Mission Concept Paper for Project A.D.I.O.S.
- 1. Mission Concept DIOS MIO - Deimos Impact & Observation Spacecraft: Mars Intercept Orbit Team 3 Jeff Anderson, Thomas Blachman, Andrew Fallon, John Franklin, Samuel Gaultney, David Habashy, Brian Hardie, Brandon Hing, Zujia Huang, Sung Kim, Jonathan Saenger Overview/Objectives Mission Objectives The main objective of our mission is to analyze the subsurface composition of mars’s moon Deimos utilizing a spectrometer during a flyby analysis of a debris cloud created by an artificial impactor. As a secondary objective and a descope the objective of taking close up pictures of Deimos has also been included. The motivation for this mission come from the decadal survey; “...There has been little mention of missions to the martian moons, Phobos and Deimos. It is clear that these two moonscould play important roles in the future exploration of Mars, especially if they turn out to be related to volatile-rich asteroids, a possibility which has not been excluded by existing data and observations.If so,they may be the surviving representativesof a family of bodies that originated in the outer asteroid belt or further, and reached the inner solar systemto deliver volatiles and organics to the accreting terrestrial Planets….measurements of bulk properties and internal structure, high- resolution imaging of surface morphology and spectral properties, and measurements of elemental and mineral composition.” Mission Overview
- 2. Utilizing a 6U cubesat and the Mars 2020 trajectory, we plan to release an impactor into Deimos to release a large dust plume. This plume will be analysed during a flyby with the spectrometer on board our cubesat. In addition, our cubesat will take images of Deimos to improve the understanding of it’s geography. This will also serve as our descope Both of these objectives fit well with the desired science objectives of the Decadal survey. The driving requirements for our mission revolve around creating a plume large enough for analysis and likewise being able to analyse a plume to determine the composition of Deimos. This affects the size of our cubesat because the impactor must be sufficiently large as to have enough kinetic energy at the velocities the probe is traveling, while keeping the total mass within the requirements. The spectrometer on the cubesat also must be sufficiently powerful to measure the composition of the plume, while also being small and light enough to still meet the mission requirement. Trajectory The initial trajectory of the Mars 2020 craft has not been fully disclosed yet so a few assumptions were made. Based on the official Mars 2020 Mission Concept powerpoint (12), the transfer orbit was approximated and slightly altered to provide a good approach to rendezvous with Deimos. There are two options for encountering Deimos: prograde or retrograde. Figure 1 shows the approximate approach for a prograde encounter and Figure 2 shows a retrograde approach. Fig. 1. Prograde Fig. 2. Retrograde In the event of being off-phase with Deimos, a ΔV will be performed at Mars’ sphere of influence (5.9x105 km away) to either slow down or speed up to converge with the moon. The maximum ΔV required to get back in phase from a maximum phase shift is 257 m/s. That is achievable by our cubesat. To calculate the speed at which our impactor will strike Deimos, the velocity of the spacecraft, Deimos, and any ΔV will need to be accounted for. In both cases, the velocity of the spacecraft is 2.4 km/s thanks to the gravitational pull from Mars, resulting in 3.75 km/s for retrograde and 1.05 km/s for prograde. These calculations were determined from GMAT. The difference in these two impact velocities greatly influences how effective the impactor will be. Kinetic energy is proportional to velocity squared, so nearly quadrupling the velocity from a prograde to a retrograde encounter results in over 12 times the impact force. This would be more effective by providing a bigger dust cloud to analyze, but it would also require much more precision regarding trajectory and pointing to collect all the data while travelling so quickly. Instruments: Morphological Matrix
- 3. - Architecture Decision Option 1 Option 2 Option 3 Option 4 Communication EWC 27 HDR-TM X Band Transmitter MarCO Spectrometer Compact Ion and Neutral Mass COSIMA Amptek X- 123SDD Propulsion System Chemical Electric Cold Gas Hybrid Chemical Hydrazine AF-M315E Attitude Control Reaction Wheels Cold Gas Electrospray Attitude Determination Star Trackers IMU Sun Sensors Command & Data Handling iOBC Cube Computer - Impactor The feasibility of the impactor was calculated using data of the Deep impact mission which impacted comet tempel-1. Deep impact, impacted with a velocity of 10 m/s and a mass of 300 kg and ejected 1.2*106 kg of material. Our impactor will hit with a 1/1000 of the kinetic energy and eject 1/1000 of the mass. This is still however 2550 kg of material. We believe this will be sufficient to analyse as it will make a 2.5 km3 cloud at a density of 1 kg/m3.(4) The impactor is heavily driven off of the driving requirement of having a plume large enough to be easily pointed at yet still dense enough to be analysed by the spectrometer. The mass of 3 kg was chosen in conjunction with trajectory to generate enough impact force as the force scales linearly with the mass. During the last portion of the flight the impactor will fly independently of the cubesat therefore will require its own short term propulsion, power, and ADCS, these are listed below. Spectrometer The spectrometer selected for our mission to Deimos was the X-123SDD complete x-ray spectrometer with silicon drift detector (SDD). The X-123SDD spectrometer was selected due to its small size (7 x 10 x 2.5 cm), low power requirement of 2.5 W, light weight of 180 grams, and a transmission
- 4. rate of 9.6 kb/s. The Compact Ion and Neutral Mass Spectrometer was not selected because of its larger weight (560 grams), about three times as heavy although it did have a smaller power requirement. For this decision we decided to sacrifice .9 Watts of power for a lighter spectrometer. The COSIMA was not selected because of it required 20.4 W which was more power then we could supply for the spectrometer alone. The COSIMA was also much larger in size with the following dimensions 39.4 x 97.3 x 37.8 cm. The X-123SDD complete x-ray spectrometer’s price, set by Amptek incorporated, is approximately 41,600 US dollars. It’s is currently being applied on the Miniature X-ray Solar Spectrometer mission launched on December 6, 2015. Propulsion The main propulsion system for this mission will require a large ∆V for a CubeSat mission. Initial calculations project this value to be approximately 250 m/s. The nature of the mission confines the orbital maneuvers to high thrust in relatively short time periods. A high degree of mobility will be needed to ensure the CubeSat can successfully position itself within range to collect data. This constraint suggests the elimination of electric and solar sail as feasible propulsive methods, as they require long burn times to achieve high ∆Vs. This leaves two options: chemical and cold gas. Inert gas propulsion has the advantage of simplicity but with an average specific impulse of 70 seconds, will not produce a sufficient ∆V with the given 6U volume constraint. This leaves chemical propulsion as the best option for the mission. Bipropellant systems are too mechanically complex to scale down to a CubeSat level, therefore a monopropellant system will be used. The propellant known as AF-M315E was selected over the hydrazine for several reasons. Most importantly, AF-M315E has a significantly lower toxicity level when compared to hydrazine. This will reduce range operations costs, as loading procedures can occur more rapidly and will eliminate the need of the waiver that is normally required with hazardous, secondary payloads. Additionally, AF-M315 is 45 percent denser than hydrazine which allows for smaller propellant tanks. It also has a lower freezing point than hydrazine which reduces the complexity of thermal management subsystem. This propellant has an expected specific impulse of 240 seconds. Aerojet Rocketdyne’s Modular Propulsion System (MPS-130) will be used as the main system. It will occupy 2U (10 cm x 10 cm x 22.4 cm) and have a wet mass of 3.5 kg. The amount of propellant it will be able to hold is 1.3 kg resulting in a 2.2 kg dry mass. Assuming all other systems on the spacecraft will total 5.5 kg, the maximum expected ∆V is 367 m/s for this configuration. This is an ample amount for the mission and will allow for 40 percent margin, in case emergency maneuvers are required. This system has a technology readiness level of 6 and will be flight tested in early 2017 when NASA’s Green Propellant Infusion Mission (GPIM) launches. This mission is intended to demonstrate the viability of using AF-M315E as an effective propellant for orbital maneuvers. System Operations Attitude Determination and Controls The primary method of attitude control and determination on the CubeSat will be through the use of thrusters and the Blue Canyon Technologies (BCT) XACT Attitude Control System. This system was determined to be the optimal attitude control system through a trade study consisting of four different systems (as shown in Table I) and comparisons between different attitude control and determination
- 5. systems in the morphological matrix. While not the cheapest system, it provides the pointing accuracy and components that are the most optimal for our mission design. The BCT XACT system contains Star- Trackers,three Reaction Wheels, and the electronics and software to run the system. Since Deimos has no magnetic field, the Magnetometer and Torque Rods that are part of the MAI-400 system are useless and the ADS will use star trackers as they will be more accurate than sun sensors. Also, since we are not near Earth, the IR Earth Horizon Sensors on the MAI-400 will not be useful. One of the biggest perks of the BCT XACT system is that it is currently being flown on the MinXSS cubesat mission and has been chosen to fly on one of the first interplanetary cubesat missions, MarCO. The flight pedigree of this control system makes this the easy choice for our primary ADC system. The best alternative is the BCT XACT-50 due to its similar specifications and components. Table I. Attitude Control System for the CubeSat BCT XACT BCT XACT Lite BCT XACT-50 MAI-400 Pointing Accuracy (deg) ±0.003, ±0.007 1 ±0.003, ±0.007 N/A Lifetime (Years) 3 3 3 N/A Mass (kg) 0.91 0.7 1.225 0.694 Volume (cm^3) 10x10x5 10x10x5 10x10x7.54 10x10x5.16 Electronics Voltage (V) 5 5 5 5 Reaction Wheel Voltage (V) 12 12 12 N/A Slew Rate (deg/s) 10 10 10 N/A Cost (FY17) $145,123 N/A N/A $72,559 Flights MarCO, MinXSS N/A N/A N/A Contains Star-Trackers Sun Sensors Star-Trackers 3 Reaction Wheels Reaction Wheels Magnetometers Reaction Wheels Magnetometer Electronics IMU Torque Rods 2 IR Earth Horizon Sensor Software Reaction Wheels Electronics 3 Torque Rods
- 6. Electronics Software Computer Software The methods of attitude control and determination on the impactor utilize some of the same components as the main cubesat, however, will be reduced to a smaller package. While the BCT XACT system is 0.5 U, to cut back on size we will focus on the use of a star tracker, an inertial measurement unit (IMU), and three reaction wheels to allow for 3-axis control. This will allow for enough attitude control even with the omission of a propulsion system. As the pointing of requirements of the impactor are less precise than that of the cubesat, all of the components of the BCT XACT system are not needed and this can also provide a cost savings. Much like the CubeSat, the use of magnetometers and torque rods are unnecessary, so only ADCS components that do not rely on magnetic fields can be used. The reaction wheels that were chosen (Blue Canyon Technologies RWP050) will have a maximum torque of about 0.007 Nm with a power at full momentum of 1 W. The design life is 5 years. The star tracker (Blue Canyon Technologies Thin Slice NST) will have an attitude solution of 5 Hz, Sky Coverage of over 99%, and peak power of less than 1.5 W. The IMU chosen is the SMG Ellipse-A with 0.2 degrees of roll and pitch accuracy over 360 degrees. Power The CubeSat will rely solely on Solar panels to generate power. It will use two Clyde Space Ukube-1 Single Deployable, Double-Sided Solar Panels. These will line the outside of the cubesat and fit within our structure. This will provide 40 W of peak power at Deimos (assuming 48% solar irradiance of Earth) and 20.8 W of Orbital Power. The power system will be managed by a Clyde Space XU Cubesat EPS, which is up to 98% efficient regulating 3.3V and 5V Busses. Also, the cubesat will use a 12V 18650B Lithium Ion Batteries with custom battery protection circuitry similar to the one used in MarCo. For the impactor a 3000mah battery will be used over solar. The short time that it needs to be used couple with the small size of the impactor makes solar infeasible. Communication There are two viable and very different approaches available in terms of communication, the first is mimicking the transmission system for the MarCO cubesat, launching in 2018. It utilizes an X-band transmission at 8 kb/sec, it is large and requires far more power than option two, around 20 W, but it will allow bypass of MRO and direct downlinking to DSN on Earth. Option two is the EWC 27 HDR-TM X- Band Transmitter, developed by SyrLinks. It has a customizable downlink speed, 5 Gb/s the option for the 3 m dish on MRO. It requires 10 W of power, and is far smaller and lighter than option one. With the downlink requirement of 10 kb/s, option two is a superior choice. It is also TL9 mission capable. Command and Data Handling For the control of the data and execution of commands, we have currently selected two possible computers. The first one is the ISIS On Board Computer (iOBC) with a flight heritage starting in 2014. The second one is the Cube Computer.
- 7. The iOBC runs with a 400 MHz 32-bit processor on a 3.3 V power supply. Its mass is 94 grams, and has dimensions 96 by 90 by 12.4 mm. It is designed to either have two 8GB high reliability SD cards or two standard SD cards of any size. Since the iOBC has heritage, its technology readiness level (TRL) is roughly 6 to 7. The Cube Computer runs with a 4-48 MHz 32-bit ARM Cortex-M3-based MCU on a 3.3 V power supply. It has a 2 GB MicroSD socket for storage. It weighs 50 to 70 grams, and has dimensions 90 by 96 by 10 mm. The Cube Computer is newer so its TRL is 6 at best. Thermal Management Structure will be the biggest contribution to thermal management on the cubesat. Utilizing a thicker structure around the payload will spread heat better. Because heat is transferred in space via radiation, removing points of contact to prevent unwanted conduction will also be done. This can be done utilizing tray inserts for payload. Materials will also be taken into account. Radiators placed over hot components will be used to disperse heat to other components. The Sun’s thermal energy will also allow heating of components during transit. Conclusion After multiple trade-off studies, the most feasible configuration of components and trajectories has been determined. The instrumentation for the mission has been found to be weight and volume efficient, a challenge considering that most cubesat technologies are not currently developed for Mars operation. The greatest challenges in weight and volume were the impactor, propulsion, and power. These rely on innovative but very possible technology. The primary science objective will be met with the impactor/spectrometer combination. The secondary objective will be met with the onboard camera. Communications will be relayed through MRO over the DSN.
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