn of the engines occurs to reduce a major part of the horizontal velocity. After that, the attitude thrusters change the thrust vector, by firing in combination, of the main engine and the rest of the (downward) thrusters in a position where both components of the velocity are reduced. When reaching a near zero horizontal velocity, the attitude thrusters fire again to adjust the main engine and the (downward) thrusters to a vertical position. While in vertical position, the powered landing starts at about 100m and the main engine and the thrusters continue to fire to reduce the vertical velocity. When the vertical velocity reaches a value of 1-2 m/s and at a couple of meters above the surface, the engine and the thrusters are cut-off and a free fall of the lander occurs. A first approach on determining the required thrust magnitude for a lunar descent is described. These preliminary calculations are based on the principles of kinetic energy and work. While in orbit the spacecraft possesses a particular speed, which depends on the altitude that the spacecraft is place at. The spacecraft shall land softly on the Moon’s surface, meaning that the final speed before touchdown is almost zero. Thus, the kinetic energy of the spacecraft decreases from its initial orbital value to approximately zero at touchdown. While descending to the moon’s surface, the gravitational force from the moon does positive work on the spacecraft since the spacecraft is getting closer to the center of the moon. How- ever, this is not the only force that acts to the spacecraft since the kinetic energy is being reduced during descent. This means that the thrusters do negative work on the spacecraft by exerting a force in the opposite direction to its motion. The equations that give the solution as well as the data (Table 18) required are given below. mMoon gMoon = G 2 (16) · RMoon

2 1 u = G m ( ) (17) orbit · Moon R + H − a ! Moon

67 K =0.5 m u2 (18) orbit · lander · orbit

W = ∆K =0 K = K (19) net − orbit − orbit

W = m g ∆y = m g ( H) (20) gravity − lander · Moon · − lander · Moon · −

W = W + W W = W W (21) net gravity thrust ⇒ thrust net − gravity

W W = F D = F = thrust (22) thrust − avg · avg − D

Parameter Value

2 gMoon 1.61 m/s H 15 km D 900 km

11 3 2 G 6.67 10− m /kg s− · · 22 mMoon 7.3 10 kg · RMoon 1740 km Semi-major axis, a 1794.6 km

mlander 263.34 kg

Table 18: Data required for the estimation of the average landing thrust.

The orbital velocity and the kinetic energy of the spacecraft are easily found with equations (2) and (3). The kinetic energy is numerically equal to the net work done on the spacecraft during its descent (with a “minus” sign) since the kinetic energy of the spacecraft upon touchdown is zero. The work done by gravity only depends on the spacecraft’s change in height, as per the approximate equation below. Subtracting the work done by gravity (positive), from the net work (negative) will give an even larger negative value for the work done by the thrusters. Once we have the work done by the thrusters, we can use the estimated path length to find the average force exerted by the thrusters on the spacecraft. The current path

68 length selected for the calculation is 900 km and is taken from [44]. Moreover, the lander mass is calculated, including margins, where the value determination is further discussed below. Notice that equation (5) is a simpler form of the gravity work since the ratio H/RMoon is about 1%, so it can act as a good approximation. The result of the average thrust required to be provided by the thrusters is around 426.26 N. This is only a first estimation of the average thrust and not the maximum thrust required. The thrust capability of the lander consists of 420 N of the main engine and eight 22 N thrusters which sums up to around 600 N. The max thrust is needed during the phase of the major braking as described in [44]. The result shows that there is a quite margin on the thrust capability of the lander. Nevertheless the thrust estimated here is the average and the maximum thrust needed for the major braking phase will be higher than this result. As it will be discussed next, another estimation of the maximum thrust needed for landing will justify the current engines configuration. In the following, a preliminary estimation of the maximum thrust and the different thrust levels through a simulation of the landing trajectory is presented. The lander dynamics are governed by the following equations system, as described in [45]:

r˙ = u (23)

θM˙ = v (24)

T (t) uv u˙ = sinβ(t) 2 (25) m − r

T (t) µ u2 v˙ = cosβ(t) + (26) m − r2 r

T m˙ = (27) −Ispg0 A nominal trajectory is produced without considering any retargeting maneuvers at this stage. In Figure 27 is shown the reference system considered for the simulation. The control variables of the system are the thrust T(t) and the thrust

69 angle β(t) while m is the lander mass, u and v are the horizontal and vertical speed respectively, r and θ are the polar coordinates of the lander position. This is the model followed and is based on [45].

Figure 27: Global landing reference system, [45].

The procedure being followed is: the trajectory is split in 3 phases where the first phase is the major braking, killing most of the horizontal velocity and reaching a value in the range of 5 m/s. After killing the horizontal velocity, the second phase starts in order to achieve an altitude of around 50m and a vertical velocity between 1 and 2 m/s as final conditions. If the conditions are met, the final phase starts which reduces further the velocities in both directions until it reaches an altitude of around 3 m. After reaching the final altitude, the engines are cut-off and a free falling for the remaining couple of meters occurs. The equations describing the lander dynamics are discretized using a constant time step (dt = 0.05) and the simulation calls in each step the different thrust control laws and thrust angle procedures to propagate the lander. These 3 phases are translated to 3 different thrust control laws and 3 different thrust angle approaches. The simulation shows that the specific final conditions are met by implementing: during the major braking phase, maximum thrust is applied constantly and the thrust angle profile follows the one of the Chang’e 3 mission [11] which is considered quite optimized. The thrust angle profile of the Chang’e 3 mission changes linearly, utilizing 2 different slopes, where in the present work it is tried to imitate this behaviour. Then, during the second phase,

70 the thrust angle experiences a large change and it was kept constant to a few degrees while for the thrust variation, the thrust control in Equation 28 is used (combined with Equation 29). In the final phase, the thrust angle is changed to 0 and the thrust variation is calculated by the thrust control in Equation 30. The 2 thrust control laws are implemented based on [12] where a description is available. Equation 28 varies the thrust value until it achieves the required final conditions of the second phase. To calculate the thrust, Equation 29 from classical physics is used, where v0 and h0 are the current speed and altitude and vf and hf are the desired state. Equation 30 is used for the final phase which is a simple ratio between the current altitude rate, v0, and the desired rate, vd. Both control laws are taking into account in every time-step the current altitude, velocity and mass to evaluate the thrust.

T =(a + g ) m (28) d M ·

v2 v2 a = 0 − f (29) d 2(h h ) 0 − f

v0 T =( ) gM m (30) vd · · Next, the results of the simulation are presented and the final conditions are shown. In Table 19 are presented the final conditions achieved. The wet mass of the lander at 15 km, prior to the initialization of the landing procedure is 263.48 kg. It considers margins on ADCS (10%) propellant and on residual (2%) propellant on top of the wet mass in order to have a more conservative (or worst case scenario) evaluation of the landing procedure to compensate for any lack of details at this stage. The maximum thrust required during landing is around 543 N which is close to the maximum capability of the lander. However, the thrust capability of the lander reaches 596 N (1 x 420N main engine and 8 x 22N assist engines) leaving a beneficial margin on the thrust level and also acts as a redun- dant scheme for the assist thrusters in case of a failure. Taking this into account, the current configuration of the PS the lander utilizes is justified.

71 Parameter Value Maximum Thrust required 543.60 N Minimum Thrust required 210.90 N Duration 874.80 sec Downrange 597.37 km Horizontal Velocity -0.74 m/s Vertical Velocity -1.00 m/s Landed mass 129.87 kg Starting mass 263.34 kg

Table 19: Final results of the simulation.

In Figure 28 and 29, the trajectory followed by the lander as a function of time and downrange are shown. The former graph shows the timeline of the landing procedure, where the 3 different phases are evident. The duration of the landing is about 900 seconds and the lander spends most of its time in the first phase which a significant reduction of its velocity occurs. The duration of the other 2 phases are smaller and they take place in low altitudes. The latter graph shows the variation in altitude relative to the surface distance travelled by the lander. It can be seen that most of the distance covered is during the major braking phase due to the associated high velocity reduction which requires more time to spend on this maneuver. The other 2 phases start to occur when the lander is close to the landing site, controlling the attitude of the lander and maintain a constant descent for a soft landing. It is worth to mention that during the main braking phase the lander gets to a higher altitude before starting to actually descent due to the associated orbital velocity at the beginning of the landing procedure. Thus, the breaking of the lander results to an altitude gain. However, after killing the majority of the velocity the lander will eventually start to descent since the gravity becomes the major factor. In Appendix B, a closer look to the last 2 phases of the landing regarding the trajectory changes can be taken (Figure 32 and 33).

72 Figure 28: Altitude as a function of time during the landing phase.

Figure 29: Altitude as a function of the surface distance covered throughout the landing phase.

In Figure 30, the thrust angle, the thrust and the mass varying with time is shown. As it can be seen, the procedure in the thrust angle variation implements as already mentioned the Chang’e 3 one during the major braking where the thrust angle varies linearly, consisting of 2 different slopes. For the next 2 phases, the

73 thrust angle is changed to a few degrees during the second part of the landing and zero for the final part. Regarding the thrust variation, maximum thrust is used during the major braking phase while for the last 2 parts the control laws presented above are used to calculate the required thrust at each moment. In the bottom sub-figure the mass variation is shown where an almost linear mass change over time is evident. The mass at start is the wet mass of the lander while at 15 km and it goes down to the dry mass of the lander at the surface. The final result of the landed mass is 129.87 kg while the calculated dry mass is 127.58 kg which gives a margin on the propellant needed to be carried on-board for any additional maneuver might be required.

Figure 30: Thrust angle, Thrust and Lander mass as a function of time during landing phase.

In Figure 31, the 2 components of the velocity are presented. The first graph shows the variation of the horizontal component of the velocity for all phases. Dur- ing the main braking the horizontal velocity is mostly reduced in order to facilitate the soft landing. Thus, a steep slope is evident during that phase. Having the horizontal component reduced, the next 2 phases reduce steadily and with smaller slopes the horizontal velocity until it reaches the desired final condition. The result shows that the horizontal velocity is close to zero. The minus sign in the final value indicates a velocity that is on the opposite side of the lander movement which could indicate a movement on the opposite direction. However, the value is rather small and have negligible effect. The second graph shows the evolution of the ver-

74 tical component of the velocity. Again, most of the variations are taking place in the main braking phase. During that phase, it can be seen the gain in vertical velocity (positive values) which is translated in a gain in altitude as explained above. After that gain, the lander gets velocity in the vertical direction which is translated into the acceleration due to the actual descent of the lander by the effect of the gravity. During the second phase, the lander reduces significantly the vertical velocity to reach the desired condition while in the last phase the vertical velocity is reduced steadily to enable the soft landing. The final condition achieved for the vertical velocity is about 1 m/s which is in the range of previous missions flown such as Chang’e 3 [11]. The minus sign is due to the fact that the positive axis of the vertical components is facing upwards. Again, in Appendix B, a closer look to the last 2 phases of the landing regarding the velocities variations is presented (Figure 34).

Figure 31: Horizontal and Vertical velocities as a function of time during landing phase.

The simulation performed has more than enough space for improvement but at this stage shows the basic physics behind the landing procedure. It is needed to highlight that a lot of assumptions based on previous studies are implemented and that the parameters chosen for the simulation are tuned in order for the simulation to give results. The simulation is sensitive in parameters changes and a further

75 investigation is going to be carried out during the next phases of the project. Also, an effort on optimization of the trajectory will be spent at later stages which will give a comparison on the fuel savings compared to the baseline landing trajectory. To summarize, a description of a landing procedure to be followed is carried out and first estimations on thrust levels are quantified. The results are giving a good first approach of the physics behind landing where basic models in combination with procedures based on previous missions are used in order to produce measurable outcomes. Further investigation on the model used and on optimization methods will give a more comprehensive and extended comparison of the results.

76 4 Conclusions

A space propulsion system for a commercial micro moon lander was proposed. Various trajectories were described and a preliminary ∆V budgets for a Hohmann and a Bi-elliptic transfer were extracted. An establishment of top-level requirements, where mass and volume showed a big impact on the design and a trade-off analysis of different PSs were carried out. Also, a design philosophy for dimensioning the PS was described and followed. Important factors affecting the mission were taken into account and the concept of a commercial mission was spec- ified. An effort to use COTS components was realized where some of the basic components were able to implement the concept. The trade-off analysis showed that a suitable PS for the mission was a unified Bi-propellant system which takes advantage of the common feeding system, sharing the same tanks which provides a system with high performance and keep the complexity relatively low. Implementation of the electric propulsion was deemed unsuitable for the mission despite the significant fuel savings. However, the selected PS violates the mass limit. This could be solved with decreasing the payload capability as shown above or implementing a bi-elliptic transfer which will decrease the fuel required due to smaller ∆V budget but it will increase the mission time considerably. An EDL approach was described and calculated where inspiration was taken from previous missions. The interfaces of the ADCS & GNC subsystems with the Propulsion subsystem were described and a preliminary architecture layout was presented. Also, a functional block diagram was introduced, describing the main functionalities of all subsystems. Finally, the KPIs were identified along with their values and a preliminary selection of components for all three subsystems was introduced and their specifi- cations were presented. Part of the future work will be the detail analysis for the design of the propulsion system including the production of schematics, redundancy and reliability concepts. The top-level requirement matrix will be expanded in a more detailed matrix where the description of the PS with numbers will be established. A preliminary functional block diagram was presented, however a more detailed functional

77 breakdown analysis is going to be carried out. Also, refinement and optimization of the trajectory using more detailed simulations will sharpen the ∆V budgets and trade-off on different trajectories will investigate the implementation of different trajectory strategies. Another area of interest is the launch scenario which takes a significant portion of the total cost. Different trades on various launch scenarios and launch vehicles will give different ∆V budgets, trajectories and thus different fuel requirements. This will drive the selection of the PS where different options could be implemented. Moreover, a trade space on different lander sizes will have a direct effect on the PS design, leaving the space for investigating various options and scenarios. The EDL approach and in particular a more detailed calculation of the required thrust levels during landing phase will be carried out. EDL phase is mission critical and thus sophisticated maneuvers and strategy are required. This will affect the configuration of the engines in order to meet the specific maneuvers and the thrust levels required during landing phase. Hence, an investigation and a trade-off on different configurations and number of engines will be carried out to meet the soft landing requirement. Also, efforts on models improvements and on optimization methods are needed for the next phases of the project.

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82 Appendices

83 Appendix - Contents

A Trajectory Calculations 85

B Simulation Graphs 87

C Requirements ranking justification 89

84 A Trajectory Calculations

Hereafter, the basic equations used in matlab to calculate the ∆Vs required for the bi-elliptic trajectory are presented. The ∆V calculation is based on the vis-viva equation. An explanation with a description of the matlab simulation code can be found here: https://de.mathworks.com/ matlabcentral/fileexchange/38885-bi-elliptic-transfer-between-coplanar-circular- orbits?focused=3790493&tab=function The orbital velocities of different orbits and different points are given below:

µ v = (31) i r ! i µ v = (32) f r ! f

2µ µ v = (33) p1 r − α ! i 1 2µ µ v = (34) α1 r − α ! a 1 2µ µ vp2 = (35) " rf − α2 2µ µ v = (36) α2 r − α ! a 2 The semi-major axes of the 2 ellipses are calculated with the equations below.

r + r α = i α (37) 1 2

r + r α = α f (38) 2 2 The transfer time from the initial orbit to the final orbit is given by:

α3 α3 τ = π 1 + π 2 (39) " µ " µ

85 The magnitude of the impulses required are calculated by:

∆V = v v (40) 1 p1 − i

∆V = v v (41) 2 2α − 2α

∆V = v v (42) 3 f − 2p

86 B Simulation Graphs

Here are presented additional graphs of the landing simulation. In particular it is shown in a zoom-in mode different results of the 2 last phases of the landing procedure.

Figure 32: Altitude vs Time.Zoom-in the last 2 phases of the landing.

Figure 33: Altitude vs Downrange. Zoom-in the last 2 phases of the landing.

87 Figure 34: Horizontal and Vertical velocities as a function of time. Zoom-in the last 2 phases of the landing.

88 C Requirements ranking justification

89 Weighting Mass Volume Performance Required ΔV Required Multiple ASAP justification Thrust ignitions compatibility Mass Same More ΔV More More Same importance important requirements important important importance due to LV due to LV more due to the due to LV due to LV requirements. requirements. important soft landing requirements. requirements. because it requirement. The lander drives the Thus, shall have a mission required mass below design. It thrust drives the ASAP shows the the success of limit. velocity the mission. increments the PS requires to deliver and acts as an index of the required propellant which drives the mass and size of the lander. Volume More ΔV More More Same important requirements important important importance due to LV more due to the due to LV due to LV requirements. important soft landing requirements. requirements. because it requirement. The lander drives the Thus, shall have a mission required volume below design. It thrust drives the ASAP shows the the success of limit. velocity the mission. increments the PS requires to deliver and acts as an index of the required propellant which drives the mass and size of the lander. Performance ΔV More Same More requirements important importance. important more due to the Both PS due to LV important soft landing performance limitations. because it requirement. for a lander The lander drives the Thus, mission needs shall respect mission required to be high and the ASAP design. It thrust drives multiple requirements shows the the success of ignitions of to be able to velocity the mission. the PS are fly. increments needed due the PS to the requires to landing. deliver and acts as an index of the required propellant which drives the mass and size of the lander. Required ΔV ΔV ΔV ΔV requirements requirements requirements more more more important important important because it because it because it drives the drives the drives the mission mission mission design. It design. It design. It shows the shows the shows the velocity velocity velocity increments increments increments the PS the PS the PS requires to requires to requires to deliver and deliver and deliver and acts as an acts as an acts as an index of the index of the index of the required required required propellant propellant propellant which drives which drives which drives the mass and the mass and the mass and size of the size of the size of the lander. lander. lander. Required Thrust More More important important due to the due to the soft landing soft landing requirement. requirement. Thus, Thus, required required thrust drives thrust drives the success of the success of the mission. the mission. Multiple ignitions ASAP compatibility is more important because it is a mission driver for the feasibility of the lander design. ASAP compatibility Storability Thrust controllability TRL/Off the shelf components Propellant suitability Temperature Chamber operation pressure

Figure 35. Requirements ranking justification. Weighting Storability Thrust TRL/Off the Propellant Temperature Chamber justification controllability shelf suitability operation components pressure Mass More important More important More important More important More important More important due to LV due to LV due to LV due to LV due to LV due to LV requirements. requirements. requirements. requirements. requirements. requirements. Volume More important More important More important More important More important More important due to LV due to LV due to LV due to LV due to LV due to LV requirements. requirements. requirements. requirements. requirements. requirements. Performance Same Less important. More important importance. More important More important More important Storability of because these High because these because these because these propellants is kinds of performance is kinds of kinds of kinds of more important missions required for missions missions missions because it is required high PS these kinds of required high PS required high PS required high PS associated with performance missions but performance performance performance safety issues for which drives the also COTS which drives the which drives the which drives the these kind of mission success components are mission success mission success mission success missions during and the important due and the and the and the both ground propellant need to the propellant need propellant need propellant need operations and and thus the size commercial and thus the size and thus the size and thus the size mission of the PS. nature of the of the PS. of the PS. of the PS. operations. mission. Required ΔV ΔV ΔV requirements ΔV requirements ΔV requirements ΔV requirements ΔV requirements requirements more important more important more important more important more important more important because it drives because it drives because it drives because it drives because it drives because it drives the mission the mission the mission the mission the mission the mission design. It shows design. It shows design. It shows design. It shows design. It shows design. It shows the velocity the velocity the velocity the velocity the velocity the velocity increments the increments the increments the increments the increments the increments the PS requires to PS requires to PS requires to PS requires to PS requires to PS requires to deliver and acts deliver and acts deliver and acts deliver and acts deliver and acts deliver and acts as an index of as an index of as an index of as an index of as an index of as an index of the required the required the required the required the required the required propellant which propellant which propellant which propellant which propellant which propellant which drives the mass drives the mass drives the mass drives the mass drives the mass drives the mass and size of the and size of the and size of the and size of the and size of the and size of the lander. lander. lander. lander. lander. lander. Required Thrust More important More important More important More important More important More important due to the soft due to the soft due to the soft due to the soft due to the soft due to the soft landing landing landing landing landing landing requirement. requirement. requirement. requirement. requirement. requirement. Thus, required Thus, required Thus, required Thus, required Thus, required Thus, required thrust drives the thrust drives the thrust drives the thrust drives the thrust drives the thrust drives the success of the success of the success of the success of the success of the success of the mission. mission. mission. mission. mission. mission. Multiple Same Same Same More important More important More important ignitions importance importance importance because the PS because the PS because the PS because they because they because they has to deliver has to deliver has to deliver are driving are driving are driving multiple burns multiple burns multiple burns factors for the factors for the factors for the during the during the during the efficiency and efficiency and efficiency and landing phase as landing phase as landing phase as reliability of the reliability of the reliability of the well as during well as during well as during PS to deliver all PS to deliver all PS to deliver all the transfer the transfer the transfer the required the required the required manoeuvres. manoeuvres. manoeuvres. manoeuvres. manoeuvres. manoeuvres. ASAP More important More important More important More important More important More important compatibility because ASAP because ASAP because ASAP because ASAP because ASAP because ASAP compatibility is a compatibility is a compatibility is a compatibility is a compatibility is a compatibility is a mission driver mission driver mission driver mission driver mission driver mission driver for the for the for the for the for the for the feasibility of the feasibility of the feasibility of the feasibility of the feasibility of the feasibility of the lander design. lander design. lander design. lander design. lander design. lander design. Storability Same Same Same importance importance importance because they because they because they More important More important affect the affect the affect the due to safety due to safety efficiency and efficiency and efficiency and and reliability and reliability reliability of the reliability of the reliability of the factors. factors. PS to deliver all PS to deliver all PS to deliver all the required the required the required manoeuvres. manoeuvres. manoeuvres. Thrust Both of them controllability drive the success of the mission. Thrust More important controllability is More important More important because the PS needed due to because the PS because the PS has to deliver the landing has to deliver has to deliver different thrust phase and the different thrust different thrust levels during the propellant levels during the levels during the landing phase. suitability is landing phase. landing phase. needed in order for the PS to be efficient and reliable. TRL/Off the Same COTS COTS shelf importance components are components are components because COTS important due important due are important to the to the for commercial commercial commercial missions while propellant nature of the nature of the suitability mission. mission. affects mission safety and PS efficiency. Propellant Same Same suitability importance importance because both because both affect the affect the efficiency efficiency delivered by the delivered by the PS. PS. Temperature Same importance because both affect the efficiency delivered by the PS. Chamber operation pressure

Figure 35. Requirements ranking justification (Continued).