Preliminary Mission Design - NANOSTAR...Reaction Wheels ADCS 0.5 Star Sensor ADCS 0.5 OBC + PCU DH...

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Preliminary Mission Design

OPEN-EYES team

Blanca Boado Cuartero Álvaro González Fariña Javier Luis González Monge Marcos Rodríguez Rodríguez Sergio Zambrano Hayas

Open-eyes preliminary mission design.

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TABLE OF CONTENTS

EXECUTIVE ABSTRACT ..................................................................................................................3

INTRODUCTION ............................................................................................................................4

APPLICABLE AND REFERENCE DOCUMENTS ..................................................................................5

MISSION OVERVIEW AND REQUIREMENTS FLOW DOWN..............................................................6

Preliminary Design Ideas ........................................................................................................................ 8

Mission phase 0-A: mission analysis and feasibility ............................................................................ 10

SUBSYSTEMS ANALYSIS AND DESIGN ......................................................................................... 11

Product Tree and Work Breakdown Structure .................................................................................... 11

Mission analysis ................................................................................................................................... 12

Systems operations modes .................................................................................................................. 17

Space propulsion Subsystem ............................................................................................................... 18

Attitude, determination and control subsystem ................................................................................. 19

Communications subsystem and ground segment ............................................................................. 21

Command and data handling subsystem............................................................................................. 23

Electric Power subsystem .................................................................................................................... 23

Mechanical design and structure ......................................................................................................... 30

Thermal control subsystem ................................................................................................................. 33

Additional payloads ............................................................................................................................. 35

RISK ANALYSIS AND MITIGATION ............................................................................................... 36

Risk management policy ...................................................................................................................... 36

Risks identification and assessment .................................................................................................... 38

CONCLUSIONS ............................................................................................................................ 40

APPENDICES ............................................................................................................................... 41


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EXECUTIVE ABSTRACT

In the context of space exploration and science, nanosatellites have proven to be a useful, cost-

efficient tool for simple short-duration missions. In this document, the preliminary mission design

analysis has been carried out by the team members of Open-Eyes.

The mission proposed by NANOSTAR project team consists on the preliminary design of a

nanosatellite with the objective of performing one (or more) lunar fly-by(s), in order to get surface

pictures from the moon. These pictures will prove useful in the determination of whether there is or

there was ice in the lunar poles.

For the accomplishment of this mission, the Open-Eyes team have worked together with a concurrent

design philosophy since the beginning of the project, working simultaneously in different subsystems

and sharing all the information, and has achieved an optimal solution according to mission

requirements.

After some iterations of research and analysis, the team took a decision. Whereas the mission

requirement is to perform a lunar fly-by, its scientific objective is to get information regarding ice

formations on lunar poles. To better serve this purpose, the team has found a transfer orbit that will

allow the satellite to stay on lunar orbit by using reasonable impulses. The team has also thought of

the End of Life, allocating a small amount of propellant that will allow the disposal of the satellite,

thus finishing the mission.


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INTRODUCTION

In the context of NANOSTAR project, this document is presented in order to share the preliminary

mission design study for the proposed mission. The document contains the following sections:

• Mission overview and mission requirements

• Subsystems analysis and design

• Risks analysis

• Conclusions

Open-Eyes team members will be acting as experts in different subsystems. For the organization of

this mission, the team has decided to split the different subsystem according to their complexity and

criticality to the mission as indicated in the Table 1:

TABLE 1. TEAM MEMBERS AND SUBSYSTEMS ALLOCATION.

Team member Priority subsystem Secondary subsystem

Boado Cuartero, Blanca Thermal Control ADCS

González Fariña, Álvaro Power Configuration

González Monge, Javier Luis Communications Payload

Rodríguez Rodríguez, Marcos Systems Engineering Structure

Zambrano Hoyas, Sergio Orbit design Propulsion

Although each member has been designated as the owner of two subsystems, Open-Eyes work as a

team, so any member will be supporting the rest of the team whenever the workload or the

complexity of the problem requires this. This is also a key point for concurrent designing, since all the

information needs to be shared amongst the different experts to achieve a successful preliminary

design.


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APPLICABLE AND REFERENCE DOCUMENTS

TABLE 2. REFERENCES.

Reference Author

State of the Art: Small Spacecraft Technology NASA

ECSS-E-ST-10C (15February2017) ESA-ESTEC

ECSS-M-ST-80C (31July2008) ESA-ESTEC

Margin philosophy for science assessment studies ESA

A3200 Datasheet On-board Computer System for mission critical space

applications NanoMind

User Manual X-Band Single Element Patch Antenna ENDUROSAT

NanoTorque GSW-600 Datasheet GomSpace

Space Sextant Datasheet ADCOLE MARYLAND

AEROSPACE MAI-SS

Sun Sensor NFSS-411 Datasheet NewSpace

TABLE 3. EXAMPLE OF A CHANGE LOG RECORD TABLE.

Edition/Revision Date Description of the change

V1.0 11/05/2019 Initial version of the document

V1.1 13/05/2019 Added Conclusions, Mass budget and IDM model.

V2.0 -

V3.0 -

-

-


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MISSION OVERVIEW AND REQUIREMENTS FLOW DOWN

In the context of space exploration, the human eyes have been set many times on The Moon. One of

the key points for human colonization is the presence of water. While liquid or vaporized water

cannot persist on lunar environment, ice could survive on polar craters.

The aim of Nanostar mission is to perform a lunar fly-by in order to take as many pictures as possible

of these possible ice formations with the minimum cost. In order to achieve this goal, a list of mission

requirements will be presented, and these requirements will cascade into smaller, more specific

requirements organized by the subsystems involved.

This requirement list and flow down is presented on Table 4:

TABLE 4. REQUIREMENTS FLOWDOWN.

Req. ID Level Subsystem(s) Requirement Description

MR001 1 -

The system shall carry and activate safely the main science payload described below, for the maximum time around

the periselenium pass, thus maximizing the total amount of science data received on Earth. Additional payloads (e.g. technological demonstrators) can be included if they do

not have a detrimental effect on the operation of the main one and the system can correctly support their operation. This will be considered positively in the evaluation of the

project.

MR001-01 2 Power System shall have enough power available to safely activate the payload when necessary

MR001-02 2 PL, Structure The system may carry additional payloads (e.g. scientific

demonstrators) as long as all other mission requirements. are met without any risk.

MR001-03 2 Thermal

System shall assure that the PL is working under the following range of temperatures:

-10ºC, 30ºC when operating -20ºC, 40ºC when inactive

MR001-04 2 ADCS PL camera shall not point directly to the sun

MR002 1 -

The altitude of the periselenium pass shall not be higher than 100 km (minimum periselenium altitude can be freely

chosen, although it should be defined according to a risk trade-off analysis)

MR002-01 2 Mission

Analysis Orbit shall pass closer than 100km above moon's surface

when performing the fly-by

MR002-02 2 Propulsion Propulsion system shall guarantee that the satellite can

reach an orbit with a periselenium pass of, at least, 100km above lunar surface


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Req. ID Level Subsystem(s) Requirement Description

MR003 1 - At least one Moon's flyby shall be performed. Additional

Moon flybys will be considered positively in the evaluation of the project.

MR003-01 2 Orbit Orbit shall assure at least one periselenium pass

MR004 1 - The science data obtained by the main science payload

shall be transmitted to Earth within the mission timeframe.

MR004-01 2 Mission Analysis Orbit shall have at least one access with Ground Stations

MR004-02 2 Communications Communication system shall be able to transmit all

mission data to Ground Stations during the total duration of the access(es).

MR005 1 - The satellite shall be capable of performing the mission

objectives, considering the space environment constraints.

MR005-01 2 Thermal The system shall assure that all hardware include is operating within a safe range of temperatures

MR006 1 - The satellite shall guarantee the correct pointing of the main science payload during the flyby(s).

MR006-01 2 ADCS System shall guarantee the correct pointing of the

satellite when the PL is operating. This pointing shall have an accuracy of, at least, 0.5 deg.

MR007 1 - The satellite volume shall not exceed 27U. A smaller

volume will be considered positively in the evaluation of the project.

MR008 1 - The mission duration from launch to end-of-life shall not

exceed 5 years. A shorter duration will be considered positively in the evaluation of the project.

MR008-01 2 Communications All mission data shall be transmitted within 5 years from the start of the mission

MR009 1 - Ground segment(s) shall rely only on the tracking stations of the ESA network

MR010 1 - The uplink and downlink frequencies shall be in the S, L,

UHF, X or C bands.nk frequencies shall be in the S, L, UHF, X or C bands.


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PRELIMINARY DESIGN IDEAS

Team’s first approach to the mission was focused on the lunar transference. The satellite is injected

into a GTO orbit, and that is the starting point for the maneuver. There were two major options so

the team had to choose to go with either electric propulsion or chemical propulsion.

Most of the nanosatellites are equipped with electrical propulsion, which is known to be effective

due to its very high specific impulses (which means that the propellant mass required is low).

However, the thrust provided by these kinds of engines is rather low, so the transfer times are

enormous. By using chemical propellant, the team could achieve two things:

• Short transfer time with the possibility to stay orbiting the moon.

• Use of innovative engines (TRL ≥ 5) with green propellants.

Thus, this mission could serve as an opportunity to spread the using of chemical engines on

nanosatellites, allowing new and faster maneuvers.

Having chosen the propulsion type, the next step was to decide on the trajectory itself. The first idea,

taken from previous lunar mission such as the Apollo program, was to perform a free-return

trajectory.

FIGURE 1: FREE RETURN TRAJECTORY.

The aim behind this trajectory is to use the gravitational pull from the moon to help the spacecraft

return to Earth, saving the cost of an impulse on lunar orbit. The problem with this maneuver was

that it is thought to be a return trip, but it is not supposed to go back to the Moon. This meant that

there was a possibility to perform one flyby and, maybe after returning to Earth orbit, perform some

corrections in order to:

• Stay orbiting Earth and include secondary payloads to be used at this point.

• Give an impulse and go back to the Moon.


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The team considered this but, given that the scientific objective of the mission is to research the

presence of ice on lunar poles, the team decided to analyze the possibility to stay on lunar orbit. This

analysis was successful, as will be presented later in this document.

At that point the orbit and propulsion system were pre-defined. However, since for the propellant

sizing the initial mass of the satellite had to be given, the team supposed an initial number of 15

cubesat units. For the rest of the subsystems:

• Operational modes can be defined, and knowing typical ranges of consumption, the power

needs can be estimated for a first iteration.

• Knowing the approximate duration of eclipses and average power consumptions of payload

and OBC, the power can be estimated, and this leads to giving a sizing for the solar panels and

battery capacity.

• With the power consumption, the amount of heat to be dissipated can be estimated,

therefore the thermal subsystem can be sized.

• With all the weights known, the number of cubesat units can be better estimated and the

next iteration begins.

The final configuration estimated consisted in 12 units, and their allocation is as following:

TABLE 1: NUMBER OF UNITS ESTIMATED AND THEIR ALLOCATION.

Element Subsystem Allocated units Propellant PROP 4

Engine PROP 1

Payload PL 2.5

Battery PWR 0.5

Reaction Wheels ADCS 0.5

Star Sensor ADCS 0.5

OBC + PCU DH 0.5

TOTAL

9.5

Available space for extra payloads

2.5

With this final configuration, and taking into account that the satellite will orbit The Moon thus being

able to extract huge amounts of information, the mission duration can be reduced to one year, and

this will imply that some subsystems can be simplified (e.g., the battery can operate on a more

aggressive range of discharge).


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MISSION PHASE 0-A: MISSION ANALYSIS AND FEASIBILITY

The objective of this preliminary design encompasses the Phase 0 and Phase A. According to the

ECSS-E-ST-10C-Rev.1, these phases comprise the following tasks:

PHASE 0: MISSION ANALYSIS-NEED IDENTIFICATION

For Phase 0, the system engineering function:

• Supports the identification of customer needs.

• Proposes possible system concepts.

• Supports the Mission Definition Review (MDR) and ensures implementation of the MDR

actions.

• Performs an analysis of the Mission Statement document and integrates this analysis and any

relevant contribution from lower level suppliers into a Mission Description document(s) in

conformance with Annex B, and maintains this latter document for the final selected concept.

• Proposes the requirements against the expressed user needs for agreement with the

customer.

PHASE A: FEASIBILITY

For Phase A, the system engineering function:

• Finalises the expression of the needs identified in Phase 0.

• Proposes system solutions (including identification of critical items and risks) to meet the

customer needs.

• Supports the Preliminary Requirement Review (PRR) and ensure implementation of PRR

actions.

• Finalises the validation of the requirements against the expressed needs together with the

customer.

Although this is the theoretical scope of these two initial phases, some things are different for this

project. The first one is due to the fact that the main mission requirements are already provided, so

the action to be taken by the team is to cascade these initial requirements into smaller, more specific

requirements. The second one is that there won’t be any review, so there no need to prepare

documents or milestones such as the PRR.


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SUBSYSTEMS ANALYSIS AND DESIGN

The aim of this section is to present the different subsystems included in the preliminary design

analysis. This study includes the subsystem design according to mission requirements, the initial

sizing and their interaction with the rest of subsystems.

PRODUCT TREE AND WORK BREAKDOWN STRUCTURE

One of the first steps of every project consists in defining the elements and work packages that will

constitute the project and final product. For this purpose, the product tree and work breakdown

structure are defined.

FIGURE 2: PRODUCT TREE.

The definition of the Product Tree is focused on all the subsystems included in the satellite. External

elements to the satellite (such as Ground station or Test facilities) have not been included. All

management and systems engineering tasks are included in the WBS, which is presented in the

following Figure:


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FIGURE 3: WORK BREAKDOWN STRUCTURE.

MISSION ANALYSIS

The orbit design is the key to manage a successful mission, as this element determines the trajectory

followed by the satellite, and therefore, the eclipses of the satellite and the access times. This fact

influences in the rest of subsystems, like the electric power one, the thermal control one or the

communication system, so the orbit design is a critical part of the mission. The main objective of this

section is to obtain the ∆𝑉 budget of the mission.

For this project, it is necessary to make a fly-by mission to the moon. The main requirement of this

section is to achieve a periselenium pass lesser than 100 km. Also, additional flybys will be considered

positively.

To make the design, GMAT was the software chosen. The study of the launcher is out of the reach of

the project, so it is supposed that it injects the satellite into free GTO. To determine the orbital

parameters of this orbit, it is necessary to study the inclination, the RAAN and the argument of

perigee of the Moon orbit around the Earth in the date of the mission (02 Dec 2025 11:59:23.000).

This data is given by GMAT, and fixing the altitude of the perigee in 600 km, the orbital parameters

of the GTO are the following:

TABLE 5. ORBITAL PARAMETERS OF THE GTO.

SMA [km] ECC INC [o] RAAN [o] AOP [o] TA [o]

24570 0.716 23.15 12.92 185 20

It is important to stand out that the inclination of this orbit may suffer changes, depending on the

inclination of the orbit of arrival, as it is thought that water can be found in the poles, so a polar orbit

is the appropriate one. The GTO will be shown in the next Figure:


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FIGURE 4. GTO IN GMAT.

The first idea to arrive to The Moon was to make a Hohmann trajectory to reach the L1 point of the

system Earth-Moon and use the instability of this point to get a boost and minimize the necessary

∆𝑉. Finally, this option was discarded because of its difficulty of implementing in GMAT.

The final design consists in making a Hohmann trajectory, starting in the perigee of the GTO to

minimize the ∆𝑉 of the impulse and ending as close as possible to the moon. It is important to stand

out that there is a moment in which the gravitational field of the Moon attracting the satellite is

greater than the Earth one, just when the satellite crosses the sphere of influence of the Moon, at

66, 100 km from the center of it. At that point, it is necessary to change the propagator of GMAT,

considering The Moon as the main body. This perturbation increases the initial impulse given to the

satellite, but it is not necessary to apply a correction maneuver. Giving a ∆𝑽 of 𝟎. 𝟕𝟎𝟒 𝐤𝐦/𝐬, the

satellite reaches 80 km of altitude in its passage through the periselenium, but the arrival inclination

was 20.2o, so it was necessary to change the inclination of the GTO to 93.15o, to manage that the

satellite passes next to the pole of the Moon. The transfer orbit is the next one:

FIGURE 5. TRANSFER ORBIT TO THE MOON.


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In this image two colours can be observed. The red one is the trajectory section in which the Earth is

the main body, and when it turns into blue, the satellite has crossed the sphere of influence of the

Moon. The satellite takes 4.1 days to complete this trajectory.

The fly-by and the pass through the periselenium (𝟖𝟎 𝐤𝐦 from the surface of the Moon) can be

observed in the next image:

FIGURE 6. LUNAR FLY-BY.

The next step is trying to obtain more flybys. Go back to the Earth was simulated in GMAT, but as the

outgoing orbit was far away from the Earth, the necessary impulse was too high for a nanosatellite

(approximately 4 km/s). At this point, many ideas were put in the table and, although it is not a

“classical” fly-by, stay orbiting the Moon was the final solution to obtain more photos of The Moon

poles. To do that, it is necessary to give another impulse just in the periselenium, in the opposite

direction of the local velocity of the satellite. Giving a ∆𝑽 of 𝟎. 𝟐𝟖𝟖 𝐤𝐦/𝐬, the resulting orbit is the

next one:

FIGURE 7. ORBIT AROUND THE MOON.

And the orbital parameters of this orbit are illustrated in the next table:


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TABLE 6. ORBITAL PARAMETERS OF THE LUNAR ORBIT.

SMA [km] ECC INC [o] RAAN [o] AOP [o] TA [o]

15078 0.872 89.43 128.70 141.57 187.45

The eccentricity of the orbit has been chosen high to minimize the last ∆𝑉 given. In the next image,

one can observe that the satellite passes through the north pole (the orbit is perpendicular to the XY

plane, drawn in blue):

FIGURE 8. INCLINATION OF THE FINAL ORBIT.

The period of the final orbit is calculated as follows:

𝑇 = 2𝜋√𝑎3

𝜇= 166140 s = 1.923 Earth days

So, the duration of the mission is determined by the duration of the transfer orbit and the number

of orbits given around The Moon. The decision to study The Moon for a year has been taken, which

may be variable depending on final adjustments.

And to conclude, in the next image one can observe the whole trajectory followed by the satellite,

seen from the Earth:


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FIGURE 9. TRAJECTORY FOLLOWED BY THE SATELLITE DURING THE MISSION.

Finally, in the next table, the impulses given in the whole mission are summarized, considering a 5%

margin according to ESA:

TABLE 7. ∆𝑽 BUDGET OF THE MISSION.

∆𝑽𝟏 [𝐤𝐦/𝐬] ∆𝑽𝟐 [𝐤𝐦/𝐬] ∆𝑽𝒕𝒐𝒕𝒂𝒍 [𝐤𝐦/𝐬] 0.703 ± 0.035 0.288 ± 0.014 0.991 ± 0.050


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SYSTEMS OPERATIONS MODES

On this section, the different system operations modes will be described. An operation mode is

defined by the subsystems that are active and the phase of the mission during which the satellite

operates in such mode.

The operation modes for Open Eyes satellite have been defined as following:

TABLE 8: OPERATION MODES.

No. Mode Description Active subsystems

OM-01 Deployment/Stabilization Initial stabilization operations after satellite deployment

ADCS/DH

OM-02 Data collection PL activation for data collection ADCS/DH/PL

OM-03 Data transmission Mission data transmission to GS ADCS/DH/COM

OM-04 Ground Communication Minor communication with GS (status check, GS control)

ADCS/DH/COM

OM-05 Power conservation Minimum power consumption (housekeeping and attitude control)

ADCS/DH

• OM-01: deployment/stabilization is defined in order to perform the initial operations of

stabilization and orientation after the satellite separates from the launcher’s upper stage.

• OM-02: data collection is defined for the mission phases when the Payload activates in order

to collect information.

• OM-03: data transmission activates when the satellite transmits the collected information

from payloads to Ground Station.

• OM-04: Ground communication activates whenever Ground Station needs to communicate

with the satellite or vice-versa in order to share secondary information (such as attitude and

trajectory parameters, battery level, etc.) or to transmit software updates.

• OM-05: Power conservation is defined for the mission phases during which the power

consumption must be the lowest, for example, during eclipses. Only basic housekeeping and

attitude control will be active.


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SPACE PROPULSION SUBSYSTEM

The propulsion subsystem is directly related to mission analysis. As an impulsive burn has been

chosen in GMAT, it is necessary to choose a chemical propellant engine for cubesats.

Open-eyes is a project adapted to the “new times”, so consciousness with the environment is one

thing to keep in mind. That is the reason why “green fuel” propellants have caught the team

attention, as they have a reduced toxicity due to the lower danger of component chemicals and

significantly reduced vapor pressure compared to hydrazine. The next image lists the current state-

of-the-art in green propellants:

FIGURE 10. CURRENT STATE-OF-THE-ART IN GREEN PROPELLANTS.

Source: NASA

The GR-22 system is the one that better adapts to the needs of the mission, since it provides the

necessary thrust to obtain low combustion times, so the impulse can be considered instantaneous.

Furthermore, its TRL status is in a good condition to accelerate its development, being able to be

“Open-Eyes” the first mission using this type of thruster.

With the specific impulse of the system, an estimation of the propellant mass can be done with the

Tsiolkovsky equation:

∆𝑉 = 𝐼𝑠𝑝ln (𝑀0

𝑀0 −𝑀𝑝) ; 𝑀𝑝 = 𝑀0 (1 − 𝑒

−∆𝑉𝐼𝑠𝑝)

The propellant mass used in each manoeuvre appears in the next table:


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TABLE 9. PROPELLANT MASS USED DURING THE MISSION.

𝑴𝒑𝟏 𝑴𝒑

𝟐 𝑴𝒑𝒕𝒐𝒕𝒂𝒍

0.251𝑀0 0.084𝑀0 0.335𝑀0

It is necessary to calculate the occupied volume by the propellant. The density of the HAN is

1.84 g/cm3 (1.84 kg/u), so:

𝑉𝑝𝑟𝑜𝑝 =𝑀𝑝𝑡𝑜𝑡𝑎𝑙

𝜌= 0.1821 ∙ 𝑀0 u

And the combustion time is calculated as follows:

𝐼𝑠𝑝 =𝐸

�̇�=𝐸 ∙ 𝑡𝑏𝑀𝑝

; 𝑡𝑏 =𝐼𝑠𝑝 ∙ 𝑀𝑝

𝐸= 32.6 ∙ 𝑀0 s

Assuming that 12U is the size of the satellite (see configuration section), estimating 1.33 kg per unit

and 1.32 kg for the solar panels (see electric power subsystem), considering a 10% margin for the

dry mass, a 2% for the propellant residuals and a 10% for the volume of the tanks according to ESA,

all the results are as follows:

TABLE 10. PROPELLANT MASS BUDGET.

𝑴𝟎 [𝐤𝐠] 𝑴𝒑𝟏 [𝐤𝐠] 𝑴𝒑

𝟐 [𝐤𝐠] 𝑴𝒑𝒕𝒐𝒕𝒂𝒍 [𝐤𝐠] 𝑽𝒑𝒓𝒐𝒑 [𝐮] 𝒕𝒃 [𝒔]

17.28 ± 1.73 4.34 ± 0.09 1.45 ± 0.03 5.79 ± 0.12 3.15 ± 0.32 563.33

ATTITUDE, DETERMINATION AND CONTROL SUBSYSTEM

The attitude determination and control subsystem plays a crucial part in a satellite’s mission, as it is

in charge of stabilizing the satellite and controlling its orientation. It is of most importance in

observation missions like this one, as the payload requires pointing to a specific direction. Another

example of its importance is that thrusters need accurate pointing for orbital maneuvers or to have

the faces of the satellite with solar panels always facing the Sun.

Usually, CubeSats use magnetorquers for the attitude control part, which are devices capable of

changing its magnetic dipole so they can interact with the surroundings magnetic field, but this

means it only works on Earth and planets that have a strong magnetic fields, so it would be of no use

on the Moon and in space.

As magnetorquers can’t be used for this mission, it was decided to use reaction wheels instead.

Reaction wheels are torque motors with high-inertia rotors that can spin in either direction and

provide one-axis control. While the standard orthogonal three reaction wheels configuration (one

for each axis) would be enough to stabilize and control the satellite, a pyramid configuration with


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four reaction wheels was chosen for several reasons, like: better performance, risk reduction and

error reduction. Any form of failure in one or more reaction wheels could result in a loss of

spacecraft’s ability to maintain its position, which may in turn cause a failure of the mission.

The pyramid configuration of reaction wheels is demonstrated in Figure .

FIGURE 11. PYRAMIDAL CONFIGURATION OF FOUR REACTION WHEELS.

The chosen reaction wheel is the GomSpace NanoTorque GSW-600, a compact and high-performance

reaction wheel designed and qualified for an equivalent of 3 years in-orbit operations that can be

purchased in a 4-wheel pyramid setup within a mounting bracket. It offers a continuous torque of

±1.5 mN·m per wheel. The instrument can be seen in the next figure.

FIGURE 12. NANOTORQUE GSW-600 4-WHEEL PYRAMID SETUP.

Besides the reaction wheels, the satellite will use a Sun sensor and a star tracker (see APPLICABLE

AND REFERENCE DOCUMENTS) for the attitude determination part.


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COMMUNICATIONS SUBSYSTEM AND GROUND SEGMENT

The communications subsystem is of the essence in any space mission. If one gathers data but is

unable to transmit it to the ground where it can be studied and interpreted, the spacecraft is

worthless.

So, one needs to ensure the telecommunication between the spacecraft and the ground stations, in

this case ESA ground stations as it is specified in the mission requirements.

The communications subsystem must ensure that the operation goes both ways, from the ground

station to the spacecraft and vice versa. That is because the spacecraft needs to transmit to the

ground station not only the scientific data from the mission, but its actual state so if corrections are

needed, the ground station can apply them and send them to the spacecraft.

To process these instructions the spacecraft has the OBDH at its core, which manage these data while

redirecting power from different parts of the spacecraft and other tasks.

Now, the process of defining the communication system of this particular mission will be described.

First, a decision will be made to choose a frequency band to send and receive data. Since this mission

will be carried out at a distance of 384,000 km from Earth, high frequency bands are preferred so

that the free space losses do not damp affect the electromagnetic wave much. Taking this into

account and the requirements for this Nanostar mission, the selected band will be an X Band.

The frequency band selected gives a downlink frequency of 7.5 GHz. In order to meet a BER of 10-5

without sacrificing much power in the process, a QPSK modulation and convolutional coding will be

selected. That gives a requirement of Eb/No = 4.4 dB, which will be increased by 3 dB as suggested in

Nanostar methodology to establish a wide enough margin to ensure the fulfillment of the mission.

As was described in the orbit section, most of the time the spacecraft will be at a lunar distance from

the Earth, hence all the calculations will be done supposing the spacecraft is at that distance during

all the mission. Next, the calculations made to be able to send the scientific and state data will be

carried out, using similar values as in other missions and typical values of ESA ground stations.

Supposing an efficiency of 80% for the ground stations antennas and a 34 m diameter, the gain of

the ground station is defined as follows:

𝐺𝑡 =𝐷2

(𝑐𝑓)2 𝜂 =

352

(3 × 108

7.5 × 109)2 0.8 = 57.6 dB

Now, calculating the received power in the Earth supposing a nominal transmitted power of 10 W

(this value will change depending on the mode of operation):

𝑃𝑟 = 𝑃𝑡 (𝜆

4𝜋𝑑)2 𝐺𝑡𝐺𝑟𝐿

= 2.49 · 10−16 W


22

where 𝑑 is 384,000 km, 𝐺𝑟 is 3 dB and 𝐿 is the link losses and a value of 5 dB will be assumed to be

conservative in the calculations.

Assuming a noise figure of 3 dB for the ground antenna, the noise temperature can be calculated as:

𝑇𝑒 = (𝑓 − 1)𝑇0 = (100.3 − 1)290 = 288.6 K

The noise can be defined as:

𝑁𝑜 = 𝑘𝑇𝑒 = 3.98 · 10−21

where k is the Boltzmann constant. Now, one can estimate the bit rate using the Eb/No set at the

beginning of the calculations and the relationship:

Using this, one gets a Rb of 14.3 kbps while doing flybys around the Moon.

𝐸𝑏 = 𝑃𝑟/𝑅𝑏

Finally, times will be computed to check if there is enough time to send the data to the ground station

between each flyby. Supposing that the scientific data is 12 MB for the camera and the spectrometer

(a minimum of 10MB is required) the time to send the data is 2 h. Given that the period of the orbit

is 24 h, that is enough time to send the data to the ESA ground stations. Also, 100 kB of data will be

taken into account for checking spacecraft status, or sending corrections maneuvers. The time to

transmit this data is 55.8 s, so if any action is required the time between the actions is short.

It should be noted that these mission times explained can be shortened if using another spacecraft

mode where more power is redirected to the communication subsystem.

Also, the system will consist of two patch cubesat antennas for X Band, to ensure that throughout

all the mission at least one of the two antennas is functioning, and the critical failure of one does not

interrupt the mission because one has the other one has a backup.

The patch antenna used will be shown in the next figure:


23

FIGURE 13. CUBESAT PATCH ANTENNA X BAND.

The datasheet of the selected antenna can be found in APPLICABLE AND REFERENCE DOCUMENTS.

COMMAND AND DATA HANDLING SUBSYSTEM

On-board Computer for Open Eyes mission will be selected based on the needs on the mission and

using well known technology for nanosatellites.

In particular, the OBC selected is NanoMind A3200. Its features can be consulted in the reference. It

requires a space of 10x10 cm approximately with a height of 1 cm.

ELECTRIC POWER SUBSYSTEM

It may be distinguished two cases: transfer orbit to the moon and lunar orbit. In the first case, it is

considered that there is no eclipse along the whole trajectory. Besides, most subsystems would not

be demanding great power, just the necessary to keep tracking the satellite and to introduce minor

corrections. In conclusion, the critical case should be while orbiting the moon.

In order to seize the electrical power subsystem (EPS) and all its components, this is mainly solar cells

and arrays, the battery and the components which distribute the power to other subsystems and

payloads, the team has begun with an estimation of the power budget that typically allocates to each

subsystem, Table 11.


24

TABLE 11. MASS AND POWER BUDGET ESTIMATION. SOURCE: Y.-K. CHANG ET AL. / ACTA ASTRONAUTICA 61 (2007)

676 – 690.

Subsystem Mass [%] Power [%]

ADCS

Passive 0.0217 · 𝑀 + 5.8198 −0.0152 · 𝑃 + 8.858 Active −0.0142 · 𝑀 + 13.748 0.0036 · 𝑃 + 18.304

C&DH −0.0079 · 𝑀 + 5.5627 −0.03 · 𝑃 + 15.39 TT&C −0.0103 · 𝑀 + 6.5935 0.0456 · 𝑃 + 25.583 TCS −0.0002 · 𝑀2 + 0.0498 · 𝑀 + 0.4785 0.0067 · 𝑃 + 0.7862 EPS −0.0084 · 𝑀 + 18.237 -

Structure −0.01 · 𝑀 + 31.079 -

Being 𝑀 [kg] the total mass of the satellite, and 𝑃 [W] the average power of the satellite.

Propulsion subsystem does not appear in Table 11 on purpose, as the duty of this section is to

characterize the power, and electric propulsion is none of the business in the design of Open Eyes’

satellite. So apparently, percentages in the table must be corrected so altogether can reach a higher

amount of the total power budget, as chemical propulsion do not demand as much power as electric.

However, the camera payload and other potential power loads are not included.

As TT&C (Telemetry, Tracking & Communications) was one of the earliest inputs in the design of the

satellite, 𝑃𝑇𝑇&𝐶 = 6 W, average power 𝑃 was then deduced from equations in Table 11¡Error! No se

encuentra el origen de la referencia.. It has been supposed a demand of 11 W for all the payloads.

Besides, according to ESA-Margin philosophy for science assessment studies, the total power budget

of the spacecraft shall include an ESA system level power margin of at least 20% of the nominal

power requirements of the spacecraft.

In the end it results in a system of equations as the following, where ∑𝑃𝑐𝑜𝑟𝑟𝑒𝑐𝑡.𝑛𝑜𝑛−𝑝𝑎𝑦.[%] and 𝑃𝑡𝑜𝑡𝑎𝑙[W]

have been highlighted in bold to point at the two unknowns.

{

𝑃𝑐𝑜𝑟𝑟𝑒𝑐𝑡.

𝑛𝑜𝑛−𝑝𝑎𝑦.[%] = 𝑃𝑡𝑎𝑏𝑙𝑒 1𝑛𝑜𝑛−𝑝𝑎𝑦.[%] ·

∑𝑷𝒄𝒐𝒓𝒓𝒆𝒄𝒕.𝒏𝒐𝒏−𝒑𝒂𝒚.[%]

∑𝑃𝑡𝑎𝑏𝑙𝑒 1𝑛𝑜𝑛−𝑝𝑎𝑦.[%]

𝑃𝑐𝑜𝑟𝑟𝑒𝑐𝑡.𝑛𝑜𝑛−𝑝𝑎𝑦.[W] = 𝑃𝑐𝑜𝑟𝑟𝑒𝑐𝑡.

𝑛𝑜𝑛−𝑝𝑎𝑦.[%] · 𝑷𝒕𝒐𝒕𝒂𝒍[𝐖]

𝑷𝒕𝒐𝒕𝒂𝒍[𝐖] = 1.2 · (∑𝑃𝑐𝑜𝑟𝑟𝑒𝑐𝑡.𝑛𝑜𝑛−𝑝𝑎𝑦.[W] +∑𝑃

𝑝𝑎𝑦𝑙𝑜𝑎𝑑[W])

∑𝑷𝒄𝒐𝒓𝒓𝒆𝒄𝒕.𝒏𝒐𝒏−𝒑𝒂𝒚.[%] +∑𝑃

𝑝𝑎𝑦𝑙𝑜𝑎𝑑[%] = 100%

∑𝑃𝑝𝑎𝑦𝑙𝑜𝑎𝑑

[%] =∑𝑃

𝑝𝑎𝑦𝑙𝑜𝑎𝑑[W]

𝑷𝒕𝒐𝒕𝒂𝒍[𝐖]=

11 W

𝑷𝒕𝒐𝒕𝒂𝒍[𝐖]

This system has no solution, but both errors in terms of total power and the sum of all non-payloads

in percentage can be minimized as it is shown in Figure .


25

FIGURE 14. ERROR IN POWER AND NON-PAYLOADS SUM PERCENTAGE.

Trying to balance both errors, the average power of the satellite hereafter will be 30 W and all non-

payloads subsystems will sum 55% of the overall power, which means for both magnitudes a 10%

of error. To reassure the goodness in the results, it is checked that is slightly over the initial 6 W (see

Table 12), which stands at the side of safety.

TABLE 12. POWER OF ALL SUBSYSTEMS.

Subsystem/ payload

Power [%] Corrected Power [%]

Power [W] Power [W] (with 20 % margin)

ADCS

Passive 8.40 6.71 2.01 2.42

Active 18.41 14.71 4.41 5.30

C&DH 14.49 11.58 3.47 4.17

TT&C 26.95 21.53 6.46 7.75

TCS 0.59 0.47 0.14 0.17

Camera - - 5.00 6.00

Ext. Payload 1 - - 3.00 3.60

Ext. Payload 2 - - 3.00 3.60

∑ 68.84 55.00 27.50 33.00


26

It is usually multiplied the average power by 2 or 3 to obtain peak power requirements for attitude

control, payload, thermal, and EPS (when charging the batteries). This can be compared with results

in Operational Modes section in this document.

POWER SOURCE

Because photovoltaic solar cells are the most common power source in most missions and are famous

for being well-known and reliable, Open Eyes will focus on this technology. The starting point of the

design is defining the mission life and the average power requirement. The former one has been

estimated in mission analysis and orbit determination, shorter than a year at most, whereas the latter

has been already obtained hereabove.

Solar cells efficiencies range from 15% to 30% in space applications. It has been guessed a

multijunction solar cell type with an efficiency of 22%, according to production references. Next, it

must be considered that an assembled solar array is less efficient than single cells due to design

inefficiencies, shadowing and temperature variations, collectively referred as inherent degradation,

𝐼𝑑. A typical value for this coefficient is 0.77 although it might vary from 0.49 to 0.88. Also, solar cells

voltage during operation must be guessed at all probability different to the maximum power

extraction point.

Power at the beginning of life, per unit of area responds to

𝑃𝐵𝑂𝐿 = 𝐾𝑆𝑈𝑁𝜂𝑐𝑒𝑙𝑙𝐼𝑑 cos(𝜃)

where 𝐾𝑆𝑈𝑁 is the solar irradiance (1,313 W/m2), which is taken as constant as the value for Earth

in summer, although apogee radius is around 15,000 km from the moon and may imply little

variation in 𝐾𝑆𝑈𝑁. In addition, cos(𝜃) introduce the loss associated to the obliquity of the solar rays

in relation to the normal direction to the cell. It is supposed an incidence angle of 50° in average

during the orbit. Then, 𝑃𝐵𝑂𝐿 achieves 142.97 W/m2.

Next, it is needed to estimate the factors that degrade the solar array’s performance during the

mission. According to ESA-Margin philosophy for science assessment studies, solar arrays and

batteries shall be sized to provide the spacecraft required power at End of Life (EOL). Life

degradation, 𝐿𝑑, occurs because of thermal cycling in and out of eclipses, micrometeoroid strikes,

plume impingement from thrusters, and material outgassing for the duration of the mission. In

general, gallium-arsenide cells power production can decrease by as much as 2.75% per year in LEO,

but a satellite located in a moon orbit can suffer greater degradation due to a higher rate of

micrometeoroid strikes.

The actual lifetime degradation in LEO can be estimated using,

𝐿𝑑 = (1 − 𝑐)𝑡𝑓


27

where c is the degradation per year and 𝑡𝑓 is the satellite lifetime in years. In Open Eyes mission, for

a 3% degradation and a lifetime duration of a year, 𝐿𝑑 is 0.97, overestimated to 0.95 because of it is

operating around the moon.

The array performance per unit of area at the end of life is,

𝑃𝐸𝑂𝐿 = 𝐿𝑑𝑃𝐵𝑂𝐿 = 0.95 · 142.97 W/m2 = 135.82 W/m2

Finally, the solar array area 𝐴𝑠𝑎 required to support the spacecraft power requirements, supposed

that cells once positioned take up 70% of the total area, is given by

𝐴𝑠𝑎 =1

𝑓𝑎𝑟𝑒𝑎 𝑃𝑡𝑜𝑡𝑎𝑙𝑃𝐸𝑂𝐿

=1

0.7

30 W

135.82 W/m2≈ 0.32 m2

It must be taken into account that other surfaces take part of the total exterior face area, such as

antennas and the radiator, and also the camera aperture, all of them going up to 0.05 m2.

It is observed in Figure the total area availability for different kinds of cubesats layouts. If it is

expected to provide all the power through solar cells mounted on the satellite exterior faces, as it is

needed an area of about 0.32 m2 plus 0.05 m2 , the smallest configuration that allow for this to

happen begins with 16U (2x2x4 U). 18U (2x3x3 U) would be a good choice too, as it allows more

unoccupied space inside de satellite, and facilitates the inertias.

FIGURE 15. TOTAL FACE AREA AVAILABILITY IN CUBESATS.


28

FIGURE 16. 16U (LEFT), 18U (CENTER) AND 12U (RIGHT).

Nevertheless, actual volumes of all elements do not overtake 10U. In conclusion, in order to leverage

the space inside the spacecraft, it is proposed a configuration of 12U, keeping in mind to reserve

some free volume to attract possible partners to load bigger payloads. Therefore, the satellite must

be equipped with deployable solar arrays. The dimensions are easily calculated as follows:

𝐴𝑎𝑣𝑎𝑖𝑙 = 𝐴12𝑈 − 0.05 m2 = 0.27 m2

𝐴𝑝𝑎𝑛𝑒𝑙 = 𝐴𝑠𝑎 − 𝐴𝑎𝑣𝑎𝑖𝑙 = 0.05 m2

Looking into IDM-CIC catalogue, generic 3U solar array fulfill an area of about 82 x 327 mm2, that is

0.027 m2, so three units of this element must be installed.

Thinking of how to implement the solar cells in the satellite, there is another approach instead of

covering nearly all free faces with them. If remembered, an incidence angle of 50° in average during

the orbit is actually too bad and does not take full benefit. Instead, it is proposed a solar array

configuration as in Figure, where two big faces of the satellite deploy a big solar rotating array

structure each. Also, some solar cells keep laying on the main structure of the satellite to produce

some extra energy in case of abnormal attitude or in transient phases of the mission. Yet, ADCS can

control the right attitude of the spacecraft body so solar incidence stays in the best range. The solar

array area needed for 20° of incidence results in 0.22 m2, 8 x 3U solar array in other words (see

¡Error! No se encuentra el origen de la referencia. left).

Other kind of solar array configuration would be as in Figure right. However, as this configuration

seems less complex because it simply deploys the arrays through hinges but do not rotate, to keep a

good incidence angle as in previous configuration, the satellite would do have to rotate instead,

reducing the observation window for the camera payload through the perigee.


29

FIGURE 17. SOURCE: HTTPS://SPACE.SKYROCKET.DE/DOC_SDAT/LUNAH-MAP.HTM

FIGURE 18. SOLAR ARRAY CONFIGURATIONS.

In conclusion, two rotating deployable solar arrays guarantee attitude flexibility and make it easier

for the camera to develop the right pointing throughout the perigee. Besides, the main body would

have practically all faces free to put more solar cells in case of need in further design analysis.

It has been considered the over cost in the mass budget referring to the same generic element in the

catalogue to redo the calculations of the propellant needed therein. The extra volume of the tanks is

maintained inside the limits available in the satellite.

POWER STORAGE

As previously told, the peak power can rise much above the nominal power. It is assumed a peak

power of 120 W, which is 4 times nominal power, so the battery must supply 90 W in difference,

assuming then that the peak power period does not coincide in time with eclipse. Time of eclipse

once in the moon orbit can extend up to 40 min.

Theoretically, two 3S1P battery modules connected in series, of three Samsung battery cells of

2850 mA ∙ h each, can provide voltage and current enough for all cases. The rate of discharge during

the eclipse must be low enough to keep power distribution to all the payloads that may be working

during the eclipse. Also, it is assumed that the peak power is not much prolongated in time, so the

depth of discharge does not decay below the minimum advised to keep the battery safe.


30

POWER DISTRIBUTION

A spacecraft power distribution system consists of cabling, fault protection, and converters to adapt

voltage to each payload. In selecting a type of power distribution, we focus on keeping power losses

and mass at a minimum while attending to survivability, cost, reliability, and power quality.

In this paper, Open Eyes has no intention to describe this matter so far because the level of detail of

the subsystem itself is not enough. Indeed, power transmission with the 2 batteries in series yet

commented for a 90 W peak power implies currents below 10 A, so small gauges are required to

transport such a current.

Just to know, the harness or wiring that interconnects the spacecraft subsystems is a large part (10 −

25%) of the electrical power system mass. Harness must be designed as short as possible to reduce

voltage drops and to reduce the total spacecraft mass.

MECHANICAL DESIGN AND STRUCTURE

Cubesat structures are typically composed of a rigid skeleton, see ¡Error! No se encuentra el origen

de la referencia.9, top and bottom panels prepared for satellite separation at the final phase of the

launch, interior panels to support shear stresses, closure panels and multiple trays to accommodate

the subsystems and the payloads.

According to ECSS-Q-ST-70-36C, among the alloys with medium-high resistance to stress‐corrosion

cracking accepted in space applications are Al-7075 or Al-6061, 𝜌𝐴𝑙7075 = 2810 kg/m3 and

𝜌𝐴𝑙6061 = 2700 kg/m3.

FIGURE 19. 12U SKELETON STRUCTURE. SOURCE: HTTPS://WWW.ISISPACE.NL/.

The only mechanisms aboard the spacecraft shall be the deployable system to position the outer

solar arrays. Moreover, rotating mechanism are far complex compared to hinge-mechanism arrays,


31

and they are in need of an intelligence to control the rotation. This intelligence will be granted via

software with the information gathered by the star tracker and sun tracker.

Now, the mass budget of the mission will be presented. It is done with the aid of the IDM-CIC

software, which enabled to elaborate the full configuration of the satellite by using sketch-up with

parts already defined typical from CubeSats standards.

FIGURE 20. OPEN EYES DEPLOYED PANELS

CONFIGURATION (1/2).

FIGURE 22. OPEN EYES FOLDED PANELS CONFIGURATION

(1/2).

FIGURE 21. OPEN EYES DEPLOYED PANELS

CONFIGURATION (2/2).

FIGURE 23. OPEN EYES FOLDED PANELS CONFIGURATION

(2/2).


32

In Figures 20 to 23, several mission configurations of the satellite are presented. This arrangement

has been made inside the IDM-CIC tool and is visualized in the IDM-View tool. One can see the

configuration with the deployed solar panels and the configuration with the solar panels folded and

next to the closure lateral panels, as it will appear at the very early stage of the mission.

Having this configuration, the mass estimations made by our team were introduced in the

calculations including the margins recommended by the ESA Agency, as can be checked in the

references.

Below, in Figure 24, one can see the detail of each subsystem involved in the mission with its

corresponding estimated mass. One may note that the total weight calculated here includes the solar

panels, which have a mass of 1.3 kg approximately. This makes this design match with the CubeSat

standards of maximum weights, even by overestimating some aspects of the design as it has been

shown in this report.

FIGURE 24. MASS BUDGET OF THE OPEN EYES MISSION.


33

THERMAL CONTROL SUBSYSTEM

The thermal control system of a spacecraft is essential to guarantee the success and optimum

performance of the mission as it allows to maintain the system’s temperatures within a set range

during its lifetime. As not all instruments in the satellite can operate at the same temperatures, if

one were to encounter a temperature outside of that range it could be damaged or its performance

could be severely reduced, affecting the mission.

The range of temperatures our satellite must operate in is set by the requirements of the mission,

which state that the satellite’s temperature must be between −10℃ and 30℃ when operating and

between −20℃ and 40℃ when inactive.

The desired temperatures are achieved by balancing the flow of heat energy across the satellite

interfaces. For that both the power dissipated by the instruments and the thermal environment of

the orbit must be considered.

The usual approach taken in the design of the thermal control of a satellite is through conventional

passive techniques, such as selective placement of power dissipating components, application of

surface finishes, used of certain materials, and regulation of conductive heat paths. In cases where

the passive approach is not enough to keep the temperatures of the spacecraft or certain

components within the set range, heaters can be placed near the instruments with more critical

temperature limits so they can meet the requirements.

The assumptions made for the initial analysis were that one of the faces of the satellite would always

be facing the Sun, that the power dissipated by the instruments is a 15% of the power consumption

and that the configuration of the satellite is a 12U.

The energy balance equation is as follows:

𝑄𝑖𝑛 = 𝑄𝑜𝑢𝑡

𝑄𝑖𝑛 = 𝑄𝑆 + 𝑄𝑎 + 𝑄𝑝 +𝑊

𝑄𝑜𝑢𝑡 = 𝜎𝑇4 𝜀𝐴

where 𝑄𝑖𝑛 is all the heat it receives, conformed by the solar input (𝑄𝑠), the solar energy reflected on

the planet/moon (albedo input) (𝑄𝑎), and the power dissipated by the internal components (W). 𝑄𝑜𝑢𝑡

it’s what the satellite irradiates though all its surfaces.

𝑄𝑠 = 𝐸𝐴𝑝𝛼

𝑄𝑎 = 𝜌𝑝𝛼𝐸𝐴𝐹𝑠𝑝

𝑄𝑝 = 𝜀𝐴𝜎𝑇𝑝4𝜀𝑝𝐹𝑠𝑝

where E is the solar constant (1366 W/m), 𝐴𝑝 is the projected area of the satellite towards the Sun,

𝛼 is the absorptivity of the surface material, 𝜌𝑝 is the planet/moon albedo factor, 𝐹𝑠𝑝 is the view

factor, 𝐴 is the total radiating surface of the satellite , 𝜀 is the emissivity of the surface material


34

(which varies from solar cells to MLI), 𝜀𝑝 is the emissivity of the planet, 𝑇𝑝 is the reference

temperature of the planet , 𝜎 is the Stefan-Boltzmann constant (5,67 ⋅ 10−8W

m2 𝐾4) and T is the

absolute temperature in Kelvin.

The material chosen for the surface was MLI, which has an absorptivity of 0.3 and an emissivity of 0.6

and provides isolation from external environment. Given that there are solar panels in the faces,

those have an absorptivity and emissivity value of 0.75 which must also be taken it into account. So,

for the external heat sources that encounters with the solar cells, the absorptivity used will be:

𝛼𝑡ℎ = 𝛼𝑝𝑎𝑛𝑒𝑙 − 𝜇𝐹𝑜

being 𝜇 the efficiency of the solar panels and 𝐹𝑜 the occupation factor of the solar panels on the

satellite. The efficiency was assumed to be 0.3 and the occupational factor 0.65.

Two cases were studied during the analysis, during the trip, far from the influence of the Earth or the

Moon, and during the Moon’s flyby. During the first case the satellite would only receive heat from

the Sun, while in the second one it would also receive the solar energy reflected on the Moon and

the heat that the Moon irradiates.

The satellite consumes approximately 30 W of power, so assuming a 15% of losses, the satellite

would dissipate 5 W. As this is just an estimation, calculations were also made for 40% of losses.

Without a heater, the results showed that the satellite would meet the requirements by small margin

when the dissipated power is 5 W. To be sure, a small 5 W heater was added to the calculations. The

results obtained can be seen in the following table.

TABLE 13. TEMPERATURE OF THE SATELLITE IN VARIOUS CASES.

Case In space During flyby

No heater, 15% dissipated −7 ºC 20 ºC

No heater, 40% dissipated 3 ºC 28 ºC

5W heater, 15% dissipated −2 ºC 24 ºC

The results showed the small heater for certain instruments would be needed so their temperatures

don’t fall below their limit while in space, allowing for a bigger margin regarding the requirements.

Also, a small radiator will be used to prevent internal components from overheating, since the coating

of those components will be such that they hold a temperature as uniform as possible. The effect of

this radiator in the overall balance is negligible since its size is very small, just for internal balance.

This is only a preliminary analysis of the thermal control, and further analysis and testing would need

to be done in the following phases of the project.


35

ADDITIONAL PAYLOADS

Additional payloads are considered throughout this preliminary analysis. Apart from the main

scientific payload, the team has decided to include an infrared spectrometer.

Its use is motivated to know more about the elements from the lunar surface by analyzing their

wavelengths. This is a good complement to the camera which takes pictures and with the

spectrometer one takes the information that cannot be seen within the visible range.

During the sizing of the satellite, there were cubesat units left blank to let more than one payload

on-board. The objective of allocating this space is to offer the possibility to embark some components

or demonstrators to third-party companies so they are able to test those components in space

environment. A maximum power consumption of 3 W has been allocated for each of these extra-

payloads. This is done to help with the funding of the mission, as we can sell the empty units in our

nanosatellite to different companies to hold a scientific payload of their choice that can match our

design and specifications.

This way the cost of the of the entire mission would be reduced, making it more feasible economically

speaking.


36

RISK ANALYSIS AND MITIGATION

The risk management is essential in every project that seeks to succeed. In the context of the mission

preliminary design, the risks analysis and evaluation has been performed according to the standards

described in the ECSS-M-ST-80C (31July2008) document for risk management.

In this document it is stated the iterative nature of risk management. These iterations comprise four

steps, which are present in the Figure:

Task 5: Decide if the risks may be accepted

Step 4

Monitor, communicate and

accept risks

Step 3

Decide and act

Step 2

Identify and assess the risks

Step 1

Define risk management

implementation requirements

Task 1: Define the risk management policy

Task 2: Prepare the risk management plan

Task 3: Identify risk scenarios

Task 4: Assess the risks

Task 6: Reduce the risks

Task 7: Recommend acceptance

Task 8: Monitor and communicate the risks

Task 9: Submit risks for acceptance. (Return

to Task 6 for risks not accepted)

R

I

S

K

M

A

N

A

G

E

M

E

N

T

C

Y

C

L

E

FIGURE 25. THE FOUR STEPS AND THEIR TASKS OF EVERY RISK MANAGEMENT CYCLE.

Source: European Space Agency

This is the methodology chosen by the team, and these steps will be followed and presented in this

document. Nevertheless, since the aim of this preliminary design is allocated among the phases 0-A,

there will be no further iterations of this risk management cycle apart from the initial one.

RISK MANAGEMENT POLICY

On this section, the risk management policy will be defined.

In order to assess and classify the risks, a scoring scheme needs to be defined. The scoring scheme

will consist on two different parameters, one for severity and one for likelihood of the risks, and the

definition of each parameter is explained on Table 14 and Table 15.


37

TABLE 14. SEVERITY-OF-CONSEQUENCE SCORING SCHEME.

Score Severity Severity of consequence: impact on feasibility

5 Catastrophic Critical impact on many subsystems. Critical impact on mission objectives.

4 Critical Major impact on primary subsystem and/or Payload. Major impact on mission.

3 Major Major impact on secondary subsystem and/or minimal impact on primary subsystem. Minor impact on mission.

2 Significant Minimal impact on secondary subsystem.

1 Negligible Minimal or no impact.

TABLE 15. LIKELIHOOD-OF-OCCURENCE SCORING SCHEME.

Score Likelihood Likelihood of occurrence

E Maximum Certain to occur, will occur one or more times per project

D High Will occur frequently, about 1 in 10 projects

C Medium Will occur sometimes, about 1 in 100 projects

B Low Will seldom occur, about 1 in 1000 projects

A Minimum Will almost never occur, 1 of 10 000 or more projects

The combination of these two parameters will be used to rate the magnitude of the risk, and

therefore the need to act or not on a determined risk. The magnitude has been defined as shown on

Figure and Table 16.

FIGURE 26. RISKS MAGNITUDE ACCORDING TO SEVERITY AND LIKELIHOOD.


38

TABLE 16. RISKS INDEX AND MAGNITUDE.

Risk index Risk

magnitude

Proposed actions

E4, E5, D5 Very High risk Unacceptable risk: implement new team process or change baseline –

seek project management attention at appropriate high management

level as defined in the risk management plan.

E3, D4, C5 High risk Unacceptable risk: see above.

E2, D3, C4, B5 Medium risk Unacceptable risk: aggressively manage, consider alternative team

process or baseline – seek attention at appropriate management level as

defined in the risk management plan.

E1, D1, D2,

C2, C3, B3,

B4, A5

Low risk Acceptable risk: control, monitor – seek responsible work package

management attention.

C1, B1, A1,

B2, A2, A3,

A4

Very Low risk Acceptable risk: see above.

RISKS IDENTIFICATION AND ASSESSMENT

Once the Risks policy is defined, the risks can be adequately assessed after being identified. The list

of risks identified, and their rating is shown on Table 17.

TABLE 17: RANKED RISKS LOG.

Project: NANOSTAR Organization: Open Eyes Date: 05/005/2019

Issue: v1.0

Rank No. Risk scenario

title

Red Yellow Green Risk

domain

Actions and status

A5 R-01 Launch Vehicle

Failure

X All No action possible.

B4 R-02 Instrument

breakdown

during launch

X All Complete test and verification of

individual and assembly normal

modes.

C4 R-03 Deviation

during lunar

transference

X Orbit Periodic orbit check and small

engines for trajectory correction

needed.

A4 R-04 Solar panel

shortcut or

malfunction

X Power Battery needed to support solar

panels.

C4 R-05 Failure in PL

activation

X Payload Possibility to do software check or

update remotely.


39

C2 R-06 Transmission

impossible due

to bad

positioning

X Comms Possibility to store data until a

communication window opens.

B5 R-07 Failure in

primary engine

ignition

X Propulsion

B5 R-08 Failure in

reaction wheel

X ADCS Redundant reaction wheels.

A4 R-09 Space debris or

micro-

meteoroids

impact

X All Monitor and control space debris.

B4 R-10 Failure in OBC

software

X Data

Handling

Possibility to upgrade software

from Ground Station

B5 R-11 Failure in solar

panel

deployment

system

X Power Deployment system testing on

Earth.

Notes

(*) Mark box as appropriate for the value of “R” (Risk index) from the risk register, according to the criteria defined in

the risk management policy.

(**) Indicate risk domain (e.g. technical, cost or schedule).


40

CONCLUSIONS

The purpose of the Open Eyes team members is to predesign a small satellite fly-by mission to the

Moon, with the objective of taking a few pictures of the lunar soil during the periselenium pass.

Using GMAT to study the best possible orbit design, the final orbit chosen consists of a Hohmann

trajectory, starting in the perigee of the GTO and ending as close as possible to the Moon. At that

point, the satellite is given an impulse so it reaches the 80km of altitude during its fly-by. Instead of

doing more fly-by’s, the team chose to maintain the satellite orbiting around the Moon (giving the

satellite a second impulse), to obtain more scientific data. A choice was made to include a

spectrometer, to not only have pictures of the surface, but information of the elements that conform

it.

The chosen propulsion system for the mission was the GR-22, which provides the necessary thrust to

obtain low combustion times and uses a “greener” fuel. Initially, the team’s idea was to use the

instability of the L1 point of the Earth-Moon system to minimize the impulse for the fly-by, but the

idea was discarded due to its difficulty and because the definitive design was thought to be better

for scientific purposes.

The satellite will consist of a 12U configuration with two rotating deployable solar arrays of 0.11m2

each that will always be facing the Sun. The material surrounding the satellite will be MLI, and a small

heater and radiator would be used to keep certain instruments warm enough during the Hohmann

trajectory and prevent them from overheating, respectively.

To stabilize the satellite and make sure the camera is always pointing towards the Moon, the satellite

will use four reaction wheels in pyramidal configuration.

For communications, the satellite will use two patch cubesat antennas for X Band, to ensure contact

throughout the mission. Besides the camera and spectrometer, the satellite will also carry some

additional payloads. These units will be sold to different enterprises who might have a specific

interest and will help funding the entire mission.


41

APPENDICES

Appendices will be used to attach additional materials that the team may find relevant to understand

the report.

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