AERO2705 Space Engineering 1

Assignment 3 1. Scientific package

The aim of the mission it is to study the composition of the comet in order to try and verify the existence

of the Oort cloud. The hypothesis is that most comets today come from the Oort cloud, and so we shall

determine the comet properties and materials using the ALICE spectrometer. We will obtain these

results and compare them to previous studies on comets and see if comet composition and properties

are similar. If so, it may lead to suggest that they came from the source, the hypothetical Oort cloud.

ALICE is a unique UV Spectrometer which allows UV Mapping within the sensors Field of Vision (FOV),

used originally at the NASA´s Rosetta Orbiter. The sensor works by allowing UV light to enter through the

Aperture door which is then reflected off to an Entrance Slit via Off Axis Parabola mirror (OAP mirror).

From the slit, the light hits the toroidal holographic grating which is then dispersed onto a micro channel

plate detector (MCP) that uses a double delay readout scheme to save the image. By having multiple

shots we can determine the elemental emission of comet with respect to time.

The Sensor itself is actually controlled by an independent microcontroller (SA 3865), which can of cause

be removed and attached to our own microcontroller if so desired. However, leaving the microcontroller

does have some benefits. The biggest advantage is that the sensing is not multiplexed with other sensors

aboard the satellite, and its RAM can be purely dedicated to the UV sensing. The downside is the

difficulty in controlling and integrating two microcontrollers (so essentially, a decision has to be made

between faster process time, and simpler integration.)

As a launch vehicle we shall choose the Atlas IIA due to its successful program since the 1950´s. The

flying path is to launch the satellite from the vehicle after arriving at comet’s orbit. It is intended a launch

from Cape Canaveral launch site at Florida, USA, to a 28.5° inclination circular LEO orbit and then make

the necessary orbit changes to launch the satellite after arriving the comet´s orbit.

2. Satellite subsystems

Attitude control In order to achieve optimal attitude control, we must a combination of both an attitude sensor and an

attitude actuation unit.

For this satellite subsystem, a fibre optic gyroscope (FOG) shall be used for attitude determination. A

FOG, performs the same function as a typical mechanical gyroscope, but instead works through the

Sagnac effect (a beam travelling against a rotation will experience a slightly shorter path), rather than

gyroscopic measurements of momentum wheels. [3]

The FOG that shall be used is the Astrix® 200 from Astrium. This inertial measurement unit will provide a

three axis measurement of the satellite’s current rotational rate with incredible accuracy (0.001 arc

seconds). The mass of the Astrix® 200 is 10kg, and only consumes a few mW of power. Moreover this

specific unit has been designed to be separated from processing electronics which will enable it to be

thermally regulated and avoid thermo-elastic effects on the fibre optic cables. [4]

The attitude actuation consists of two sections, a fine control and a slew control unit. For slew control,

small hydrazine thrusters shall be placed at critical points around the satellite in order to maximise the

moment arm of thrusters relative to the centre of gravity (cg) of the satellite.

Moreover the hydrazine thrusters shall be placed with symmetry. This is so that instead of having one

thruster to produce a net moment, two may work in collaboration to produce the same net moment

with less fuel and power consumed. The hydrazine thrusters to be used for the satellite shall be 1N

thrusters developed by Astrium [5].

For fine attitude control, we shall use a system of gyroscopic reaction wheels. The particular final

product is still to be decided upon. However, the purchased gyroscopes must allow for three-axis

stabilization. There must be three light weight gyroscopes positioned perpendicular to one another, with

a fourth one positioned skew to the remaining three. This fourth wheel shall function to remove spin

saturation from any other, and also serve as a backup reaction wheel, should any of the others fail

Orbit Transfer and Position Control The satellite will need a main engine for large changes in velocity. We need to therefore use a large

thruster suited to the job, with the best option being a monopropellant 400N [6] Engine for the satellite.

The thruster shall be designed for both long term steady state and pulse mode operation.

For fuel savings we need to use a propellant with a high specific impulse, thus Hydrazine (once again) is a

good option. The structure of the thrusters must also be designed to protect them from damage and

serve as a heat barrier for propellant valves. For the storage of the propellant we are going to use a BT 0-

01 [7] platform Hydrazine Bladder Tank, so the system will not need an external pressurized system to

make the fuel flow to the thrusters since the propellant tank have a mechanism that keep contracting

the internal volume as the fuel is consumed. The tank has two operational modes: either blow-down

mode, or pressure regulated mode.

Thermal Equilibrium Balancing

Since the satellite is working in a vacuum environment, we need to maintain the internal temperature to

make all the components work in the operating temperature, avoiding the temperature to reach the

minimum and maximum of the survival temperature to make an optimum fluctuation of temperature as

necessary for the components.

For this, we must use some components to avoid heat transfers between the satellite and space since we

have the influence of external agents, as the sun. For protect the satellite from thermal radiation we

must use a Multi-Layer Insulation of aluminum and kapton™ [8] to reflect solar radiation and also

prevent heat to enter the satellite structure. Also we can add louvers in the external structure, since they

are thermally activated shutters that regulates the structural and electronic thermal environment during

the flight, they sense the temperature of a space radiator and react to control temperature and are

capable of operation within an external environment range of -85°C to 120°C over 30,000 cycles with no

degradation in performance [9]. Internally on the satellite we need also to add a cryogenic multi-layer

insulation [10] at the propellant tank to avoid heat transfer with other components, since the propellant

is at a low temperature and we want to maintain an optimum temperature on the satellite interior.

Power generation, storage and regulation

Power will need to be generated to support the onboard electronics of the satellite to carry out the

mission. This energy will be generated by solar panels made by Emcore and this will have to generate

enough power to supply energy to:

Communication systems

Central control systems

Science instruments

Positioning and attitude control electronics

The satellite will fly further away from the sun therefore the mission duration is estimated to be the

window between satellite meeting the comet and till they pass beyond the orbit of Mars, at which point

solar panels are deemed to be insufficient to power the entire system.

Given that the solar panels will be at 29.5% minimum average efficiency [17], and intensity of light at

Mars is 589.2 W/m^2 [16] minimum surface area of solar panel needed at average distance of Mars orbit

will be about 3 m^2 (rounding up). This will ensure sufficient power when the satellite and comet are

closer to the sun. This is necessary to account for damaged solar cells and decrease in efficiency over

time. This electrical energy will be stored into Li-ion (or Lithium Polymer) battery so all the power output

is from the battery to regulate voltage supplied.

Communications

The factors that have been considered for the communications subsystem include:

Data transmission rate

Power consumption

Directional considerations of transmissions

Interference due to Earth’s atmosphere.

The transponder that will be used is the X/X Deep Space Transponder, produced by Thales Alenia Space.

X-band frequency signals will be transmitted, as higher data rates can be transmitted, although they will

require more power. The transponder’s uplink frequency is in the range 7145-7235 MHz, and the

downlink frequency is 8400-8500 MHz.

The Antennas that will be used include four Low Gain Antennas, for receiving and transmitting signals to

and from Earth (uplink and downlink). Four LGAs have been decided upon so that there will always by an

antenna receiving signal from Earth, regardless of the spacecraft’s attitude. In addition, a 0.6m-diameter

High Gain Antenna will be responsible for transmitting data back to Earth (downlink). As the spacecraft

travels towards the comet, the Low Gain Antenna, which is omnidirectional, will be used for tracking the

spacecraft and sending simple correctional thrust commands to the spacecraft. After analysing the

comet’s composition and making measurements using the scientific instruments, the Low Gain Antenna

will be used to receive commands to rotate its attitude so that the High Gain Antenna, which is one-

directional, faces the Earth, at which point, the High Gain Antenna will transmit all the scientific data

(such as images and spectral observations) back to Earth.

Computer Control Systems

Choosing a microcontroller for any application (including satellite) is completely subjective. Each family

of micro controllers are programmed in different ways and have different hardware ports for

multiplexing. This means that programmers will often select their microcontroller purely due to

familiarity. So, for our selection, we are purely going to decide on the microcontroller FAMILY (not the

model itself) based on the following factors:

Voltage/power consumption: very important that none of our component drain the power from

our satellite. The voltage consumption must be as low as possible.

Pin count: Very important as more pins will allow more sensors and actuators to be integrated

Flash memory: this is where all the commands are programmed and stored. While not critical in

having a larger Flash memory, it does lend itself to more flexibility when more sensors and

actuators are integrated in to the controller. Generally having a bigger Flash memory is better.

Clock Speed: this determines the processor’s speed. Like the Flash memory, it is not critical in

having high clock speed, but it does improve our processing speed and number of iteration

sampling. This does give us a more accurate data faster, which is useful in emergency altitude

change, but does little in getting our scientific data since our data is independent of time.

Some possible candidates includes:

RX family

o has Pin count range between 80~100 for a operating voltage range between 1.5~2

o Clock speed is decent with frequency between 32~50 MHz

o Memory is also decent with 256~512KB ad RAM range between 8~64 KB

RL78 family

o has Pin count range between 80~144 for a operating voltage range between 1.5~2

o Clock speed is decent with frequency of 32 MHz

o Memory is also decent with 256~512KB ad RAM range between 8~64 KB

78K Family

o has Pin count range between 80~144 for an operating voltage range between 1.5~2

o Clock speed is 20 MHz

o Memory is decent with 128~512KB ad RAM range between 8~64 KB

Structure and Component Layout

For the satellite bus structure we anticipate that a cube-like structure will be the most efficient (and

easiest) to deal with. However in regards to component layout, all components must be arranged around

the satellite to achieve two aims:

1. Minimize overall mass moment of inertia of the satellite

2. Keep centre of gravity as close to the geometric centre of the satellite as possible (at all times)

In order to achieve this, the most logical arrangement would be to place the heavy mass components

symmetrical about the geometric centre, and have the light weight components spread out along the

edge of the satellite.

3. Life of the Satellite

From when we meet the comet till we exceed the orbit of Mars

The mission length is planned to encompass the trajectory from the launch on Earth, to the comet,

which is approximately 200 days and also an allotted period of time to analyze the comet and transmit

data back to Earth, which is planned to be 14 days. Thus, the total mission period is about 214 days, after

which the spacecraft can be utilized for further missions, if fuel and power still allow for it. Many of the

spacecraft components are designed to last for up to 15 years, far greater than the planned mission. In

addition, the spacecraft may lose effective operation or communication when the spacecraft is moved

beyond the radius of Mars’ orbit (estimated to be some number of days from the start of the mission),

due to the lack of intensity of light on the solar panels and that the batteries are predicted to last only a

few days if operated at 500 W.

4. References

[1] (http://www.esa.int/esaMI/Rosetta/SEMRHF374OD_0.html)

[2] (http://www.esa.int/esaMI/Rosetta/SEMYCF374OD_1.html)

[3] http://www.astrium.eads.net/media/document/astrix-200.pdf

[4] http://www.yorku.ca/bquine/pages/attitude_determination_subsystem.htm

[5] http://cs.astrium.eads.net/sp/spacecraft-propulsion/hydrazine-thrusters/1n-thruster.html

[6] http://cs.astrium.eads.net/sp/spacecraft-propulsion/hydrazine-thrusters/400n-thruster.html

[7] http://cs.astrium.eads.net/sp/spacecraft-propulsion/propellant-tanks/39-litre-hydrazine-bladder-

tank.html

[8] http://www.dunmore.com/pdf/aerospace-brochure.pdf

[9] http://www.orbital.com/NewsInfo/Publications/TSD/Thermal_Louvers_Brochure.pdf

[10]

http://www.ruag.com/space/Products/Satellite_Structures2C_Mechanisms_Mechanical_Equipment/The

rmal_Systems/Cryogenic_Insulation_Coolcat

[11] http://cs.astrium.eads.net/sp/spacecraft-propulsion/hydrazine-thrusters/20n-thruster.html

[12] http://cs.astrium.eads.net/sp/spacecraft-propulsion/propellant-tanks/bladder-tanks.html

[13] http://pdf.directindustry.com/pdf/holroyd-components/kapton/56398-149062.html

[14] http://www.thermacore.com/design-center/documents/datasheets/Loop-Heat-Pipe-Data-Sheet.pdf

[15] http://www.thermacore.com/design-center/documents/datasheets/k-Core-Data-Sheet.pdf

[16] page 8 (http://www.nasa.gov/pdf/163008main_SESE_TeachersGuide__Part_dc3.pdf)

[17] http://www.emcore.com/wp-content/themes/emcore/pdf/ZTJ_datasheet.pdf

[18] http://www.eetimes.com/design/microcontroller-mcu/4235484/How-to-go-about-selecting-a-

microcontroller