Comet and

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Name Signature SID Date

Alexander Bunting Alexander Bunting 311190162 5/10/12

Anthony Zeater Anthony Zeater 311193080 5/10/12

Bastiaan Uytterhoeven-Spark Bastiaan Uytterhoeven-Spark 310220122 5/10/12

Daniel Farhat Daniel Farhat 310205565 5/10/12

Michael Holmes Michael Holmes 311246044 5/10/12

Name SID Allocated

Alexander Bunting 311190162 GROUP MANAGER, Scientific package, Thermal equilibrium, Power requirements

Anthony Zeater 311193080 CPU, Data bus, Logic units, Mass and Inertia

Bastiaan Uytterhoeven-Spark 310220122 Position and attitude control, structure and layout, Mass and Inertia

Daniel Farhat 310205565 Communications Michael Holmes 311246044 Position and Attitude sensing

AERO2705 1 [Type text]

AERO2705 Assignment 4 Comet and Martian Intercept Mission

Alex Bunting (311190162) Anthony Zeater (311193080) Bastiaan Uytterhoeven-Spark (310220122) Daniel Farhat (310205565) Michael Holmes (311246044)


A E R O 2 7 0 5 A S S I G N M E N T 4

TABLE OF CONTENTS

1. Comet Intercept Mission and Launch Vehicle: Energia ................................................................................................... 3

1.1 Mission objective ........................................................................................................................................................................... 3

1.2 Launch................................................................................................................................................................................................ 3

1.3 Comet-Intercept Orbit ................................................................................................................................................................. 3

2. Scientific Package ............................................................................................................................................................................ 5

2.1 Proposed Mission .................................................................................................................................................................. 5

2.2 Instrument List ....................................................................................................................................................................... 5

3. Power Generation and Storage .................................................................................................................................................. 6

3.1 Power Balance ................................................................................................................................................................................ 6

3.2 Power Regulation .......................................................................................................................................................................... 7

4. Thermal Regulation ........................................................................................................................................................................ 8

5. Communications .............................................................................................................................................................................. 9

6. Computer control system.......................................................................................................................................................... 11

6.1 Microprocessor ........................................................................................................................................................................... 11

6.2 Clock ................................................................................................................................................................................................ 11

6.3 Data Buses ..................................................................................................................................................................................... 12

6.4 Storage ............................................................................................................................................................................................ 12

7. Position and Attitude .................................................................................................................................................................. 14

7.1 Sensing Components ......................................................................................................................................................... 14

7.1.1 Issues with translational movement determination:......................................................................................... 15

7.1.2 Long Range Position Determination: ........................................................................................................................ 16

7.2 Actuating Components ..................................................................................................................................................... 17

7.2.1 Attitude control ................................................................................................................................................................. 17

7.2.2 Position control ................................................................................................................................................................. 18

8. Structure and Layout .................................................................................................................................................................. 19

9. System Integration ....................................................................................................................................................................... 20

9.1 Heat Control ................................................................................................................................................................................. 21

9.2 Position and Attitude Control ............................................................................................................................................... 21

9.3 Data Collection ............................................................................................................................................................................ 21

10. Mass and Polar Inertia .......................................................................................................................................................... 22

11. Range finder .............................................................................................................................................................................. 24

11.1 Principles .................................................................................................................................................................................... 24

11.2 Design ........................................................................................................................................................................................... 24

Bibliography ............................................................................................................................................................................................. 29


1. COMET INTERCEPT MISSION AND LAUNCH VEHICLE: ENERGIA

1.1 MISSION OBJECTIVE

The satellite has been designed to gather data from a comet on an elliptical orbit with an aphelion of 1,433,449,370 km and a perihelion of 83,058,550 km. The purpose of the satellite is to obtain readings of the comet’s water and chemical composition, and transmit the data back to Earth for analysis. The satellite is expected to operate for approximately 3-5 years minimum, not including the travel time to intercept the comet.

1.2 LAUNCH

The satellite has been simulated to launch on the Russian Energia launch vehicle, from the Baikonur cosmodrome in Kazakhstan. The Energia rocket used was the original 4 booster configuration, used in the 1987 and 1988 launches. The satellite will be part of a shuttle attached to the launch vehicle.

Launch will be conducted from Baikonur Cosmodrome using Energia launch vehicle directly into a polar orbit trajectory with a total payload mass of 140 tonnes on 22nd June 2020 at 9am. Within Low-Earth Oribt 70 tonnes of other missions shall be released into polar orbit, and the satellite and comet-intercept stage deployed.

1.3 COMET-INTERCEPT ORBIT

The satellite will be attached to an intercept vehicle with a separate fuel canister of approximately 60tonnes and an Athena 1 engine. This vehicle will provide the thrust necessary to escape earth’s gravity and later match the speed of the comet as it moves along its orbit. The orbit consists of two main thruster burns: the first to escape earth at escape velocity and trajectory parallel to earth’s motion, and the second to make minor adjustments to the flight path to match the comet once within a close proximity. At this stage, the satellite is released close to comet aphelion and the satellite and has the same orbit as the comet and is within 50km of the comet’s centre. This journey is expected to be completed by April 2024, with a journey time of three years and 10 months.

For the simulated flight path, see Figure 1 over the page.


FIGURE 1: COMET INTERCEPT ORBIT TRAJECTORY

- Blue: Comet Path - Black: Earth Orbit - Red: Rocket Intercept Trajectory


2. SCIENTIFIC PACKAGE

2.1 PROPOSED MISSION

The mission shall investigate the role of comets in assisting the development of life on Earth. The question of how Earth alone of all the known planets came to support life is one of great interest to science. It has been proposed that some of the requirements for life to form were not native to planet Earth, but were instead supplied by the impact of comets near the beginning of the solar systems. These include basic amino acids, the building blocks of proteins essential to cell interactions, and water. As comets often have very stable orbits over hundreds of millions of years they are of interest because they can provide a glimpse into the history of the solar system.

Spectral analysis and prior comet probe missions have shown that comets created near and beyond Neptune have much higher proportions of heavy water than found in Earth’s oceans. However Comet Linear, theorised to have formed around the orbit of Jupiter, had a similar composition to water on Earth. [1] As the target comet currently orbits between Mercury and Saturn, it is likely to have been formed in a similar area. Close analysis of this comet will provide interesting results to the possibility of comets seeding water on earth essential for life.

The second objective is to focus on attempting to identify the presence, type and concentration of any amino acids present within the comet. Amino acids were first detected by the Stardust probe on Comet Wild 2 [2]. It has also been theorised that in the intense environment of a comet impacting the Earth complex proteins were created by chemical reactions of organic molecules like the amino acids [3]. To investigate this possibility data about the composition, size, shape and density of the comet shall be gathered in order to simulate comet impact.

The instruments have been selected for the purposes of this mission. Space missions usually have a completely custom-built scientific package given their very specific operating criteria. The details provided for the instruments below have been derived from comparison to the Rosetta Spacecraft mission currently in transit to its target comet.

2.2 INSTRUMENT LIST

Dust mass spectrometer – the most important instrument for determining the presence and quantity of organic molecules like amino acids. Its ability to collect samples ejected from the comet surface allow it to perform a much greater level of detailed analysis with the addition of determining masses of molecules to identify them. Mass 20kg. Power 20W at 30V. [4]

Far Ultraviolet Spectrometer –Capability to also measure noble gases tells us thermal history of comet –how close to the sun it has approached based upon what solid noble gases remain. Also able to detect major components of organic molecules: sulphur, nitrogen, carbon, oxygen and hydrogen atoms all. Mass 3kg, Power 4W at 30V. [5]

Microwave Spectrometer – Primary unit for detection of water and carbon monoxide molecules. Detection of isotopes will be used to examine deuterium concentration for comparison to earth water bodies. Mass 20kg, Power 43W at 30V [6]

Visible and IR Camera – This gives shape, size and rotation of comet by visual reference. Also used for visual observation of comet erosion over time and sources of dust from surface. Some investigation of chemical composition, dust and gas emission near nucleus surface. Total mass 35kg includes 22kg of camera lenses, with the remainder image sensing and signal processing as well as secure harness and shielding. Power 57W at 30V [7]

Radio Sounding Instruments – determines internal structure. This is conducted by inference on the impact of transmitting radio signals through from the satellite to earth through the body of the comet nucleus.


3. POWER GENERATION AND STORAGE

At the extreme distance from the sun of the comet intercept point solar power is unfeasible to operate the satellite. To power the comet probe a radioisotope thermoelectric generator (RTG) will be used, which employs thermocouples to generate an electric current using a thermal gradient within the material – hot near the radioactive source and cooler near the heat-sink, to create a flow of charge carriers. It will be of comparable size to those used on the deep-space Cassini, Galileo and Ulysses Missions, being about 60kg mass, producing an electrical power output of 300W at 30V while producing 4400W of heat during operation. [8] The power supply is usually Plutonium-238, with a 90yr half-life and power output of 0.5kW/kg. This power generation method was chosen for its high continuous output over a long life– losing around 25% output power after 20 years of operation, [9] with high reliability due to its mechanical simplicity.

This device does have the disadvantage of producing large amounts of heat that need to be dissipated however its ability to generate power independent of the distance to the sun outweighs this drawback. During the transfer orbit the RTG circuit loop will not be closed so electrical power is not generated. In order to dissipate the heat from the radioactive source heat dissipation systems will be built into the satellite delivery system that carries the instrument package to the comet.

Batteries will be used to store backup energy. These will primarily be used to provide surge currents for running heating coils when a position manoeuvre is required and fuel pumping and control components are too cold to operate effectively. In case of generator failure the batteries will power the entire satellite for a short time, long enough to send a well-defined termination message to earth to report mission failure as without a power source the satellite cannot operate.

The batteries will be sourced from GS YUASA from their LSE Series. An array of four LSE175 batteries shall be used, each weighing 5kg with a capacity of 800Wh. [10] These were chosen for their reliability – the LSE series is currently in use in many operational satellites; for their low-mass large energy storage capacity; and for their life expectancy of 18yrs well longer than the mission duration.

3.1 POWER BALANCE

TABLE 1: POWER REQUIREMENTS OF COMPONENTS

Device Voltage Required (V) Power Required (W)

Mass Spectrometer 30 20

Far UV Spectrometer 30 4

Microwave Spectrometer 30 43

Visible and IR Cameras 30 57

Transponder 30 15

Communications Amplifier 30 20-165

CPU and Atomic Clock 5 5

Twin Star Trackers 28 7.5-9

Inertial Measurement 28 24

Control Moment Gyros 28 30

MAX TOTAL POWER DEMAND 372

OUTPUT POWER 300


The power balance shows that we cannot run every device simultaneously, as close to the maximum possible power draw of the communications amplifier will be required to transmit over the great distances from the comet to earth. However, the IMU as a redundant system and the Control Moment Gyros do not need to be run continuously, and as such there is only a small restriction placed on the scientific instruments in that at least one of the microwave spectrometers or the visible/IR cameras cannot be run at the same time as the other instruments while transmitting data to earth.

3.2 POWER REGULATION

As devices require different power levels yet the generator produces a constant 30V supply regulation is required to supply power to each device at its correct operating level. The most important of these is the CPU which needs only 5V supply and 30V will cause it to fail. All other devices however able to be operated within a 25-35V range (the voltage requirement listed in Table 1 is the voltage at which the power demand has been specified by the manufacturer) and so no voltage regulation is required.

Simple resistor-based voltage dividers shall be used where appropriate to output the required voltage to the computing elements, while impedances to lower the current to the required levels will also be implemented to not overload electronic circuitry.


4. THERMAL REGULATION

Temperature sensors shall be placed on all key components including instruments, fuel and propulsion systems to ensure satisfactory operation within their operating regions of -10 to 50 °C.

The main forms of temperature control are use of the main rocket engine and heating coils for heating, and infrared radiators for cooling. At the extreme distance from the sun that the spacecraft operates heating is more likely to be required. Existing heat sources including the RTG power source, internal electronics and the chemical fuel engines will often be sufficient to keep the satellite within optimal operating conditions; however key instruments like fuel pumps will be fitted with heating coils.

If excess heat does need to be dissipated infrared radiators with a high emissivity factor is the primary form of cooling. The heat output rate at a given surface temperature due to IR radiation is given by

. Where is the emissivity of the surface, is the Stefan-Boltzmann constant, A is the surface area and T the surface temperature. Thus a large-surface area coated with a high-emissivity material such as carbon encased in glass [11] is used for cooling, with heat dissipated related to duration of emission.

Electronic device operator or engine burns need to be reduced or stopped during a cooling period as the cooling is time dependent and cannot be controlled [12]


5. COMMUNICATIONS

The communications system of the satellite will have to be designed for deep space transmission through the deep space network (DSN). The satellite will use the X and Ka [1] bands to provide uplink and downlink speed. The antennas will consist of : A parabolic High Gain Antenna (HGA) Multiple Low Gain Antennas (LGA) A Medium Gain Antenna (MGA). [2] The HGA is mainly used for all data transmissions, although the other antennas are used at different stages of a spacecraft’s mission and in case of an emergency. The high power amplifier chosen for telecommunications system is a traveling wave tube amplifier (TWTA) as opposed to a Solid State amplifier (SSA)[3]. The TWTA was chosen as in general they are smaller, provide a stronger signal and use less power, making them more efficient. The TWTA is The high gain parabolic antenna uses X band at 8.2GHz communication for earth deep space network uplink and instrument data downlink. This can obtain a link with a much high transfer rate due to the high gain and narrow beam [4]. Rates of (130kb/s to ~3mb/s)[5] can be obtained depending on the proximity to earth. The high gain reflector is fed by a cone antenna positioned about the focus of the parabolic dish to produce a HGA with 40-60dB gains. The medium gain antenna is used for communication when the HGA is not directly in line with the earth. This happens when the satellite is in thrust manoeuvres or travelling close to the sun. The MGA is used in these cases, as its transmission beam allows 10-20 degree range allowing it to transfer data without being pointed directly at earth. However the gain achieved is only 15-22dB The low gain antennas have a broad beam being able to transmit in all directions. These antennas are used for emergencies and when the MGA is not in close direction to the earth. (E.g. orbital manoeuvres.) The LGA have a small gain of 0-5 dB. This results in very slow data rate where only small amounts of information can be sent and received, such as manoeuvre changes. The three low gain antennas are placed around the satellite to achieve the widest coverage for the satellite. The two helix LGA’s are place on the aft and front of the satellite while the pipe antenna gives coverage in-between the two helix antennas. This provides an addition of three coverage areas [6].

The medium gain antenna and the high gain antenna are placed in pair to allow a constant data transfer when the high gain antenna drifts slightly. The medium gain antenna is used to ensure there is no break in transmission when the high gain antenna cannot provide a data feed. The medium gain antenna is placed on top of the feed antenna for the dish reflector to ensure that it is pointed towards earth when high data rates occur. The X-band transceiver is used to provide a low power controlled output to the amplifiers with the desired frequency. The transceiver is capable of using Ku and Ka bands to provide Tracking and Data

Pipe

Helix Helix


relay with the TDRSS. This is done for satellite tracking and to provide a data connection with higher data rate.

TABLE 2: COMMUNICATIONS PARTS LIST

Part Power Length Width Height Diameter Mass(Kg) TWTA 20-165W 393 70 64 n/a .97

EPC (Regulator)

n/a 178 80 95 n/a 1.05

Transceiver (HRT-440)

15 W 203 168 76 n/a 2.3

2*Helix LGA n/a n/a n/a 250 100 .35

Pipe LGA n/a n/a n/a 216 286 .9

MGA n/a 686 216 300 n/a 3.72

Reflector Dish n/a n/a n/a 2.7(m) 3.3(m) 25

Reflector feed n/a 520 n/a n/a 280 2.7


6. COMPUTER CONTROL SYSTEM

There are many electronic devices within a satellite, included a multitude of sensors, scientific packages, thruster drivers and radio transponders. In order to control all these electronics and process their inputs and outputs in order to achieve a desired result, a computer control system must be used. It consists of four main parts: the microprocessor, the clock, data buses and storage. The microcontroller chosen to fulfil this need is the RAD750 6U CompactPCI single-board computer built by BAE systems. It specifically contains radiation hardened devices that are able to be used in the outer space environment, and as such is appropriate for the satellite.

6.1 MICROPROCESSOR

The microprocessor handles all arithmetic and logic flow within the system. A basic microprocessor is made up of three parts:

Arithmetic Logic Unit (ALU): handles all arithmetic and

logical operations such as addition, subtraction, AND,

OR, etc.

Control Unit: deals with all instruction sets that the

microprocessor has been provided with; decoding

them and sending the appropriate data and low level

operations to the ALU to process.

Registers: stores intermediary data, instruction sets

and flags which symbolise the current state the ALU.

The microcontroller chosen for the satellite uses a RAD750 microprocessor, which is a “radiation-hardened” version of the popular IBM PowerPC 750. It has been designed to be clocked at 200MHz, withstand a larger temperature range than most microprocessors (-55 to 125 degrees Celsius) and resist up to 1Mrad of radiation. As such it is specifically designed for space applications, as far as being used in the current Mars Curiosity Rover and the comet chasing Deep Impact probe.

6.2 CLOCK

The clock synchronises cycle times to ensure that every process occurs at the correct moment. The microcontroller has an inbuilt quartz clock, but due to inaccuracies during spaceflight and relativistic effects, this will be compared with an atomic clock in order to ensure the accuracy of timing with regards instructions being received and executed as necessary.

One such clock is the Symmetricom SA.45s chip scale atomic clock (CSAC) which is a commercially available atomic clock in an electronic chip format, having a speed of 10MHz. This does reduce the speed at which our microprocessor can operate by a factor of 20, however specific use of the SA.45s can be restricted to time related operations, rather than being used at all times for menial tasks where timing is not an issue, e.g. compression of data during recording. The SA.45s has a mean time between

FIGURE 3 EXAMPLE OF 6U COMPACTPCI COMPUTER

FIGURE 4 IMAGE OF A RAD750 MICROPROCESSOR

FIGURE 2: EXAMPLE OF 6U COMPACT PCI COMPUTER


failures of 50,000 hours and can withstand temperatures between -40 and 85 degrees Celsius, which is well within the ranges of the outer space environment.

FIGURE 5 DETAILED VIEW OF THE SA.45S CSAC

6.3 DATA BUSES

Data buses are external connections to other peripherals, such as sensors, transmitters and actuators. The primary data bus that the microcontroller uses is called SpaceWire, which is a very adaptable serial interface; having a capacity of up to 400 Mbit/s.

The SpaceWire Endpoint must be used to convert the bus into other, more common data bus types, such as SPI, UART, I2C and PCI. It operates at a speed 100MHz, and can withstand temperatures of -55 to 125 degrees Celsius and at minimum 1 MRad of radiation. A SpaceWire Router can also be used in order to increase the amount of ports that are able to be connected to the microcontroller.

6.4 STORAGE

Storage is the location where data is kept for either further analysis/calculation or as a temporary location before being transmitted back to Earth. It also stores instructions and precompiled code for the microcontroller.

For internal storage, the microcontroller contains up to 48MB of RAM, up to 1MB of ROM, and up to 4MB of EEPROM. These forms of storage will primarily be used for temporary data and calculation storage. The size of storage can be expanded by connecting an external storage device to the microcontroller, which can further be used to temporarily store images and data for future transmission.

FIGURE 6 SPACEWIRE ENDPOINT


The external storage device that has been chosen for the satellite are the FLASH NAND Space Grade Radiation Tolerant Memory Stacks from 3D Plus. They are available in sizes up to 128GB, can withstand 50 kRads of radiation, and have a lifespan of 15 to 18 years in an outer space environment. As such, they are the most appropriate form of external storage for the satellite.

FIGURE 7 EXAMPLE OF THE 3D PLUS RADIATION TOLERANT MEMORY


7. POSITION AND ATTITUDE

7.1 SENSING COMPONENTS

2 x EADS SODERN SED26 Star Tracker [1]

Mass 3.1kg each

Dimensions: 160x170x290mm each

Power Consumption: 7.5W at 20deg, 9W at 40deg, each

Output Interface: RS422 Serial Protocol

Input Voltage Range: 20-50V

Operating Temperature Range: -20deg → 50deg Celsius

Update Rate 10Hz

Tracking up to 10 stars each simultaneously

The purpose of this component is to serve as the primary system for attitude determination. It operates by analysing an image taken by its lens with respect to a stored database of star maps for accurate determination of the current rotation of the craft in the XYZ dimension space.[1] Although one unit would be sufficient for basic operation. Two units increases the field of view visible to the craft, allowing for faster alignment determination of the craft with respect to either the comet or the Earth; as well as offering redundancy, since the Astrium IMU cannot determine absolute attitude position alone. [4] Attitude determination with respect to the comet aids in appropriate attitude control whilst in close proximity, whereby the attitude reading is coupled with distance readings from the IR proximity unit. Conversely, attitude determination with respect to the Earth is necessary for the proper alignment of the antennae for communications and long-range position tracking.

EADS ASTRIUM ASTRIX 120 Fibre Optic FOG Inertial Measurement Unit [2]

Mass: 2kg Inertial Core Unit (ICU), 4kg Gyro Electrical Unit (GEU)

Dimensions: 215diam x 180h mm (ICU), 270x150x145mm (GEU)

Power Consumption: 6W/channel, 24W total

Output Interface: 1553/RS422 Protocols

Input Voltage Range: 22-50V

Operating Temperature Range: -10deg → 50deg Celsius

This component is primarily tasked to act as a redundancy to the primary Star Tracking attitude sensors, which cannot operate effectively if the craft encounters regions of dust particulates, where the collection of light through the lenses is inhibited by particle interference. It is more responsive to the change in rate of velocity than its counterparts, albeit requiring more power to operate. As such it is not running constantly, only when failure is detected in either of the 2 Star Trackers, or when passing through a region of low visibility. This device operates by exploiting the Sagnac Effect, whereby two beams of light are fired around a circular loop in opposite directions. Even if the craft is rotating at a slow rate of 0.01 degrees per hour, the incredibly minor change in length of the path of the beam of light requires more time for the beam to complete its revolution, causing a phase difference. [3] The phase difference between each beam can then be used to determine the rotational velocity, as one will complete its circuit faster, the difference between equalling twice the change in phase. The device runs with 4 skewed planes of detection for very accurate tracking of attitude, however, these devices operate with an inherent bias error of 0.001 degrees per hour, which occurs due to sensory and computational limitations. As such it must be recalibrated through use of the Star Trackers to determine the absolute attitude at least every few hours. Because of this it is not designed to operate as the sole means of attitude detection indefinitely, as it is only capable of tracking the rate of change of the angular velocity, not the absolute velocity.


The ASTRIUM unit has the lower life expectancy, of 15 years, as opposed to the 18 years expected from the SODERN units. As the mission expected duration is approximately 4 years, the devices should be fully operational throughout the course of the mission.

7.1.1 ISSUES WITH TRANSLATIONAL MOVEMENT DETERMINATION:

The determination of the attitude of the craft with respect to time is relatively simple, as outlined above; with the known orientation of the Star Tracker with respect to the craft, the current rotation of the craft with respect to the surrounding stars can be found, and then adjusted to coincide with the craft's current task, whether it be to position a particular edge of the craft towards the comet for sample collection/ image capturing, or to position the antennae towards Earth for maximum error reduction in communications. Theoretically this can all be achieved with a single Star Tracker, and as such the second Star Tracker plus the Fibre Optic IMU were included for efficiency and redundancy issues.

The determination of translational position change however, is much more difficult, and cannot be achieved by one Star Tracker alone. Similar to how a single human eye cannot correctly perceive depth without its counterpart, a lone Star Tracker is incapable of discriminating translational movement from rotational, attitude movement. This scenario is presented with some simple 2-dimensional example sketches below:

In the sketches above, the solid lines represent the current position of the craft/tracker, whilst the broken lines mark the previous position, with translational movement on the left, and rotational movement on the right. As can be seen, the stars within the lone tracker's field of view will appear pretty much identical in both scenarios, making it impossible to distinguish a difference.

The solution is to mimic the human eye, whereby a second Star Tracker is used. Whilst it could theoretically be positioned in any other position on the craft, the direct opposite is chosen as it yields access to a completely different star field. This is illustrated in the third sketch shown below. By using two different reference star fields from two different Star Tracking cameras, rotation can be distinguished from translation by comparing the change in rotation recorded by both cameras. If the craft is undergoing pure rotation, the measured rotations of each camera would be equal in magnitude and sense (i.e. both trackers would see an anticlockwise rotation). Conversely, a purely translational movement in the y direction would yield rotations of equal magnitude, but opposite sense (right camera would see an anticlockwise movement, left would see clockwise). The difficult lies in separating combined translational and rotational movement. Such a manoeuvre would incorporate attitude changes of different magnitudes and opposite sense. Also this concept is difficult to translate

FIGURE 8: REQUIREMENT FOR DUAL STAR-TRACKERS


into the 3 dimensional domain as the craft could be rotating/translating about any 3 of the reference axis simultaneously. Thus a complex algorithm would be required, yet it would be feasible.

7.1.2 LONG RANGE POSITION DETERMINATION:

Long range position determination is more trivial, and for this particular craft is determined by a combination of star tracking and communications time-encoding. At interception, the craft is too far into deep space to bounce waves reliably off of planets/other objects in order to triangulate its position. Rather, it relies on the time-encoding of data streams sent to and from the satellite. With the angle of the craft relative to the Earth is logged at the instant a data stream is received, as well as the time of receive, the time taken for the data stream (travelling at the speed of light) to reach the craft can be calculated, and thus the distance. Obviously a position cannot be triangulated off one point, and there are no other points of reference, so a position can be ascertained by taking subsequent measurements and determining the position by generating a theoretical path of positions. Of course the main obstacle with this approach is the correction for the asynchronous nature of the clocks due to relativistic effects. This requires a complex mathematical model to predict the change of the on-board clock with respect to an atomic clock on Earth. Even with a good model, the adjustment is only approximate, and therefore the actual position of the craft with respect to the Earth will incur some significant error. With more successive measurements comes greater accuracy, thus this method will be adequate for making mid-course corrections on the way to the comet. Final approach towards the comet however would require more accurate position sensing and adjustment, most likely through the use of a larger-scale equivalent of the IR sensor, albeit custom made/non-commercial. Once final adjustments have been made, the position of the craft with respect to Earth can be effectively treated as the position of the Comet with respect to Earth, with minute adjustment from the attitude sensors/IR sensor.

FIGURE 9: USING TWO STAR-TRACKERS


7.2 ACTUATING COMPONENTS

7.2.1 ATTITUDE CONTROL

Attitude control in satellites used in deep space orbits usually fall into three categories: a thruster based system, a gyroscopic motion based system, or the use of a solar sail [18]. However, solar sail technology is still under testing, and has a large size requirement, and hence was not considered. A thruster based system was not chosen as these systems have a limited operational lifetime, as each of the 12 thrusters requires fuel to operate. This fuel would also increase the mass of the vehicle, as well as use a considerable amount of space.

There are two sub-categories of gyroscopic control: a reaction wheel control system, which varies the spin rate of the wheel to rotate the vehicle; and a Control Moment Gyroscope (CMG) system, where satellite spin is controlled by the rotation of the flywheel [19].

Due to the satellite’s small size and weight, using a set of 4 reaction wheels may be the best component to use for attitude control, one wheel for each axis direction, and a redundant wheel to allow for limitless rotation. Reaction wheels were chosen over a CMG system due to space consideration, extra reliability, and more precision for smaller angle changes, which will most likely be the main use of the attitude control system [20]. The extra power requirement is minimal for a satellite of this size. A possible reaction wheel to use would be the MicroWheel 1000 Reaction Wheel from Microsat System Canada Inc., with an optional rate sensor which can be used instead of an external attitude sensor. This is all controlled through a common serial interface to the satellite’s main logic computer. Each MW1000 has a size of 130 x 130 x 90mm, a mass of 1.4kg, a power consumption of at most 10W at 28 Volts, and a maximum torque of 30 milliNewtons metres. Each wheel has an operation lifetime of at least 5 years [21].

For the final satellite design, a system consisting of 6 MicroWheel 1000 reaction wheels was chosen. Two reaction wheels will be placed along each of the three major axes, for a secondary redundancy factor in the case that the primary wheel malfunctions. Each of these six wheels will contain the optional rate sensor provided by the manufacturer. Similarly, these rate sensors act as a redundant attitude sensor system in case of primary operation failure. Using two wheels on each axis also removes the need for the fourth bi-axial reaction wheel as the same methods of almost unbounded rotation can be created.

7.2.1.1 COMPONENT SPECIFICATIONS

Actuator: 6 x MicroWheel 1000 Reaction Wheel

Manufacturer: Microsat Systems Canada Inc.

Size: 130x130x90mm

Mass: 1.4kg

Speed range: +/- 10,000 RPM

Torque capacity: 30 milliNewton metres

Power consumption: <10 Watts

Input Voltage: 28V +/- 6V

Operational Temperature range: -30 to 60 degrees Celcius


Data interface: RS-422/485 serial port

Rate sensor performance: 2-500 degrees/sec, with a drift rate of less than 0.35 degrees/sec

Operational life: approx. 5 years

7.2.2 POSITION CONTROL

For small adjustments in the translational position of the satellite, a single thruster should be used and mounted on the back end of the satellite, to minimise the effect of created interference on the communication array. Only a small thruster is needed for orbital manoeuvrability, so using an efficient liquid fuel to easily store on the satellite, or an ion thruster would be an appropriate choice [22]. Example thrusters would be the Atlantic Research Corporation’s Leros 2R thruster, the Diamler-Benz 400S, or the Astrium 400 N thruster. Each of these weight approximately 3-4kg, produce 400-550N of Thrust, and are fuelled by Monomethylhydrazine (MMH) and oxidised by MON-3 combination [23] [24].

7.2.2.1 COMPONENT SPECIFICATIONS [24]

Actuator: 400N Bi-Propellant Engine Model S400-12

Manufacturer: Astrium

Nominal Thrust: 420N

Specific Impulse: 318s

Flow rate: 135g/s

Inner fuel diameter: 16.4mm

Nozzle diameter: 244mm

Mass: 3.6kg

Fuel : MMH with MON-3 oxidizer

Fuel Mixture Ratio: 1.65

Size: Inner components: 180x85x160mm (lxbxh) Outer flow nozzle: Cone base radius 248mm, height 340mm Centre of gravity: 161mm Total length: 503mm


8. STRUCTURE AND LAYOUT

The design layout of the satellite is designed to maximise heat flow around the satellite. The fuel for the main thruster is best placed at the centre of the satellite, as the majority of the satellite's mass will be in the fuel. The reaction wheels for attitude control should be placed near the centre of mass, on one side of the fuel tank, along with the lithium ion batteries, and the control systems. The placement of the nuclear power generator should be placed on the opposite side of the fuel tank, to have even mass distribution. The communication array and scientific tools would be placed on the top face of the satellite, with the two large cameras placed on two opposite sides, and the main engine on the bottom side, to reduce interference caused by thrusting. Position sensing equipment can be evenly distributed around the satellite, to positions where they are most useful, as the weight and size of the equipment is not incredibly large. In order to contain the system components, as well as heating coils, the main internal chassis of the satellite would be a cube with approximate side lengths of 2 metres, with thrusters, communications and cameras extruding from the chassis. The aim of the distribution of components is to keep mass symmetrical across all three axes.

The chassis would be constructed from steel for strength and heat conduction, with an internal casing around the fuel tank and thruster. The outside of the chassis would have lead plating to reduce the effect of radiation on internal equipment.

FIGURE 10: SPACECRAFT LAYOUT


9. SYSTEM INTEGRATION

FIGURE 11: SYSTEM DIAGRAM


In order for the spacecraft systems to not interfere with each other they have a priority order, and lower priority systems only run on conditions of the higher.

Larger-scale systems that do not affect the everyday operation of the satellite, such as commands from earth to terminate the current mission and start towards a different comet, have not been included here given their irregular occurrence.

9.1 HEAT CONTROL

Controlling heat is a primary concern of the spacecraft, as it is a physical quantity that can damage the craft and yet is able to be managed with an appropriate response. As such the heat sensing have the ability to override and stop the operation of other systems as long as they are not essential for the life of the craft. Systems that can be shut down include the scientific instruments, communications and use of the thruster until the craft has sufficiently cooled.

Heat control logic performed by the CPU takes analogue inputs from the various heat sensors around the craft and compares them to the required values. As individual components approach temperatures towards the top of their operating range they are switched off. If the temperature drops towards the bottom of the operation range, particularly for pumps or motors, then heating coils are activated – possibly using additional power stored in the on-board batteries, until operating temperature is reached.

9.2 POSITION AND ATTITUDE CONTROL

Getting the instrument package to where it needs to be to perform its mission is the next priority once heat is stabilised to within the operating range of components. However, as the use of the chemical rocket engine produces a cloud of expelled propellant it will interfere with the mass and optical spectrometers. As such the position control system has override control over the scientific instruments, which are only reactivated when the craft is sufficiently far away from the location where the burn was made.

The CPU analyses the position compared to where it is supposed to be according to its mission plan. If the attitude requires adjustment the gyroscopes have their speed controlled proportional to the amount of angular change required. If the position needs to be changed, the gyroscopes must first be employed to rotate the craft so the unidirectional rocket engine points along the required line of motion. Then the thrusters can fire for a time sufficient to provide the required total impulse.

9.3 DATA COLLECTION

Once heat has been managed and position adjusted, and all chemical propellants cleared from the operation regions of the instruments, then the scientific instrument package can begin collecting data. Communications and commands received from Earth as well as pre-programmed mission parameters will determine which instruments are employed and how they are used. The data is then output back to the CPU either as an analogue signal or a digital (some instruments perform internal A/D conversion). Either the data is then compressed and sent back to Earth, or stored in memory if communications are temporarily down until they are restored.


10. MASS AND POLAR INERTIA

TABLE 3: POLAR INERTIA CALCULATIONS

Glo

ba

l

Izz 4

.91

3.7

7

3.7

7

0.0

9

3.7

4

0.0

1

0.1

6

0.0

2

0.9

9

0.2

8

0.2

3

81

.06

5.3

4

0.2

5

0.0

3

0.1

1

0.1

3

To

tal

Izz

10

4.9

0

Glo

ba

l

Iyy 0

.04

0.0

4

0.0

4

0.0

0

0.0

3

0.2

0

0.3

3

0.0

0

0.2

6

0.0

0

0.0

0

81

.06

5.3

4

0.1

6

0.1

4

0.1

1

0.1

3

To

tal

Iyy

87

.88

Glo

ba

l

Ixx 4

.05

1.8

9

1.8

9

0.0

3

4.0

4

0.0

1

3.5

9

0.0

2

3.6

1

0.0

9

0.2

1

24

1.9

4

21

.00

0.0

5

0.8

0

2.2

6

5.1

4

To

tal

Ixx

29

0.6

3

zz

Off

se

t

3.1

8

3.1

6

3.1

6

0.0

5

2.5

9

0.0

0

0.0

0

0.0

0

0.8

1

0.1

9

0.0

1

0.0

0

0.0

0

0.2

2

0.0

0

0.0

0

0.0

0

yy

Off

se

t

0.0

0

0.0

3

0.0

3

0.0

0

0.0

0

0.2

0

0.3

1

0.0

0

0.2

6

0.0

0

0.0

0

0.0

0

0.0

0

0.1

3

0.1

1

0.0

0

0.0

0

xx

Off

se

t

2.3

0

1.3

7

1.3

7

0.0

0

3.1

2

0.0

0

3.5

5

0.0

0

3.4

4

0.0

0

0.0

2

16

0.8

8

8.1

7

0.0

0

0.7

5

2.1

6

5.0

3

Lo

ca

l

Izz 1

.73

0.6

1

0.6

1

0.0

3

1.1

4

0.0

1

0.1

6

0.0

2

0.1

8

0.0

8

0.2

2

81

.06

5.3

4

0.0

3

0.0

3

0.1

1

0.1

3

Lo

ca

l

Iyy 0

.04

0.0

1

0.0

1

0.0

0

0.0

3

0.0

0

0.0

1

0.0

0

0.0

0

0.0

0

0.0

0

81

.06

5.3

4

0.0

3

0.0

3

0.1

1

0.1

3

Lo

ca

l

Ixx 1

.75

0.5

3

0.5

3

0.0

3

0.9

2

0.0

1

0.0

4

0.0

2

0.1

7

0.0

9

0.1

9

81

.06

12

.83

0.0

5

0.0

5

0.1

1

0.1

1

Dep

th

a

0.9

7

0.7

5

0.7

5

0.3

3

0.6

8

0.2

0

0.2

2

0.1

3

0.2

9

0.3

9

0.3

0

0

.87

dia

met

er

dia

met

er

dia

met

er

dia

met

er

dia

met

er

Wid

th

l

0.3

9

0.2

6

0.2

6

0.0

9

0.3

0

0.1

7

0.3

0

0.0

9

0.1

6

0.1

8

0.1

5

0

.87

0.4

2

0.2

2

0.2

5

0.2

8

0.2

5

r rad

ius

Heig

ht

b

0.3

8

0.4

1

0.4

1

0.1

5

0.4

8

0.0

8

0.6

9

0.1

3

0.1

7

0.1

5

0.2

0

0

.87

1.1

4

0.2

2

0.1

0

0.5

2

0.5

0

l len

gth

Ma

ss

1

9.1

0

10

.00

10

.00

3.0

0

19

.90

2.3

0

3.7

2

1.4

0

3.1

0

6.0

0

20

.00

10

0.0

0

65

0.0

0

60

.00

2.0

0

1.6

0

2.7

0

3.6

0

Pa

rt

Mass

Sp

ec

Wid

e C

am

Narr

ow

Cam

UV

Sp

ec

Mic

roS

pec

Tra

nsc

eiv

er

Med

ium

Gain

6xR

eacti

on

W

2xS

tarT

rack

IMU

Batt

ery

Str

uctu

re

Fu

el

Th

erm

alG

en

Co

ntr

oller

Lo

wG

ain

Hig

hG

ain

En

gin

e


The preceeding data is calculated based upon a number of simplifications. Each individual component has been approximated as either a homogeneous rectangular prism or cylinder. The centre of mass of each component has been approximated to be the centre of the component. Due to the highly complex shape of the structure/shielding, it has been omitted from the calculations, as it weighs 100kg it is a considerable omission, though with proper distribution the balance of moments about the different axis would remain approximately the same. The moment of inertia calculations have been taken about a reference axis centred on the array of reaction wheels. With the X direction nominated as the axis from the reaction wheels to the engine, the Y axis being parallel with the Thermal Generator, and the Z axis perpendicular to these two, coincident with the narrow and wide field cameras.

Before global arrangement was undertaken, certain parameters were set to ensure the proper orientation this included:

Keeping the engine and fuel tank in relatively close proximity.

Keeping the core electronics well shielded from the exterior and relatively close together.

Keeping the narrow and wide angle cameras pointing in the same direction so that both can be used simultaneously to photograph the comet without attitude adjustment.

Having the Star Trackers pointing in opposite directions so as to employ a diverse field of view.

Grouping the antennae together so they can run simultaneously without attitude adjustment.

From here, components were placed in order of magnitude of mass. Thus the fuel tank was first centred, followed by the generator and spectrometers. The engine was then employed on the opposite side through the theoretical centre of mass so that its operation would incur as little rotational movement as possible. The cameras were then distributed laterally around the fuel tank to balance out inertial movement in the Z and Y directions. The antennae were grouped along the X axis to minimise inertial movement about the X axis. The remaining hardware was distributed laterally about the reaction wheels. At this point, the on-board micro-controller is well shielded from inbound radiation via the fuel tank and thermal generator. Although additional radiation hardening will be necessary for reduction in fault performance. It would be preferable for the cameras to be further away from the engine, however this was less feasible due to balancing issues with the generator and spectrometers.

Moments of inertia were calculated according to the following equations [1]:

Rectangle

Ixx 1/12*mass*(a^2+L^2)

Iyy 1/12*mass*(b^2+L^2)

Izz 1/12*mass*(a^2+b^2)

Cylinder

Ixx 1/4*mass*r^2+1/12*mass*L^2

Izz 1/2*mass*r^2

Where a = depth, L = width, b = height, r = radius; when looking upon the craft from the left face (see drawing).

Due to the thermal generator having a different orientation, its inertial X and Z dimensions were altered to X/Z and Y, respectively, otherwise the calculations remain very much the same once the components are oriented properly.


11. RANGE FINDER

11.1 PRINCIPLES

The task to accomplish consists of building a range sensor to measure a distance between 0cm and 100cm from a large object. The design choice was made to use an infra-red rangefinder due to its high accuracy and ease of integration. Another possible option would be a laser rangefinder, but due to the increased cost of the sensor, using infrared light was found to be a viable alternative.

The sensor works by using an infrared emitting LED to produce an infrared wave of a specific frequency to bounce off the object in question. The light is reflected into the photosensitive sensor and received. However, due to the separation distance between the LED and receiver, the incoming wave is detected at an angle dependent on the distance away from the measured object. [25]

FIGURE 12: OPERATION OF THE RANGE SENSOR [25]

The main disadvantages of such as system are the operating range of each sensor, and incorrect data due to noise. To compensate for the first, the system uses two IR rangefinders with overlapping operational ranges to give the system a total range of approximately 0-150cm. Due to the large distance away from the sun, infrared noise should be much smaller than it is on earth. However, to insure correct data, a band-pass filter could be implemented to only receive data from waves with the same frequency as the LED emits.

11.2 DESIGN

The main component of the sensing system is the Arduino microprocessor, which allows the system to have additional functionality if required. Firstly, to power the IR rangefinder, a voltage regulator is attached to the 9V line from the microprocessor to give the sensor the constant 5V that it requires.

The sensor sends the received signal back to an input port on the microprocessor. However, due to how the sensor operates, the signal is received as an analogue signal with a varying voltage. So, to properly process the data, the signal is passed through the board’s analogue-to-digital converter. The data is then mapped to a distance, which is found through the rangefinder’s datasheet. This is the same design for the second rangefinder. As the object moves to different differences, and processor uses the appropriate rangefinder data for accurate results in situations where the object is not detected by both sensors. To check contact with the option, a simple long-lever bump sensor is used. If this sensor is hit, the system is obviously in contact with the object.


The microprocessor can either then submit the digital data via a serial port, or by outputting a Pulse-Width Modulation (PWM) wave. Either option allows the system to be directly interfaced with the satellite’s processing unit, in an incredibly easy manner.

For the demonstration, a LED display may be attached to the sensor to easily show the audience the distance of the object from the system.

FIGURE 13: RANGE SENSOR BLOCK DIAGRAM


FIGURE 14: IR DETECTION CIRCUIT

FIGURE 15: ANALOGUE OUTPUT

FIGURE 16: VOLTAGE REGULATION


These look up tables give the Analogue-Digital Conversion result at each cm distance. 0 represents a 0V input, 1023 a 5V input.

TABLE 4: DISTANCE LOOK-UP TABLE – SHORT RANGE

Distance Value Distance Value Distance Value

0 0 34 24 68 0

1 325 35 21 69 0

2 557 36 18 70 0

3 567 37 15 71 0

4 464 38 12 72 0

5 401 39 9 73 0

6 350 40 7 74 0

7 313 41 6 75 0

8 270 42 5 76 0

9 244 43 4 77 0

10 220 44 3 78 0

11 197 45 1 79 0

12 180 46 1 80 0

13 167 47 1 81 0

14 157 48 1 82 0

15 146 49 1 83 0

16 137 50 1 84 0

17 128 51 1 85 0

18 119 52 1 86 0

19 110 53 1 87 0

20 101 54 1 88 0

21 94 55 1 89 0

22 87 56 1 90 0

23 80 57 1 91 0

24 73 58 1 92 0

25 66 59 1 93 0

26 61 60 0 94 0

27 56 61 0 95 0

28 51 62 0 96 0

29 46 63 0 97 0

30 40 64 0 98 0

31 36 65 0 99 0

32 32 66 0 100 0

33 28 67 0


TABLE 5: DISTANCE LOOK-UP TABLE - LONG RANGE

Distance Value Distance Value Distance Value 0 0 34 346 68 176 1 75 35 336 69 173 2 150 36 329 70 170 3 225 37 322 71 169 4 300 38 315 72 168 5 376 39 308 73 167 6 408 40 298 74 166 7 440 41 291 75 165 8 472 42 284 76 162 9 504 43 277 77 159

10 536 44 270 78 156 11 538 45 263 79 153 12 540 46 258 80 151 13 542 47 253 81 150 14 544 48 248 82 149 15 545 49 243 83 148 16 535 50 240 84 147 17 525 51 235 85 145 18 515 52 230 86 142 19 505 53 225 87 140 20 495 54 220 88 139 21 485 55 216 89 137 22 475 56 213 90 136 23 465 57 210 91 134 24 455 58 207 92 133 25 445 59 204 93 132 26 433 60 200 94 130 27 421 61 197 95 129 28 409 62 194 96 127 29 397 63 191 97 126 30 386 64 198 98 125 31 376 65 185 99 123 32 366 66 182 100 121 33 356 67 179


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