Fall 2017 Fuel Resupply Mission Final Report

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Fall 2017 Fuel Resupply Mission Final Report 401A – Preliminary Spacecraft Design Team 7: ResuppLIONS Team Members: Adam DiPillo, Alejandro Genis, Kevin Cahill, Nick Smith, Santino Zangaro, Takugo Amy

Abstract Orbital resupply represents one of the most important next steps in space mission research. The mission to Mars will require propellant resupply while the Transport vehicle remains in-orbit. Orbital resupply remains relatively untouched, with use only on the ISS, in the form of Progress M1. ISS refueling does not accurately reflect the extensive capability that would be required for a Mars mission. The “Asclepius” mission by “Team ResuppLION” therefore intends to provide a complete mission design that implements orbital resupply of hybrid fuel for a Mars mission spacecraft. Asclepius will support the Transport craft designed by University of Texas at Austin planned to launch in 2033. As this project is in progress, this report includes the amount of research, development and design completed, thus far. Structures and Mechanisms provides an overall structure for the spacecraft that stands up to the stresses of the mission. It also interfaces with the target spacecraft using a robotic arm and the docking standard. In order to launch Asclepius into a translunar trajectory, a super heavy launch vehicle is needed, such as the SLS Block 2. The Block 2 has a launch capability of 130,000 kg and uses cargo fairings that support up to a payload width and height of 10 m and 27.43 m, respectively. These capabilities satisfy Asclepius’ launch requirements, with the main setbacks being each Block 2 launch will cost approximately $500,000,000 and the amount of launches is limited to three per year. Other important parameters that will be analyzed, are the translunar trajectory injection accuracy and the launch dates.

Propulsion used the previous decisions of using Monomethyl Hydrazine and Nitrogen Tetroxide to select an engine. The new Super Draco thruster that produces a thrust of about 71,000 N will be used as the primary type of engine. The massive amount of weight that needs to be transported to the moon has produced a huge amount of required thrust. The amount of thrust varied from 2,137,342.9 N to 774,399.6 N. The number of engines used were estimated using V estimates the burn time of the engines and the amount of thrust needed. The maximum amount of V with the shortest time of flight required 30 Super Draco engines. The minimum amount of V required a 11 Super Draco Engines, which was a more reasonable number.

Several ground stations were identified for the mission, the mission control center and space flight operations facility will be located at the Jet Propulsion laboratory, while the payload operations control center will be located at NASA Goddard. Communications will use a 0.6m wide high gain parabolic dish using S-band frequencies to communicate with ground control through the Deep Space Network. The RAD5545 SpaceVPX single-board computer will be the workhorse to process and distribute the spacecrafts' data. The system architecture will be constructed in a bus configuration that will dynamically operate, varying with each mission phase. Through environmental protection, redundancy, timekeeping and memory allocation with the proper devices, this system will avoid many of the common flaws that can pop up in space missions.

Guidance, Navigation, and Control will be capable of pointing and guiding the spacecraft during all phases of the mission using thrusters and a specialized sensor suite.

Power produced an array size that will be used as the primary source of power. The array size will need to be about 5m2 to supply 6000 W of power using the Azure Space 30% Triple Junction cells. A secondary power source was also selected. A Lithium-Ion battery from ABSL named 8s52p 28V will be used as the secondary battery source. It will be able to produce about 2000 W of power which will be more than enough for normal operating power. Two batteries will be used in the event that one were to fail or if the spacecraft would have to be under the secondary power source for an extended amount of time.

Through preliminary thermal analysis of Asclepius, it was determined that the spacecraft will require a combination of active and passive control methods. Some of the control methods will include: multilayer insulation (MLI), electric heaters, louvers, and coatings. The current design is to use MLI on the command module to prevent excessive heat loss or absorption during all phases of the mission. When exposed to the Sun, louvers will be utilized to radiate heat out of the command module. When in an eclipse period of orbit the MLI may maintain operational temperatures inside the command module but electric heaters may need to be utilized. The detailed analysis to be completed will help determine if these active methods are necessary. A coating on the fuel tank may be needed depending on the final material choice and that materials properties. Because Asclepius utilizes different instruments to accomplish various tasks over the course of the entire mission, an analysis with regards to what type and how many of these instruments will be used must be done. Due to the ResuppLION missions’ lacking a scientific objective, the specialized instruments used by Asclepius will mainly consist of equipment for the Communications subsystem. Additionally, in order to take advantage of the Space Launch System’s Exploration Upper Stage payload services, Asclepius will be using a customized payload adaptor that consists of a Payload Attach Fitting, to act as the structural interface to the Exploration Upper Stage, and a Payload Separation System, to act as the separation interface for the payload mounted on the Payload Attach Fitting.

The end result will be a complete mission design and plan from launch to mission end.

Acronyms LDRO Lunar Distant Retrograde Orbit GN&C Guidance, Navigation, and Control MMH Monomethyl Hydrazine NTO Nitrogen Tetroxide PVMQ Phenyl Vinyl Methyl Silicon Rubber NDS NASA Docking System SLS Space Launch System LEO Low Earth Orbit PPL Primary Payload PLF Payload Fairing TLI Trans-Lunar Injection NRHO Near Rectilinear Halo Orbit EUS Exploration Upper Stage KSC Kennedy Space Center PAF Payload Attach Fitting PSS Payload Separation System GaAs Gallium Arsenide Cell STC Standard Temperature Conditions BOL Beginning of Life EOL End of Life LDHEO Lunar Distant High Earth Orbit LGA Lunar Gravity Assist WSB Weak Stability Boundary STK Systems Tool Kit OMS Orbital Maneuvering System C&DH Command & Data Handling RAM Random Access Memory ROM Read-Only Memory UART Universal Asynchronous Receiver Transmitter SMAD Space Mission Analysis & Design I2C Inter-Integrated Circuit (pronounced I-squared C) MHz Megahertz

iii kHz Kilohertz I/O Input/Output SPI Serial Peripheral Interface BAE British Aerospace MLI Multilayer Insulation MCC Mission Control Center SOCC Space Craft Operations Control Center POCC Payload Operations Control Center

Table of Contents Abstract ...... i Acronyms ...... iii Introduction ...... 1 Mission Requirements, Concept of Operations, and Planned Trajectories ...... 3 Requirements ...... 3 Concept of Operations ...... 4 Planned Trajectories ...... 4 1. Structures ...... 6 Asclepius Architecture ...... 6 Fuel Transfer ...... 6 Berthing with Target Spacecraft ...... 8 Collapsible Tank ...... 9 Collapsing Actuators ...... 12 Docking Module & Robotic Arm ...... 13 Command Module ...... 13 Weight Estimation ...... 13 2. Launch Vehicle ...... 15 3. Propulsion ...... 18 4. Ground Control ...... 24 5. Communications ...... 25 6. Command & Data Handling ...... 26 Hardware Design ...... 26 Software Design ...... 27 System Architecture ...... 27 On-board Computer ...... 28 Mission Phases ...... 29 7. Guidance, Navigation & Control ...... 30 Attitude Determination and Control ...... 30 Orbit Determination and Control ...... 32 8. Power ...... 34 9. Thermal Control ...... 40 10. Payload ...... 42

Summary Tables ...... 43 Weight, Cost, and Power Estimates ...... 43 Recommendations for Future Work ...... 44 Structures and Mechanisms: ...... 44 Launch Vehicles:...... 44 Propulsion: ...... 44 Space Payloads: ...... 44 Ground Control: ...... 44 Communications: ...... 45 C&DH ...... 45 Power: ...... 45 Thermal ...... 45 Conclusions ...... 46 Appendix ...... 47 References ...... 50 Launch Vehicle ...... 51 Power: ...... 51 Ground Control ...... 51 Communications ...... 51 Command and Data Handling ...... 52 GNC ...... 52 Power ...... 53 General ...... 53

Introduction ResuppLION is currently working on theme three of the NASA RASC-AL competition to aid in the future plans of transporting humans to Mars. The purpose of this theme is to design an in- space transportation system that will supply another spacecraft with electric and chemical fuel. The primary requirements given for this mission are:

 Refueling of the Deep Space Transport occurs in cis-lunar space via propellant launched from Earth, on SLS or any other launch system

 Propellant resupply must support one crewed orbital Mars mission every other opportunity beginning in 2033 and Mars surface mission cargo delivery beginning in 2037, with the ability to be expanded to support crewed Mars surface missions later

 Nuclear-powered propulsion concepts are not allowed.

Theme three also has the opportunity to work with another team. ResuppLION decided to work with the University of Texas at Austin. Texas is working on theme one of the competition. They were tasked with designing a reusable hybrid propulsion stage that will transport a baseline habitat with crew and logistics to Mars. ResuppLION will transport fuel to the hybrid propulsions state via Asclepius (resupply spacecraft) in Lunar Distant Retrograde Orbit (LDRO).

This mission is a crucial first stage that will enable the transportation of humans to Mars. It requires integration of various subsystems to ensure that the spacecraft, Asclepius, functions correctly and is able to complete its mission. Asclepius will have seven major modules used, structures is handling the integration of those seven modules which includes the docking station, the robotic arm that will be used for berthing, the collapsible tank used for transferring fuel to the Hybrid Transport, and others. Launch Vehicles is responsible for placing Asclepius in a necessary orbit that will minimize future maneuver corrections. Additionally, since multiple launches will be required, launch timing and scheduling of the ResuppLION missions is also being integrated into this subsystem. Propulsion is handling the type of engines that will be used on Asclepius as well as the amount of V required. Ground Station is handling all aspects of the mission on earth from the command center to analysis the data returned from the spacecraft. Communications will be managing the interface between Asclepius, earth, and the reusable hybrid propulsion state. Command & Data Handling (C&DH) will be looking into the communication between every subsystem to appropriately handle and deliver commands. Guidance Navigation and Control (GNC) is exploring the major phases of attitude control during the mission as well the orientation hardware used onboard. GNC is also exploring the protocols and steps that need to be taken during certain phases in the mission. Power is handling the supply of power to the spacecraft. Power primarily looked into the solar cells, panels, and batteries that will be used for the missions. Thermal Control tackling the regulation of temperature of the spacecraft. Thermal control is looking into the different temperatures that must be maintained throughout the spacecraft. Space Payloads outlines the various payloads Asclepius will be using in order to satisfy its mission objective.

The success of the mission is contingent on the integration success of these subsystems. As the preliminary design phase keeps advancing various aspects of the mission are subject to change. Up to this point each subsystem has provided an initial assessment of how their subsystems will ensure that Asclepius has a successful mission.

Mission Requirements, Concept of Operations, and Planned Trajectories Requirements 1. Delivers fuel to spacecraft in cis-lunar space  Our partners at University of Texas have requested a rendezvous in Lunar Distant Retrograde Orbit. This is an orbit around the moon to be used as a refuel point before their transfer to mars. 2. Capable of fueling a Mars mission at every other opportunity starting in 2033  This gives 4 years and 4 months between successive ResuppLion missions (twice the synodic period of Mars with respect to Earth) 3. Capable of transporting both Electric and Chemical Propellants  The target spacecraft will use a combination of chemical and electrical propellants. Team ResuppLION will need to resupply both of these. Special care will be needed for the chemical propellants because these will have some risk of coming in contact and igniting. ResuppLION’s partner team at University of Texas is requiring Monomethyl Hydrazine, Nitrogen Tetroxide, and Xenon Fuel. 4. Onboard technologies must be fully ready for deployment by 2029.  This is approximately 4 years before the first mission. All technologies must be fully developed before then and a finalized design must be finished. It’s important for the success of the Mars mission to have this fully developed well in advance. Team ResuppLION will aim for a test flight as a “technology demonstration” flight sometime in 2030. 5. Deliver “many” metric tons of fuel to target spacecraft  This was interpreted to be approximately 65,000 Kilograms of fuel. This number comes from the hybrid propulsion theme of the RASC-AL competition. The target spacecraft to be resupplied should be similar to this. How this weight will be partitioned by fuel type is still flexible as University of Texas continues to iterate on their mission design.

Concept of Operations

Figure 0.1

The target spacecraft will launch first and be placed into LDRO where it will wait for a fuel resupply. Asclepius will then be launched and deployed into a translunar orbit. Asclepius will then perform a few maneuvers to bring it into a lower LDRO than the target spacecraft. Asclepius will trail behind the target in this lower altitude orbit. The faster orbital period will allow Asclepius to catch up to the target spacecraft. At this point, Asclepius will begin its approach to the target spacecraft. The two will rendezvous in the higher LDRO where Asclepius will transfer its fuel. Once all fuel has been transferred, Asclepius will depart and perform a maneuver which will bring its periapse back down into low earth orbit. The altitude of periapse will be inside the atmosphere so that aerobraking will bring the apoapse down to a level where reentry can be safely performed. Aerobraking will not be used to completely deorbit the spacecraft because this would cause an unpredictable landing area. Instead, a final maneuver will be performed before natural reentry to put the spacecraft on a desired descent trajectory into the ocean where it can be safely disposed.

Planned Trajectories

The launch vehicle will deploy Asclepius into a translunar injection orbit From there, Asclepius will perform a transfer that will be bring it into LDRO. The target orbit will have Asclepius trailing behind the target at a slightly lower altitude. This orbit transfer will need to be timed so that injection into this trailing orbit has Asclepius trailing behind the target where it will take a few days for it to catch up with the target. More details on this transfer are in the propulsion section. Once the craft naturally catches up to the target, Asclepius will begin a close approach

4 where it will match velocity with the target and work its way up radially toward the target until the spacecraft mate occurs.

At any point along the approach, Asclepius must be capable of safely aborting away from the target. Since Asclepius is approaching from below, if the spacecraft fails completely due to some system failure, the orbit will cause it to naturally fall away from the target. Alternatively, if the mission requires Asclepius to perform an active abort that allows for a second chance at rendezvous, it could execute a delta V burn that will bring its apoapse higher than the target spacecraft as seen in figure 0.2. This will cause Asclepius to move past the target and above the target. This abort orbit will have a similar period to the target orbit and so when the craft come back around, Asclpius will be set up for a second attempt.

Figure 0.2

Once Asclepius departs from the target, it will move radially away from the target and allow it to drift a short distance away before beginning its descent back to earth. At this point, Asclepius will execute a burn that brings its periapse down into earth’s atmosphere. Asclepius will then use aerobraking to bring the apoapse down as seen in figure 0.3. The purpose of this is to create a slower and safer reentry velocity. To bring the apoapse down to a level that is close to circular is expected to take several passes.

Figure 0.3

1. Structures Asclepius Architecture

Figure 1.1 Asclepius has six major modules that are used. They are laid out as seen in figure 1.1. Each section contains instrumentation and mechanisms related to its respective purpose. This gives the spacecraft a modular build for ease of testing and integration. The docking module will route the fuel into the docking station for transport to the other craft. This segment also contains the physical docking mechanism and the robotic arm that will be used for berthing with the target spacecraft. The resupply tanks contain the fuel that will be transferred. The electric propellant will be placed in the center tank. Electric propellant will be a chemically inert fuel (xenon) and so it will act as a safety barrier between the fuel and oxidizer to minimize any chance of fuel and oxidizer mixing in the event of a leak. Collapse actuators and structural panels will serve as the physical mechanism by which the fuel will be transferred. The command module will contain all the mission systems and sensors that control Asclepius. This includes all hardware for command and data handling, power electronics and solar panels, communication hardware, and GN&C instrumentation. A propulsion unit with an engine that can be turned on and off will be used for all orbital maneuvers. A propulsion unit with an engine that can be turned on and off will be used for all orbital maneuvers. While the exact number of engines and their locations on Asclepius are currently being research, more information on the types of engines being considered can be found in the Propulsion section of the report.

Fuel Transfer Fuel will be transferred via collapsible tanks. The material will be made of a flexible silicon material. This was chosen because it will be able to withstand the pressure and remain flexible at the temperature the fuels will be held at. This may have issues with permeability which will result in pressure and fuel loss over time. To combat this, the tank will be kept at an optimal temperature and be of a thickness that decreases the permeability and diffusivity of the propellants with the

6 silicone rubber to an acceptable value. More details on the permeability of the silicon rubber tank can be found in the Collapsible Tank sub section.

Figure 1.2 The fuel will be sent over at a controlled flow rate using feedback control as shown in figure 1.2 The controller will control the mass flow rate using the pressure sensors. The walls of the tank will have a maximum static pressure before damage or rupture occurs. Because of this, an additional safety measurement of the static pressure will be taken to limit the output of the collapse controllers. This will prevent the collapse actuators from over pressurizing the tanks if the commanded mass flow rate results in a static pressure that exceeds a safety range.

Figure 1.3

For this model of feedback control to be implemented, Asclepius will need some way of measuring the fuel flow rate through the lines. Fuel is approximately incompressible and the tank diameter is much wider than a fuel line. Because of this, Bernoulli’s equation can be applied to get a mass flow rate equation (Equation 1.1)[1.1].

푚̇ = 퐴√2𝜌∆푃 Equation 1.1

The controller will use digital logic implemented on the primary processor on Asclepius (more details on the processor in the C&DH section of this report). This will make the collapse control logic essentially a subroutine of the main flight software. This reduces the build complexity and number of components used to build Asclepius. The actuators and sensors will receive commands from the flight computer via one of the I2C busses on the processor.

Berthing with Target Spacecraft The interface between the two spacecraft will use the International Docking System Standard[1.2]. This is an androgynous docking system designed to be used with human spaceflight systems. It is intended to be used as a standard for mating spacecraft in the future for both docking (no arm) and berthing (robotic arm assisted). This system allows for an easily understood interface between Asclepius and the spacecraft produced by University of Texas. This system can also be used for crew evacuation. In the event that Asclepius fails, the crewed missions could be evacuated and returned to earth. This decision is up to the engineers at University of Texas; the option is just a factor in this choice. A robotic arm will be used to assist the spacecraft mating. This system was chosen because it’s safer than free flight docking between very massive spacecraft. The arm will be located on the docking module of Asclepius and will attach onto a specialized hard point on the target spacecraft. Arm assisted berthing has been used on the resupply missions to the International Space Station[1.3].

Collapsible Tank

As stated before, Asclepius will be using a collapsible fuel tank system designed to store and transfer the three Deep Space Transport propellants: Monomethyl Hydrazine, Nitrogen Tetroxide, and Xenon. As a result, the fuel tank must be designed with a material that can withstand the imposed collapsing strain and prevent leakage.

The optimal material thus far has proven to be silicon rubber, which is a type of elastomer [1.4]. This is due to several reasons. Generally, silicon rubber has low chemical reactivity and toxicity, and its properties are constant over a wide range of temperatures [1.5]. A specific advantage, though, is that silicon rubber’s chemical structure can be changed by adding certain molecules to its primary polymer chain. Changing its chemical structure can affect interesting properties. One property that can be controlled is the silicon rubber’s glass transition temperature (Tg), which is the temperature at which elastic materials act brittle. By making phenyl, vinyl, and methyl substitutions on the polymer chain, the silicon rubber’s Tg can be lowered to -120° C, the lowest of all elastomers [1.4-1.5]. This type of silicon rubber is called PVMQ and it is able to sustain its large flexibility at very low temperatures, which is key for this collapsible tank system.

Unfortunately, PVMQ is known to be one of the most permeable types of elastomers. This is due to “free volume” or “holes” that “thermally form and disappear with the movement of polymer chains” [1.4]. When the rubber comes into contact with a fluid it proceeds to diffuse through the membrane, via the free volume holes, until it has fully passed through the rubber. The fluid then resumes its state before it entered the membrane. This phenomenon is in agreement with Henry’s law: 푐 = 푆 ∗ 푝 (1.2)

Where, c represents the concentration of molecules in the rubber, S is the solubility in the rubber, and p is the pressure of the fluid [2.4]. In aerospace applications, the rate of permeation can be characterized by the permeability of the membrane, which can be denoted as:

푣∗훿 푃 = (1.3) 퐴∗푡∗(푝1−푝0) Where 푃 represents the permeability for a given fluid in a given membrane, 푣 is the volume which penetrates the membrane, 훿 is the membrane thickness, 퐴 is the area of the membrane, 푡 is the time, 푝1 is the partial pressure of the fluid on the higher-pressure side of the membrane, and 푝0 is the partial pressure of the fluid on the lower-pressure side of the membrane [2.4]. There are a few factors that must be considered when combatting permeability in silicone rubber. One factor is the type of fluid inside the tank, as solubility and diffusivity differ between types. It has been demonstrated that the permeability of silicone rubber is affected by the solubility more so than the diffusivity of a given fluid. This is because if a molecule has a similar polarity to silicone, it will have a higher solubility in silicone rubber. If it has a dissimilar polarity, it will have a lower solubility in silicone rubber, and therefore, the permeability will be lower even despite a high diffusivity [1.4-1.5].

Temperature also has a strong effect on permeability as silicone rubber’s free volume depends on it. Low temperatures result in less free volume and lower polymer chain mobility, which results in lower permeability and restriction of diffusion, respectively. In addition, lower temperatures result in slower molecular movement, which in turn ultimately causes less diffusion. When the ambient [1.4] temperature is lower than the polymer’s Tg, permeability is reduced significantly .

In addition, the solubility of a fluid in silicon rubber is dependent upon the temperature and enthalpy of solution, ΔHs, as shown below: −∆퐻푠⁄푅푇 푆 = 푆0푒 (1.4)

Where S represents the solubility, S0 is “pre-exponential factors”, R is the gas constant, and T is temperature in Kelvin. One can see that if the enthalpy of solution is negative, solubility increases with decreasing temperature, while if the enthalpy of solution is positive the opposite is true [2.4].

While permeability is independent of factors such as pressure, and membrane thickness and area, the amount of fluid that diffuses through the membrane depends on these variables. Essentially, a lower pressure differential, and a thicker, smaller membrane area will result in less overall diffusion [1.4].

In order to adapt PVMQ for low permeability applications, the silicon rubber must be modified. Fortunately, Arlon Silicone Technologies Division has developed “proprietary technology” that modifies PVMQ so that it still retains low temperature flexibility, but also has a lower permeability than PVMQ [1.4].

In The Permeability Characteristics of Silicone Rubber by Haibing Zhang, experimental results provided by Arlon are analyzed. One experiment in the report involves the permeability of different silicone rubbers with helium at 23 °C. Using sample sizes that weighed four ounces per square yard, the permeability of Arlon modified PVMQ was 49.2 liters per meter squared per twenty-four hours per atm. This is a significantly lower permeability than non-Arlon modified PVMQ, which was found to be 130 liters per meter squared per twenty-four hours per atm [1.4].

Using the testing method ASTM 6182, the low temperature flexibility of the different silicone rubbers was also studied. The flexibility of a membrane was characterized by its elongation during testing when held at –100 °C. In the resulting stress-strain curve, the Arlon modified PVMQ had “approximately 30%” elongation and did not rip during testing. On the other hand, non-Arlon modified PVMQ had an “excellent elongation of approximately 500%” [1.4].

With all considered, the Arlon modified PMVQ will be the specific type of silicon rubber used for the collapsible tank material. This is because the Arlon modified PMVQ “optimizes the compromise between low permeability and good low temperature flexibility” [1.4], which would be very advantageous for the ResuppLION missions.

Based on the amount and types of propellant the University of Texas needs, a preliminary design of the collapsible tank can made. Assuming fuel residuals of 3.5% for MMH and NTO and 2% for Xenon, Table 1.1 summarizes the propellant specifications.

Fuel Type Density (kg/m3) Weight (kg) Volume (m3) Xenon 1500 19584 13.06 Monomethyl hydrazine 880 17595 19.99 Nitrogen Tetroxide 1440 27738 19.26 Total 64917 52.31

Table 1.1 Fuel Weight and Volume Estimates

Since the collapsible tank is to be cylindrical in shape, the optimized diameter and height that will efficiently use minimal amount of silicon rubber are 4.06 and 4.04 meters, respectively. Of course, the thickness of the tank walls will not be negligible, so these are approximate dimensions for the tank structure. Additionally, structure panels that are attached to the collapsing actuators will between each “sub tank” section to prevent the walls that are applying the pressure from deforming. A proposed material for these panels is Carbon Fiber because of its high stiffness and light weight [1.6]. At the moment, there will be four-disc shaped structural panels, each with a radius equal to that of the silicone rubber tank's and a thickness estimated at .003175 m, installed throughout the tank. One panel will be at the top and bottom of each sub tank, with each sub tank sharing a top and/or bottom structural panel with the adjacent sub tank.

Due to the propellants being stored separately, the temperature and thickness of the tank sections can be optimized based off the propellant stored inside.

One can begin this optimization analysis by recalling that solubility is the dominant variable in determining the propellant’s permeability with silicone rubber. So, if the propellant is initially very soluble with silicone rubber, greater measurements must be taken to prevent leakage.

Fortunately, one can decrease the solubility if it is too high. Recalling equation (1.4), if the propellant has a positive enthalpy of solution it will become less soluble with silicone rubber as temperature decreases, and consequently less permeable according to (1.2). So, the analysis can continue by finding out the sign of MMH’s, NTO’s, and Xenon’s enthalpy of solution with silicone rubber. Based on this, one can know how changing the temperature of the sub tank will affect solubility. If the propellant has a negative enthalpy of solution, then its solubility, and therefore permeability, with silicone rubber decreases with increasing temperature. In this event, there is no trade-off between flexibility and permeability as silicone rubber’s flexibility also increases with increasing temperature [1.4-1.5].

When the enthalpy of solution is positive though, a trade-off between flexibility and permeability occurs with a changing temperature. One can begin by defining the temperature ranges in which each propellant is a liquid. These are the ranges at which each sub tank can be kept, respectively. Then, an initial temperature can be selected and incrementally decreased until the permeability of the gas is acceptable. Once this temperature is found, the silicone rubber’s flexibility would then be tested. If the flexibility is not acceptable, then the temperature could be increased again if possible. In the event that the temperature cannot be increased, the silicone rubber’s wall thickness can then be decreased to increase flexibility. Of course, if one recalls equation (1.3), permeability is normalized by thickness. So, while the value of permeability is independent of thickness, decreasing the wall thickness will increase the total propellant diffusion across the membrane [1.4].

So, the wall thickness can be decreased, to increase flexibility, without adding a significant amount of propellant loss if the solubility is held at a low value to decrease permeability [1.4]. This is an iterative process that must be repeated until the selected temperature and wall thickness enables acceptable performance. In the event that one cannot find a wall thickness that enables acceptable flexibility and diffusivity, the walls of the tank must be structurally modified so they are less strained during collapse, which is currently being researched. This will result in a thickness that prevents significant diffusion but enables flexibility.

As an initial estimate, the silicone rubber tank's walls will be the same thickness as the structural panels at approximately .003175 m. When taking into account the amount of silicone rubber it will take to cover the four structural panels, the total amount of silicone rubber needed for the tank is about .32813 m3.

Once Asclepius initiates the fuel tank collapse, the propellant will be transferred via piping made of the same Arlon modified PVMQ as the tank. One of the pipe’s ends will be attached to the docking module, where the exiting fuel will be routed to the target’s fuel tank, while the other end will be attached to its respective sub tank at approximately mid height. This location was chosen on the tank as this is where the pressure is concentrated during collapse. The piping will be able to deform during compression so that it can move with the tank wall, as it buckles outwards, and the docking module, as it translates.

While the exact dimensions of the tubing are still being developed, initial parameter estimations can be made. Assuming that the tubing will have the same thickness as the rest of the tank, one can place the tubes initial inner radius at ten times that of its thickness, which would be .03175 m. If each tube needs to be long enough to reach the docking module and then be routed through it, approximately.00802 m3 more of silicone rubber will be needed. While this initial tubing diameter seems small relative to the diameter of the tank, the goal of this is to decrease the amount of propellant that will be left in the tubing after the tank completely collapses. These parameters are subject to change in order to find an adequate mass flow rate that will transfer the propellants to the Deep Space Transport in a reasonable amount of time and without rupturing the tank.

Collapsing Actuators

While it is currently unknown what specific collapsing actuator system Asclepius will be using to physically collapse the tank, an estimation of its parameters can be based off industrial mechanisms used today. For example, a telescopic linear actuator called the "RollBeam" is currently in production by the company Serapid and are designed to operate in "harsh or unusual environments" [1.7]. These operate on a "push-pull system" with a chain that is stored in a magazine upon complete retraction [1.8]. Multiple mechanisms of this nature could be installed and orientated in a way so that they and their chain/physical retracting mechanism will not come in contact with the silicone rubber tank at all during collapse. The tallest sub silicone tank is the one that stores MMH at approximately 1.544 m in height. Assuming the silicone rubber wall's will perfectly fold in half upon collapse, a linear actuator will have to have its magazine opening located about .722 m away from the rigid structural panels to not touch any part of the tank's walls upon collapse. As a preliminary design, four of these actuators can be installed at four equidistant points along the

12 bottom structural panel of each sub tank. Their retracting mechanisms would then extend upward toward/contract away from the top structural panel of their respective sub tank to cause the collapse. As a result, a total of twelve telescopic linear actuators and approximately 16.16 m of chain/physical retracting mechanism will be needed for this system.

Docking Module & Robotic Arm

Asclepius' docking module will route the propellants pumped from the silicone rubber tanks to the Deep Space Transport as it is the physical interface connecting the two spacecrafts. Since the International Docking System Standard is being used to facilitate the berthing, a docking component similar to the active NASA Docking System (NDS) described in the Overview of the NASA Docking System and the International Docking System Standard by George Parma [1.9], will be installed onto Asclepius' docking module. This publication also mentions the NDS's petal height being .175 m [2.9], which can be used as an initial height estimate for the docking module. Furthermore, the SpaceX CRS-7 Mission Overview manifest states that the NDS stored in the unpressurized cargo section of SpaceX's Dragon spacecraft weighed 526 kg [1.10], which is suitable for an initial mass estimate of Asclepius' docking module.

In addition, the robotic arm Asclepius will be using is similar to the "European Robotic Arm" (ERA) that is to be attached to the Russian segment of the International Space Station. In the 42nd Aerospace Mechanism Symposium: The European Robotic Arm: High-Performance Mechanism Finally on its way to Space publication by H. J. Cruijssen, the ERA is described as being 11.3 meters in length and 630 kilograms in weight [1.11]. These values are subject to change with further research into Asclepius' requirements but act as good initial estimates for the robotic arm's parameters.

Command Module

Currently, only preliminary estimates can be made about Asclepius' command module, such as the total size and shape. Since this component of the spacecraft is to contain all the mission systems and sensors that control Asclepius, including all hardware for command and data handling, power electronics and solar panels, communication hardware, and GN&C instrumentation, it must have the volume to do so. As an initial estimate, the command module will be cylindrical in shape with a diameter of 5.604 meters and a height of 2.02 meters. These initial parameters, which result in a volume of 49.82 m3 and surface area of 84.89 m2, are chosen so the command module has the minimal diameter needed to make contact with all the collapsing actuators and a height that is at least half that of the silicone rubber tank's total height to allow adequate room for storage of all the hardware. Since the specific details to how the equipment will be mounted to the command module and what materials will be used for construction are unknown, a mass estimate will be produced once more progress is made.

Weight Estimation

By analyzing the structural parameters outlined in this section, a total weight estimate was derived and compiled in Table 1.2.

Component Mass (kg) Silicone Rubber 773.15 Carbon Fiber Panels 82.18 Collapsing Actuators 307.01 Docking Module w/ Robotic Arm 1156.0 Command Module TBD Total Structure/Mechanism Mass 2318.34 Table 1.2 Structure/Mechanisms Weight Estimate

2. Launch Vehicle

When evaluating which launch vehicle is best suited for the ResuppLION missions, a few factors must be considered or an inefficient, or even unsuccessful, Earth launch is probable. For the scope of the ResuppLION missions, the most important factors are the launch vehicle’s payload weight and size limit. Other major factors include the launch vehicle’s insertion orbit capability, launch date and site, and cost per launch. The remaining factors, such as launch vehicle reliability, payload loading, and vehicle scheduling and availability, will be analyzed in the future.

First and foremost, the launch vehicle must have the capability to carry Asclepius’ payload into a low earth orbit or, preferably, beyond. The weights of Asclepius' different components were estimated, when possible, and can be found in Table 2.1.

Component Mass (kg) Structures 2,318.34 Power 68.0 Engines TBD Propellant 10,000 C&DH 5.0 GNC 30.0 Communications 30.0 Thermal Control 71.62 Primary payload 64,917.0 S/C Gross 77,439.96 S/C Dry 67,439.96 Table 2.1 Asclepius Weight Estimate

Because the gross spacecraft weight is estimated at 77,439.96 kg without the engines, a super heavy-lift launch vehicle must be used during these missions.

In addition, the launch vehicle’s cargo fairing must be large enough to accommodate Asclepius. Since Asclepius is estimated to be roughly 5.61 m in width and 6.24 m in height, without its propulsion unit, it will require a cargo fairing that has a much larger payload volume capability than most in use today.

Due to the rendezvous between Asclepius and the Deep Space Transport occurring in cis lunar space via a Lunar Distant Retrograde Orbit, Asclepius would require a certain amount of ΔV for maneuvering from a low earth orbit into a translunar injection trajectory. Because Asclepius’ primary payload weight is so large, any reduction in weight to other components of the spacecraft will be greatly beneficial to the mission. So, a launch vehicle that is capable of carrying its payload into an injection orbit that minimizes Asclepius’ required ΔV is advantageous.

Aside from these factors, the launch cost is also worth analyzing. This is simply because it often uses a significant portion of the total mission budget. Also, selecting an optimal launch date and

15 site for the planned insertion orbit can decrease injection error, and consequently, decrease the required reserve ΔV. Therefore, optimizing the launch according to the mission requirements will save money that can then be distributed to other parts of the mission.

As a result, the only launch vehicle that is confirmed in development, and can satisfy these requirements is the Space Launch System (SLS). Specifically, Asclepius will be using the SLS Block 2, due to its expected advancements relative to the first planned SLS launch vehicles, the Block 1 and Block 1B. The Block 2 is expected to be available “no earlier than 2028” [2.1], making it available hardware for the ResuppLION missions. Another launch vehicle that is being considered, in case the Block 2 will not be available, is SpaceX’s “BFR.” This launch vehicle was announced in July 2017, which means it is too early in development to know for sure whether it will meet the specifications. For the sake of this analysis, the SLS Block 2 will be the main focus since it is the most promising vehicle at the moment.

The SLS Block 2 is estimated to have a LEO capability of 130,000 kg, which is more than enough to support Asclepius. The Block 2 is able to carry such a high mass to orbit because it will be using an “Evolved Booster” [2.1] to replace the solid rocket boosters used by Block 1B. In contrast, the four RS-25 main engines from Block 1B will also be featured on the Block 2. These engines will produce a thrust of 7,428,530 N while running on a combination of liquid oxygen and liquid hydrogen [2.1].

In addition, the Block 2 will feature a customizable “Integrated Spacecraft/Payload Element” [2.1] so the payload fairing and adaptors can be tailored to the spacecraft. To accommodate Asclepius, a Primary Payload (PPL) interface will be used. PPL is a payload fairing group that accommodates un-crewed spacecrafts, and has the largest fairings available. Specifically, ResuppLION will be using the largest SLS cargo fairing planned, the “10m PLF” [2.1]. This fairing is 10 meters in diameter and 27.43 meters in height, which will produce an available payload volume of approximately 1320 m3 [2.1]. This volume is enough to accommodate the collapsible fuel tank system, with plenty of room for the spacecraft’s other components and any needed payload adaptors. the SLS’s available payload adaptors are engineered to efficiently attach the payload to the SLS and its payload support systems [2.1]. These payload adaptors will be expanded upon in the Space Payloads section.

Another advantage of the SLS is its planned Exploration Upper Stage (EUS), which will be first available on the Block 1B. The EUS is engineered to “provide both ascent/circularization and in- space transportation for payloads” [2.1]. This stage features “four RL10-C3 [2.1] LOX/LH2 engines” for power and provides a standard interface for “various Payload Adapter/PLFs” [3.1]. During the ResuppLION missions, the EUS is extremely useful because it can be used to perform a Trans-Lunar Injection (TLI) to a lunar vicinity [3.1], which is where the rendezvous point is. This TLI will put Asclepius into a Near Rectilinear Halo Orbit (NRHO), where it can make its own trajectory correction maneuvers for the rest of the mission [2.1]. Asclepius’ ascent profile geometries will be tailored so that the altitude and velocity of its lunar trajectory will reduce the ΔV’s needed for future correction maneuvers.

Currently, the SLS is estimated to cost $500,000,000 per launch. While this is relatively expensive, launch costs can be reduced by allowing other missions to occupy the unused Secondary Payload

16 attachments that will be available on Asclepius’ payload adaptor [2.1]. Of course, this is only worth the effort if there are no new constraints that result from sharing the payload space. Furthermore, the ResuppLION missions will currently be launched from the SLS’s default launch location, the Kennedy Space Center (KSC). Since the University of Texas has chosen 04/04/2033 as their launch date, and the SLS is limited to three launches per year, with the constraint of no more than five launches in two years [3.1], the first ResuppLION mission currently has a working launch date of approximately 08/04/2033. This date is subject to change as Asclepius' earth to moon transfer plan is still being developed, which affects the total ΔV and time of flight of the mission. This is explained further in the Propulsion section of the report.

In regards to most of the secondary launch vehicle factors mentioned, more development on the SLS and research must be conducted before it can accurately be gauged if they will satisfy requirements. These factors included the nature of loading the payload, vehicle scheduling availability, and vehicle reliability.

In Figure 2.1, one can find an architectural reference for the SLS Block 2. This figure is included to help visualize the Block 2’s equipment and was retrieved from the “Space Launch System Mission Planner’s Guide” NASA publication [3.1].

Figure 2.1 SLS Block 2 Architecture [2.1]

3. Propulsion Due to the rendezvous being done in Lunar Distant Retrograde Orbit (LDRO), Asclepius will need to travel about 384,000 km at perigee. This has complicated the previous concept of design. Asclepius was initially designed to rendezvous with the hybrid spacecraft in Low Earth Orbit (LEO). With the new rendezvous location orbit transfers have become more complicated. Initially a Hohmann transfer was thought to be a good way to estimate the amount of V that would be required for the mission. After initial calculations, the V that was calculated was too high. A Hohmann transfer would not be the correct way to go. Other methods need to be used to calculate the V required to transfer from the earth to the moon. Due to Asclepius not being reusable the V for a transfer from the moon to earth may not be required. But, it is still a possibility, so it will be included in this report and looked into moving forward as a possibility.

RASC-AL has an initial transportation plan for the hybrid transportation system. They have two methods outlined in their End to End Trajectory for Conjunction Class Mars Missions Using Hybrid Solar-Electric/Chemical Transportation System [3.1]. The first transfer into LDRO uses the Space Launch System (SLS). The SLS will launch the payload into a trans-Lunar injection orbit. Figure 3.1, is an example of what the injection will look like. If this orbit injection is used Asclepius will need to perform a mid-course maneuver to adjust the trajectory and an insertion burn to be placed into the LDRO orbit. Figure 3.2 provides an estimated V and time of flight depending on what day of the month is chosen for launch.

Figure 3.1: LDRO Injection Orbit 1

Figure 3.2: V and Time of Flight LDRO Injection Orbit 1

Another alternative is also provided by RASC-AL for LDRO injection [3.1]. In this second scenario, the payload will be launched by the SLS into a trans-lunar injection or be put into Lunar Distant High Earth Orbit (LDHEO). If the payload is put into a trans-lunar injection orbit then the spacecraft will aim for a low lunar approach instead of aiming for the Earth-Moon L2 point that is seen in Figure 3.1. If the spacecraft is already in LDHEO then only a phasing maneuver will need to be done to target the Lunar Gravity Assist (LGA). Figure 3.3 is a representation of the what the alternative orbit injection may look like. Figure 3.4 has estimates of the amount of V that may be required for the orbital insertion into LDRO. The V shown in Figure 3.2 and Figure 3.4 is only for a transfer from the earth to the moon. It does not include a transfer from the moon back to earth.

Figure 3.3: LDRO Injection Orbit 2

Figure 3.4: V and Time of Flight LDRO Injection Orbit 2

There is also a third approach for an earth to moon transfer, that would be using weak stability boundary (WSB) theory. In Calculation of Weak Stability Boundary Ballistic Lunar Transfer Trajectories, it states that “Ballistic lunar capture transfers from the earth to the moon have the unique property that upon arrival at the moon, a spacecraft is automatically captured in an elliptical

20 orbit without the use of rockets” [3.2]. The theory was first developed in 1987 and since then new WSB transfers have been completed successfully. This is another viable option that will be looked into along with the two possible options RASCA-AL has outlined. These orbit injections along with the WSB transfers were calculated using the Satellite Tool Kit (STK). STK will be used in the future to map out the complicated lunar transfer to achieve the smallest amount of V. All three options only have an outline for an earth to moon transfer. The V required from the moon to the earth will require a different approach to the problem. A Hohmann transfer was attempted to calculate the amount of V, but it was too high. Due to the complexity of the orbit transfers STK will also be used for the moon to earth transfer. In the meantime, the V from the RASC-AL options will be used to estimate the amount of thrust needed for Asclepius. The V estimate for the earth to moon transfer will also be used for the moon to earth transfer; this will not be an accurate number, but it will give a good estimate. 77,439.96 Figures 3.2 and 3.4 can give rough estimates of the V. The V from Figure 2, is about .69 km/s. The highest V from Figure 3.4 is about .34 km/s without the orbital injections and .33 km/s with the orbital injection. These values can then be used to calculate the amount of thrust needed for the V during the earth to moon transfer. The mass of this spacecraft without the fuel it’ll use is expected to be about 67 tons, in An Integrated Hybrid Transportation Architecture for Human Mars Expeditions provided by RASC-AL it states that they expect that the refueling tanker will only need up to 10 tons of NTO and MMH [3.3]. This means that the total mass of the Asclepius is expected to be around 77 tons at the start of the mission. This will drop down to below 10 tons by the end of the mission once fuel is transferred to the hybrid transportation system This is important to keep in mind because that would mean that the amount of thrust needed for the moon to earth transfer will dramatically decrease.

Table 3.1: NTO/MMH Engine Comparison

Super Draco OMS Aestus2 Thrust (N) 71,000 26,700 29,400 Engine Mass (kg) Development 118 111 Length (m) Development 1.96 2.20 Specific Impulse (s) 235 316 324

Table 3.1 has a list of NTO/MMH engines are available for use on Asclepius. The Orbital Maneuvering System (OMS) and Aestus2 are older engines but they have been used for missions in the past and are very reliable. The Super Draco is a newer engine that is currently being developed by SpaceX. The Super Draco’s thrust and specific Impulse are far better than the other two engines by a large margin so it would be the best choice for this mission. Unfortunately, due to the developmental phase it’s in, not much information on the engine can be provided. The Super Draco also has a burn time of about 25 seconds [3.4]. Using the mass of Asclepius, the total V and

21 the burn time a rough estimate of the thrust needed can be calculated.

푭 = 풎 ∗ 풂 (3.1)

Equation 3.1, can be used to calculate an estimate of the thrust. The first option for an earth to lunar transfer seen in Figure 3.1 will require a V of .69 km/s. The Super Draco engines can fire up to 25 seconds so dividing .69 km/s by 25 seconds will give you an acceleration of .0276 km/s2 or 27.7 m/s2. Multiplying the acceleration to the mass of the spacecraft (77 tons or 77,439.96 kg’s) you’ll require a force of 2,137,342.9 N. Dividing this number by the 71,000 N of thrust delivered by the Super Draco thruster means the spacecraft will require at least 30 Super Draco thrusters.

Table 3.2: Max V required with 25 second burn time and a weight of 77 tons

V (Max Required) Thrust Needed (25s # of Super Draco burn) Engines LDRO Injection .69 km/s 2,137,342.9 N 30 orbit 1 LDRO Injection Orbit 2 (Direct .33 km/s 1,022,207.5 N 14 Insertion LDRO) LDRO Injection Orbit 2 (LDHEO to .34 km/s 1,053,183.5 N 15 LDRO)

Table 3.3: Minimum V required with 25 second burn time and a weight of 77 tons

V (Min Required) Thrust Needed (25s # of Super Draco burn) Engines LDRO Injection .58 km/s 1,796,607.1 N 25 orbit 1 LDRO Injection Orbit 2 (Direct .25 km/s 774,399.6 N 11 Insertion LDRO) LDRO Injection Orbit 2 (LDHEO to .27 km/s 836,351.7 N 12 LDRO)

Table 3.2 shows how many Super Draco engines will be needed for the maximum amount of V

22 that will be required during the earth to moon transfer. Table 3.3 represents the same information but with the minimum amount of V seen in Figures 3.2 and 3.4. The total V can be estimated to be what the V from the earth to moon transfer is multiplied by two; it’s a rough estimate. An accurate total V will be presented at a later time once STK is used. As mentioned earlier, the weight of the spacecraft will decrease dramatically so a different amount of thrust will be needed for the V from the moon to earth transfer. That thrust isn’t a big concern because it will be something that the engines on Asclepius will be able to manage. Attitude thrusters will also be attached to the body of the spacecraft to keep it stable, the location and amount of thrust needed will be decided after the number of engines is decided on.

The V estimates and thrust produced from those estimates have helped in further understanding the type of engines that Asclepius will need to perform its mission in the most efficient manner possible. Due to the projected amount of Super Draco engines needed it may be of benefit to look into engines that produce more thrust, specifically cryogenic engines that use liquid oxygen (LO2). This will reduce the number of engines needed on Asclepius. After a more careful V is calculated for the transfers (earth-moon and a possible moon-earth transfer), the amount of Super Draco Thrusters will be revisited, hopefully it will reduce the number. On the other hand, 11 thrusters may be a reasonable number. The dimensions of the Super Draco thrusters will hopefully be available and accurate dimensions of Asclepius will be provided in the future to know how many Super Draco thrusters can be fitted onto the spacecraft.

4. Ground Control The Ground control subsystem dictates the major aspects of the mission on the Earth. While a mission of this type has never been attempted before, with the exception of launch and rendezvous there is no need for constant allocation for the full resources of a dedicated ground station, and thus the ResuppLION mission will not be creating a dedicated ground station system from scratch and instead use and share existing ground control architecture. The Mission Control Center for the ResuppLION mission will be located at the Jet Propulsion laboratory(JPL) in California. Given the mission’s use of the Deep Space Network for all communications, which is operated by JPL from their Deep Space Operations center, it is an obvious choice to use JPL’s facility. Furthermore, JPL has a long history of acting as the mission control center for a variety of uncrewed missions and will be more comfortable working with the planned autonomous rendezvous than other facilities[4.1] JPL’s Space Flight Operations Facility will act as the mission’s Spacecraft Operation Control Center(SOCC). The facility will oversee the monitoring and analysis the telemetry and other mission data from the spacecraft bus and common systems which dictates the attitude and dynamics of the spacecraft[10.1]. Furthermore, it is the only part of the ground system which will be directly sending commands to the spacecraft and thus is beneficial to be located in the same facility as the Mission Control Center. The Payload Operations Control Center(POCC) would ideally also be located at JPL, so it can easily coordinate with mission control and the spacecraft operation control. However, while JPL’s resources are expansive, it may be in the interest of the mission to have the POCC elsewhere to ensure enough resources are available for the mission. The primary candidate would be Goddard Space Flight Center in Maryland, which handled the Lunar Reconnaissance Orbiter in the past, which much of the communication subsystem of Asclepius is based on. This would require significant changes in the existing ground communication’s architecture as the Goddard Space Flight Center primarily operates the Space Network and Near Earth Network, and has little history working with missions which utilize deep space network[4.2]. However, the space center has more than enough resources to house a POCC. The POCC is responsible for monitoring and analyzing the payload instruments not monitored by the SOCC, and coordinates with the SOCC to send commands to these instruments.

5. Communications The communication sub system is the interface between the Asclepius, ground control, and the target spacecraft. Original plans to use a set of low gain antennas as the primary means to communicate with earth have been discarded due to concerns that a low gain antenna would use 퐸 a large amount of power to meet required energy per bit to noise power spectral density ratio ( 푏) 푁0 for the desired Modulation and Coding schemes. A set of low gain antenna will still be placed on the spacecraft as backup should the main high gain antenna fail. The number and type of low gain antenna is still being decided upon, however initial candidates include Horn and helix antenna due to higher gains and low mass. The preliminary uplink budget was decided to be 64bits sent once every second and a downlink budget of 1000kb sent every 10s, resulting in a 64bps and 100kbps budget respectively. The total data requirement is still subject to debate as each subsystem has not compiled a data requirement and not all sensors have been identified at the point of this report. The current 100kbps downlink budget was estimated to be a maximum data rate given that the downlink will primarily be health and telemetry and the lack of a science payload[10.1]. After link budget calculations (see Appendix 1. for full link budget calculations) it was found that S band would be sufficient for the required data rates and would require a relatively small parabolic dish. However, with the current configuration Ka band can easily be utilized with minimum increases in dish size if higher data rates are desired. QPSK signal modulation will be used on Asclepius as it allows for an increased spectrum utilization, and prior use on the Lunar Reconnaissance orbiter which much of Asclepius’ communication’s subsystem is based on5.1.

퐸 In order to accommodate QPSK’s energy per bit to noise power spectral density ratio ( 푏) and an 푁0 extra 3db to account for implementation losses or errors, and with a set power of 20W, Asclepius will require a 0.6m wide parabolic dish. It is currently being determined how the dish will be mounted to Asclepius to hold continuous coverage with earth without affecting rendezvous and being shadowed by the spacecraft. Preliminary considerations are to mount the parabolic reflector on long arm on a gimbal much like the Lunar Reconnaissance Orbiter. Asclepius will be utilizing the Deep Space Network to communicate with ground control in order to accommodate the tremendous distance from the earth to the moon, while a small low gain antenna will be used to communicate during rendezvous with the target spacecraft. However, given the fact that the target spacecraft will similarly be using the Deep Space Network for communications, if a working protocol cannot be determined during rendezvous for communications to the Deep Space Network between the two spacecraft it may be in the interest of both parties for the construction of another set of parabolic reflectors on Earth. The current proposed protocol is for the target spacecraft to act as a relay for Asclepius during rendezvous given the more powerful antenna on the target spacecraft.

6. Command & Data Handling The Command and Data Handling (C&DH) subsystem communicates with every subsystem in the spacecraft to appropriately handle and deliver commands, considered the 'brain' of the spacecraft. This system is built to maximize the number of decisions that C&DH makes autonomously while ensuring that enough information is available to operators on the ground to make informed decisions. The C&DH system includes an onboard computer which controls the operation of all systems by monitoring and responding to conditions as well as executing written-in functions. Installed software manages the programs written to handle various tasks, firmware for the permanent read-only memory and the hardware components with the capability to fit mission requirements.

C&DH main tasks: 1. Subsystems and payloads control 2. Receive commands from Earth and prepare transmission data for Earth 3. Manages collection of solar power and battery charging 4. Keeps and distributes spacecraft time 5. Spacecraft position calculations 6. Carries out commanded maneuvers 7. Autonomously monitors and responds to a wide range of potential onboard problems 8. Allocated memories for data and software storage 9. Docking interaction 10. Controlling the payload sensors

Throughout Asclepius' mission, there will be multiple phases - each requiring different focuses of the C&DH system. Once launched into LDRO, the GNC subsystem will go to work for orbit maintenance and attitude control with responding thrusters while the thermal control monitors the conditions. During the fuel transfer phase, these systems will continue to operate and the robotic arm and fuel transfer system will execute functions to successfully refuel the Transport vehicle. Following undocking, the robotic arm and fuel transfer system will be disabled and the computer will be dedicated to the remaining subsystems. All telemetry data collected by the C&DH system will be available to Ground Control using the radio antenna on the spacecraft. Through this, Ground Control will have the ability to intervene in the mission when necessary.

Hardware Design As mentioned by Military and Aerospace Electronic, one difficulty when experiencing an atmosphere with high levels of radiation is bit-flips in temporary memory, which results in unpredictable software errors. For this issue, a programmable safety timer will be used to reset the processor if a bit-flip occurs and the processor will reload the boot-software from ROM to RAM. This means the boot-software must be stored in memory (bit-flips cannot occur in memory). [6.1] The operating system software will be stored in the temporary memory and a copy in the permanent memory. Another area of correction will be implemented to find hardware and software design errors. A delay interface will be used to correct the errors by uploading new software to temporary memory.

The on-board computer will be connected to the communications system to communicate with the ground station. To accomplish this, a Universal Asynchronous Receiver Transmitter (UART) will be used. UART is the most used serial communication protocol for integrated circuits. This device has configurable data format and transmission speeds, allowing serial and parallel data conversion and handles serial timing.[6.2] UART will be used as the interface for all serial transmission ports – this looks to be thermal control, fuel transfer and the robotic arm.

Software Design All the programming will be written in the C language. This is perhaps the most popular language for programming embedded systems according to Colin Walls, author for the EE times. It remains a very popular language for microcontroller development due to the code efficiency, reduced overhead and development time. The C language is written and compiled into binary that the processor understands and then downloaded into the microprocessor.[6.6] An embedded compiler must be available for the processor.

An important part of software development is writing several subroutines or protocols to avoid inherent problems for on-board computers. Subsystems rely heavily on each other, potentially causing fatal errors that execute incorrect parts of the program, written for other subsystems. In a worst case scenario, this could cause a failed mission. [6.6] This will also be avoided using the bus line architecture to decrease reliability across the spacecraft and decrease processing speed.

System Architecture A multi-drop data bus system architecture with a distributed processing resource was chosen for the Asclepius mission. As stated in SMAD, this architecture allows direct data transmissions across a common network with interconnect to the various components integrated into the bus structure. Based on efficiency, dependency and reliability, the Centralized or Ring Architectures are inferior to the Bus for the ResuppLION mission[10.1] As explained by I2C Bus, the inter-integrated circuit (I2C) is a multi-master bus providing arbitration and collision detection, multi-slave, packet switched, single-ended, serial computer bus invented by Philips Semiconductor. I2C was designed to allow easy communication between components which reside on the same circuit board. Each device connected to the bus is software- addressable by a unique address. The original communication speed was defined with a maximum of 100 kHz but faster modes have been developed: 1 MHz, 3.4 MHz, and 5MHz. These allow for 10-bit address space and 112 devices on the bus. This allows for shared bus lines that enable multiple masters to transmit data on the same line. [6.7]

Asclepius will use a multi-bus I2C for communication between the components of the spacecraft to decrease the reliability that systems have on each other by having separate bus lines for each component, linked by cable. When multiple masters are driving the bus (shared line), each device needs to be able to recognize that another device is currently talking and the bus is therefore busy. This means they follow arbitration logic and may detect ongoing bus communication, as to not interrupt it. Dedicated lines will be used for time sensitive systems while the others will use shared lines. Shared lines take the commands and line them up in a queue by the order they were received

27 and the commands will be executed in this order. This causes a time delay for the commands that a dedicated line does not have.

On-board Computer At the core of the Asclepius' C&DH system is the onboard computer. The desired features for this were a static read-only memory as well as a random-access memory to cover both written-in functions as well as responsive functions, high processing speed necessary for the mission functionality, C compiler for software, a 32-bit architecture minimum, controllable I/O for ground control purposes, programmable UART interface for serial data transmission, master and slave SPI Serial Interface, programmable timers and be operable in extreme conditions.

The BAE Systems RAD 5545 SpaceVPX single-board computer is the perfect fit for this mission. BAE states that their space computer line has provided more than 900 computer across 300 satellites, accumulating more than 9,000 years of flight heritage.[6.4] The product specifications from the BAE release publication are shown in Figure 6.1 describe the features of 5545 with one of the most distinct ones being a total ionizing dose of 100 Krad, latch-up immunity which is a type of short circuit in an integrated circuit and a low single-event upset where one ionizing particles strikes a sensitive node in the microprocessor.[6.5] This is appropriate protection for the known conditions that the spacecraft will be exposed to during the mission. Another key characteristic seen here are the diverse protocols built-in including UART, SPI and I2C.

This microprocessor will be radiation hardened against ambient radiation, meaning it is capable of working reliably in the outer space conditions. According to Military & Aerospace Electronics, radiation accelerates the aging of electronics, leading to degradation of electrical performance or permanent [6.1] failures. Figure 6.2 RAD 5545 specifications [5]

Note: No pricing was available for this computer on the manufacturer’s website. Per 2002 reference, the RAD750 is a comparable system that gives an estimated price of $200,000. [6.3]

Mission Phases For the entire mission, at least 3 bus lines will be dedicated. The allocation of dedicated versus shared lines will change with the mission phases to accommodate different functions and system responsibilities. The configurations for all mission phases mapped out in Figure 6.2. These bus lines are described in Figure 6.3. The following are the phases that Asclepius will go through to complete the ResuppLION mission:

1. Orbit Maintenance and Attitude Control 2. Close Proximity Entry and Docking 3. Refueling and Undocking 4. Close Proximity Departure 5. Orbit Maintenance and Attitude Control 6. Earth Reentry/Moon Disposal

Mission Phase GNC Thrusters Thermal Robotic Arm Fuel Transfer Comm 1 D D D Locked Off Off D 2 D D S D S S 3 S D S Locked Off D D 4 D D S Locked Off S S 5 D D D Locked Off Off D 6 D D D Locked Off Off D

Figure 6.2. Phase Diagram showing the allocation of the Dedicated (D) and Shared (S) lines throughout all 6 phases of the mission. Dynamic allocation increases efficiency and reliability to focus on the important systems of each phase.

Figure 6.3. Detailed view of the spacecraft data bus lines that will be allocated to either dedicated or shared I2C throughout the mission.

7. Guidance, Navigation & Control Attitude Determination and Control Attitude Control will be accomplished using unidirectional reaction control thrusters. These thrusters will be the sole source of stability; passive stabilizers such as a gravity gradient or spin stabilization will not be used. The additional weight and hardware required for these methods are not worth the stability gained. These thrusters can also function as the position and fine orbit control that can be used to maneuver Asclepius close to the target spacecraft. The primary concern with thrusters as attitude is the coupling of attitude maneuvers and positioning maneuvers. To control n degrees of freedom, a minimum of n+1 unidirectional thrusters are needed. Because Asclepius is operating in 6 degrees of freedom, it will need at least 7 thrusters[7.1].

Thrusters have a possibility of failing and so more than 7 thrusters will be necessary on Asclepius. Choosing a thruster configuration (set of positions and orientations) is the current topic of study. It is anticipated that thrusters will fail in two possible modes: stuck on and stuck off. The latter is trivial to deal with if there are redundant thrusters on board. A “stuck on” thruster would be significantly more difficult to deal with because this will exert a large constant torque on Asclepius as well as consume large quantities of fuel. The Asclepius will need to have a highly reliable kill switch for each thruster. Asclepius will also need to be able to autonomously maintain safe angular rates during the failure detection and kill process.

There are five major phases of attitude control in the mission detailed in table 7.1.Each phase will use a set of sensors, actuators, and control algorithm specific to its stage. For sensors, Asclepius will contain a single Ring Laser Gyro for attitude rate sensing. The gyro is only critical for spacecraft control during the deployment stage. It can be used during the other phases. To increase control accuracy but it is ultimately not necessary because nominal spacecraft angular rates will be near zero. Because of this, a redundant gyro will not be used. There will be two star trackers for redundancy. Star trackers are optical devices that can be damaged by micrometeorites and cannot be pointed directly at the sun. Because of this, two star trackers are standard on many spacecraft[7.2,7.3]. Lidar and retroreflectors will be used for state estimation when Asclepius gets into close proximity with the target spacecraft.

It is expected that Asclepius will be deployed from the launch vehicle with some initial angular rates on all three spacecraft axes. The spacecraft needs to null these rates before reliable communications and power generation can begin. A ring laser gyro which can sense angular rates will be used as the feedback sensor to initially null these rates.

In freeflight, Asclepius will have two pointing attitudes: sun pointing and communications pointing. Sun pointing will maximize the power to the solar panels and communications pointing will maximize the communication ability of the antenna. The spacecraft will slew to these attitudes and maintain them using the thrusters.

Table 7.1 Mission Phases

The attitude mode with the largest constant torque on Asclepius will be maneuvering pointing. All orbital maneuvers are impulsive and therefore the engines are high thrust. The spacecraft design will have the thrust vector pointing through the center of mass; however, small angular misalignments of the engine combined with error in center of mass placement will result in a constant torque on the spacecraft during orbital maneuvers.

As Asclepius approaches the target spacecraft, it is more important to maintain the attitude with respect to the target than Asclepius’ attitude with respect to an inertial or earth frame. Three options are laid out in figure 7.1. The first option would be to use the current attitude sensors and build up an attitude with respect to the target from communication of their state. Alternatively, it could be assumed that they are in the expected attitude and approach their spacecraft under this assumption. The last option is to measure the relative 14 attitude directly using Lidar and retroreflectors. The first two options were too computationally intensive, subject to error, had greater time delay, and had more possible modes of failure. Because of this, Asclepius will use Lidar and retroreflectors to directly measure the attitude with respect to the target. This data can be cross checked against inertial attitude measurements to verify it’s validity. Optical targets were considered instead of the Lidar system. These were ruled out for a few reasons. Autonomous image processing is more computationally intensive than the Lidar system, optical systems would have a shorter effective range for position and velocity sensing, and there are fewer companies developing optical systems for rendezvous[7.4]. Using a Lidar system will allow team ResuppLION to buy a system built by another company rather than designing one.

Figure 7.1 ProxOps Attitude Determination Options

Orbit Determination and Control For spacecraft navigation, Asclepius will use the Doppler Orbitography and Radio Positioning Integrated by Satellite (DORIS) system[7.5]. This is a system of Radio emitters on earth. Asclepius will have a receiver that can measure the radio waves and their doppler shift to determine the position and velocity to determine an orbit. This has an advantage over ground based systems of being integrated into the satellite. Each ground antenna has a limited cone that it can transmit on. The coverage will be suitable at the altitudes that Asclepius will be operating at. The coverage at 1335 kilometer altitude is given in figure 7.2. Once Asclepius is in close proximity to the target spacecraft, it will use the Lidar system for position and velocity relative to the target spacecraft.

Figure 7.2 DORIS Coverage

For major orbital maneuvers, Asclepius will use a large primary engine (more details in the propulsion section of this report). This will get Asclepius into an orbit close to the target spacecraft. At this point, Asclepius can use the attitude thrusters for position control. They are much lower thrust compared to the main engine and so they’re well suited to the fine position adjustments required in berthing with the target. When Asclepius departs, it will use the main engine to bring it back down into an orbit that passes through earth’s upper atmosphere. Aerobraking will then be used to bring the apoapse down even further before one final main engine burn is used to get Asclepius to reenter the atmosphere.

8. Power Theme three has allowed this team to work with another team for this project. Asclepius will be resupplying the hybrid transport spacecraft in LDRO during the length of its mission. In the mission’s requirements, it states that the Hybrid Transport System should have a lifetime of at least 14 years. At the moment, the University of Texas isn’t sure how long they expect their lifetime to be so 15 years will be used as an estimate the longevity of the mission.

Subsystem Estimated Power GNC 270 W Thermal Control TBD Propulsion TBD Communication 30 W Command and Data Handling 35 W Total 330* *Not including Thermal Control and Propulsion Table 8.1: Asclepius Power Estimates Unfortunately, there aren’t any exact values for the amount of power that each subsystem will require yet. A total of 330 W will be needed from current estimates, not including Thermal Control and Propulsion estimates. Once the other two subsystems have an accurate power output, the amount of power needed could be 400 W. The collapse actuators that will transfer the fuel into the other spacecraft will require at least 2000 W of power. The robotic arm which will be used to help with the berthing process will require a similar amount of 2000 W of power.

Asclepius will be using solar panels as the primary power source and secondary batteries as a power source when the solar panels are not available during eclipse periods. The solar panels will supply power directly to the system and that power will be regulated using a voltage regulator and a heat sink. Any excess power will be dispersed as heat which could be used to help Asclepius' thermal environment to keep other systems at their operating temperature.

The solar cells that will be used for this mission will depend on their, efficiency, radiation tolerance, lifespan, structural robustness, reliability and cost [8.1]. There are different types of cells that are available: single crystal silicone cells, single junction based cells, multi-junction based cells, and amorphous/polycrystalline cells. Table 8.2 shows the typical efficiencies of those cells.

Cell Efficiency (%) Single crystal silicone 12.7-14.8 Single junction based (GaAs) 19

Dual junction 22 Multi-junction based (GaAs) ~27 Amorphous/polycrystalline ~10 Table 8.2: Different Type of cell Efficiencies Due to the longevity of the mission and where the mission will take place, a high cell efficiency will be preferred. Dual and multi-junction cells have the highest efficiency, table 8.3 highlights the attributes of both cells. Under STC* Dual Junction Multi Junction Efficiency (%) 22 ~27 Operating Voltage (V) 2-2.5 2-3 Radiation Tolerance 0.80 0.84 Cell Thickness(micro-meters) 140-175 140-175 Absorbance 0.91 0.92 Cell Weight (mg/cm2) 80-100 80-100 Cost (1/cm2) ~$21.00 ~$26.00 *STC= Standard Temperature Conditions Table 8.3: Properties of Dual Junction and Multi Junction Cells Table 8.3 shows that multi-junction cells are a better type of cell due to their efficiency, operating voltage, and radiation tolerance. The cost for multi-junction cells are more expensive but the efficiency that it provides along with the other attributes it holds make it the best choice.

There are various venders that sell different types of multi-junction cells. Two top venders; Azur Space and Spectrolab supply triple-junction space cells. Table 8.4 shows the properties of the two different types of cells for both companies. Azur Space 30% Triple Spectrolab 30.7% Triple Junction Junction Efficiency 29.5 30.7 Volts (V) 2.7 2.175 Current 17.23 18.1 (mA/cm2) Solar .9 .88 Absorbance Emittance .87 .85 Operating 15-60 (optimal at 25) 15-75 (optimal at 28) Temp (C) Power Density 419.264 418.264 (W/m2) Weight 86 84 (mg/cm2) Table 8.4: Azur Space and Spectrolab cell comparison

Using the basic power equation (8.1), 푷풐풘풆풓 = 푽풐풍풕풂품풆 ∗ 푪풖풓풓풆풏풕 (8.1) the power output for a cell can be calculated using the voltage and current given for that cell. According to AzureSpace, their TJ Solar Cell 3G30C can supply an average of 465.38 W/m2 at beginning of life (BOL). The degradation factor for these cells is about 0.90 so a power average of about 419.264 W/m2 can be expected after 15 years. Spectrolab states on their website that their XTJ Prime Cell can supply 491.231 W/m2 at beginning of life with a degradation factor of 0.85, it's expected to produce 418.98 W/m2 after 15 years. These values are rough estimates and were calculated assuming that the cells are operating at their optimal temperature as well as in direct sunlight. These values can and will probably change due to temperature fluctuations and the sun’s angle of incidence on the solar panel. Looking at both cells, Azur Space’s TJ Solar Cell 3G30C will be used for the solar arrays on Asclepius. Spectrolab has a better BOL with a power of 491 W/m2 compared to Azur Space’s 465.39 W/m2, but Azur Space has a better operating voltage, better solar absorbance, and emittance. Azur Space and Spectrolab have a close EOL power rating for 15 years, but Azur Space has a better degradation factor of 0.90 so if the mission needs to be extended more than 15 years then Azur Space will have a much better EOL power rating than Spectrolab. Using the data supplied from Azur Space the size of the arrays can be estimated. As mentioned earlier the estimated required power for this mission will be 400 W at all times and 2000 W during the docking phase of the mission. Each solar cell will be able to produce an average of 1.404 W, while each solar cell is about 30.18 cm2. That info was then used to generate a value of 465.38 W/m2 at BOL. This value cannot be used to determine the solar array size due to degradation of the cells. The solar array size will need to be calculated using the expected EOL value to make sure that Asclepius will have enough power to complete its mission.

푷푹풆풒 (8.2) 푨푹풆풒 = 푷풐

Using equation 8.2, you can calculate the amount of area required for the solar panel. Dividing 2000W by 419.264 W/m2 will give you an area of 4.77 m2. Rounding this number to 5 m2 means that Asclepius will need either two 2.5 m2 solar arrays or one 5 m2 solar array. The solar array size is an estimate that doesn’t take into account the direction in which the solar arrays will be pointing nor the amount of sunlight that the cells will be able to absorb due to the angle at which the sunlight will hit the cells, modifications will need to be done.

There are different types of solar array panels that can be used for the solar cells. The solar array panels that have been used in previous missions are, body-mounted arrays, rigid panel arrays, flexible panel arrays, flexible roll out arrays, concentrating arrays, and high temperature arrays.

36 according to Solar Cell and Array Technology for Future Space Missions by Roa Surampudi and others [8.1]. Body-mounted arrays are used for smaller satellites that may require a few hundred watts. Rigid Panel arrays are used for missions that require several hundred watts to tens of kilowatts of power. Flexible arrays have about the same properties as rigid panels arrays but they also have high packing efficiency, and a simple deployment system. Flexible roll out arrays can provide tens of kilowatts of power, it is packed into a cylinder during launch which allows it to also have high packing efficiency. Concentrating arrays are primarily used for outer planetary missions that require a high specific power. High Temperature arrays are used for missions that have close encounters with the sun. This mission only requires the spacecraft to travel from the earth to the moon and its estimated it will need at least 2000 W of power and the size of the arrays will need to at least be 5 m2. Rigid panel arrays and flexible panel arrays would meet those requirements the best. A rigid panel can range from $500 to $1500 per watt and a flexible panel can range from $1000 to $2000 per watt. Currently a decision on which panel to use cannot be made due to the constraint of not knowing the specific dimensions of the spacecraft and where all the components will be placed which is a huge influence on what type of panel will be permitted. A secondary battery will also need to be used on Asclepius. A secondary battery can be used during “launch and post launch until the deployment of solar panels, for firing pyros and firing rockets for attitude control, during cruise anomalies or trajectory control maneuvers of the spacecraft and during eclipse periods” [8.2]. A good secondary battery will need to be reliable, robust, as well as the ability to operate under extreme temperatures, and have a tolerance to radiation. In addition to those, a lightweight and compact battery is highly preferred. In the previous report, there were four different battery types that were being evaluated. Those were, Nickel-Hydrogen, Nickel Cadmium, Lithium-Ion, and Silver Zinc. Nickel-Hydrogen and Nickel-Cadmium have been the main type of batteries that have been used in previous space missions, and have proven to be very reliable. Battery Specific Energy Operating Life-Span Cycle-Life Energy Density Temp (C) (year) (partial (Wh/kg) (Wh/L) depth discharge) Ni- 35 100 -10 to +20 <5 >30,000 Cadmium Ni- 40 80 -10 to +30 5 to 10 >40,000 Hydrogen Ag-Zinc 100 200 -10 to +25 <1 <100 Li-Ion 100 240 -30 to +40 4 1000 Table 8.5: Battery Properties Comparison Table 8.5 compares the four different battery types, Lithium-Ion stands out with its operating temperature range, its energy density and specific energy. NASA believed that “lithium-ion batteries offer significant advantages in terms of mass, volume, and temperature range and were perceived as the appropriate next generation technology” [8.2]. The life-span of Lithium-Ion

37 batteries are sub-par compared to Nickel-Cadmium or Nickel-Hydrogen but it still offers better properties in other areas than the other batteries. Lithium-Ion is a new type of battery that is still being developed today. It has been tested on the Mars Lander under the 2001 Mars Surveyor Program. It was a successful test that helped solidify Lithium-Ion as the best in class secondary battery during that time. Due to the initial advantages seen in Lithium-Ion, the heritage that it carries and the continued performance upgrade of the battery, it would benefit the mission to use Lithium-Ion as the secondary power source on Asclepius. There are various different types of Lithium-Ion batteries that can be purchased. ABSL and EaglePicher are two highly reliable vendors who have them available. Table 8.6 has two batteries from both companies that can be used on Asclepius. The price of these batteries were not given. These batteries must be able to produce at least 400 W of power and it would be preferred that they also produce 2000 W of power. Eagle SAR 10197 and ABSL8s104p produce well over 2000 W of power. While Eagle SAR 10199 and ABSL 8s52p produce just above 2000 W. Due to the requirements of this mission. Eagle SAR 10197 and ABSL 8s104p will not be used. They both produce more than 4000 W of power which is too much for the nature of this mission. Eagle SAR 10199 and ABSL 8s52p both produce around 2000 W of power. They have similar qualities but because weight is a big factor on Asclepius then ABSL 8s52p is the best choice. Two batteries will be used for this mission in case one malfunctions, or fails. It’ll also be available in case there’s an issue with the solar arrays and the spacecraft would have to stay under the secondary power source for an extended amount of time. Eagle SAR- Eagle SAR- ABSL 8s104p ABSL 8s52p 10197 10199 28V 28V Nominal Voltage 33.3 33.3 28 28 (V) Specific Energy 104.9 94.6 90.02 98.5 (Wh/kg) Shock Level (G) 1135 1135 1907 2000 Weight (kg) 63.5 35.2 49.9 22.8 Operating Temp -5 to +35 -5 to +35 5 to 35 0 to 45 (C) Power (W) 5128 2564 4368 2184 Table 8.6: ABSL and EaglePicher Batteries This mission is not expected to be ready for deployment until 2029, which is about 11 years into the future. Currently the equipment selected; two batteries that weigh a total of 45.6 kg, a solar panel that’s estimated to weigh 22 kg will sum up to a total weight of about 68 kg. Technology will improve, solar array cells will become more efficient and Lithium-Ion batteries will also become more efficient as well. This could mean that the weight of the batteries can be reduced in the future. Moving forward, an investigation into technology that is being researched and developed now will be looked into and evaluated for its use in 2029. A deeper investigation will also be conducted into the type of solar panel that will be used, be it rigid or flexible solar panel. When a decision on the type of solar panel is made, then a more extensive analysis will need to be conducted. The analysis will focus on the orientation of the solar panel and the expected amount

38 of sun that the panels will receive during the mission.

9. Thermal Control Asclepius is an uncrewed spacecraft and therefor the thermal control has been designed around maintaining operational temperatures for internal components and fuel. By gathering the baseline temperature requirements found in table 9.1 from Space Mission Design and Analysis (SMAD)[10.1], it was determined that within the mission systems portion of the spacecraft the temperature must range between 10 and 30 degrees Celsius to satisfy all component needs. Based solely on the temperature requirements of the fuel, the tank must maintain a temperature between 15 and 40 degrees Celsius. Because the tank design requires the material to be flexible, the material selection will also have an impact on the thermal control of the tank. This tank material has not been finalized yet, so thermal analysis of that will be incorporated at a later time. It is possible that the tank will require an active heating system.

Operational Component Min Max Fuel 15 40 Antennas -100 100 Solar Arrays -150 110 Batteries 10 30 GNC -40 70

Avionics Baseplates -20 60

Table 9.1 Component Operational Temperature Requirements (Celsius)

To regulate the Asclepius’ temperature, active and passive control methods will be used. The command module portion of the spacecraft will be covered with multi-layer insulation (MLI) to keep the internal components at an operational temperature during eclipse and full sun exposure periods of orbit. Because of the heat generated from the internal components, the spacecraft will also incorporate one or more louvers, allowing it to radiate heat out when the sun exposure is extreme. The exact number and location of the louvers has yet to be determined. These louvers will be the main part of the active thermal control system and will be automated using data fed from Thermistors. Thermistors were chosen over Resistance Temperature Detectors (RTD) because they have a quicker response time and are less expensive, according to Space Health Monitoring Using Analog Multiplexers and Temperature Sensors [9.2].

An initial spacecraft thermal analysis was performed using the equations 9.1, 9.2, and 9.3, ignoring internal heat radiation for the time being. Equation DD was derived to calculate the temperature of the spacecraft at the 3 different orientations shown in Figure 9.1. Assuming an absorptivity of one, emissivity of one, and preliminary spacecraft dimensions the temperature of Asclepius (Table 9.2) was able to be estimated for the three different attitude orientations shown in Figure 9.1. As seen in Table 9.2, the majority of the temperatures are below operational standards, so with further analysis and research including internal heat radiation, more accurate absorptivity, and emissivity constants; a more detailed thermal control system will be able to be designed.

푄푖푛 = 푄표푢푡 (9.1)

푄푖푛 = 훼휑푠퐴푒푥푝표푠푒푑 + 푄푖푛푡푒푟푛푎푙 (9.2) 4 푄표푢푡 = 𝜎휀푇 퐴푡표푡푎푙 (9.3) 훼휑 퐴 푠 푒푥푝표푠푒푑 1/4 푇 = ( 4 ) (9.4) 𝜎휀푇 퐴푡표푡푎푙

Figure 9.1 General Spacecraft Orientation Towards the Sun

Mission System Temp. w/o internal heat Tank Temp. w/o internal radiation heat radiation -273 -22 -23 -273 -8 -168 Table 9.1 Spacecraft Temperature (Celsius)

10. Payload Due to resupplying spacecrafts rarely being used, one can consider the objective of the ResuppLION missions uncommon. Asclepius is not performing a scientific mission, so it uses a small number of specialized instruments to perform its task.

Because Asclepius’ primary objective is to deliver propellants to the Deep Space Transport, its primary payload consists of the resupplying propellants. The instruments of the collapsible fuel tank system, including the silicone rubber tanks, the structural panels, the collapsing mechanism architecture, and the docking module with its robotic arm, are what deliver the primary payload. These instruments are active payloads, as they are directly communicating with the target subject (Deep Space Transport) [9.1]. One can see that the primary payload and the instruments that deliver it make up a large part of Asclepius' architecture. The details of these instruments can be found in the Structures and Mechanisms section of the report.

Also, Asclepius will have multiple communication payloads for telemetry, tracking, and control. The communication payloads consist of a 0.6 m wide parabolic dish as a primary means of communication, with a set of low gain antennas in case of the high gain antenna failing. These instruments were chosen in order to facilitate communication with ground control at the distance Asclepius will be in its Lunar Distant Retrograde Orbit. The nature of the information transferred consists of the relevant data produced and received by each subsystem [9.2], and is expanded upon in the Communications section of the report. While the specific types of antennas Asclepius will use are currently being researched, manufacturing specifications are still being developed.

As stated before, the 10m PLF will be the fairing used to accommodate Asclepius and its customized payload adapter that will securely attach it to the EUS. As a preliminary design, the payload adapter will consist of a Payload Attach Fitting (PAF), to act as the structural interface to the SLS’s EUS, and a Payload Separation System (PSS), to act as the “structural separation interface for a payload mounted on the PAF” [9.3]. The PAF is “constructed of composite sectors with horizontal and vertical joints” [9.3]. In order to meet specific payload requirements but also maintain an efficient mass, the PAF’s composite sectors can be lengthened or shortened. As of right now, the maximum load a PAF/PSS adaptor can hold is 40.0 metric tons. While this maximum load is below Asclepius’ total gross weight, these adaptors were designed to adapt over time to the variety of payloads the SLS will inevitably need to carry [9.3]. Therefore, a custom payload adaptor can be engineered for Asclepius if enough time is allotted for the request.

In addition, SLS payload adaptors can use the EUS to provide resources and interfaces to the PPL fairing during ground and flight operations. This includes the electrical interfaces that provide power or the flight data services for PLF jettison and spacecraft separation commands. If needed, the payload can be electrically power post lift-off by mounting payload batteries onto the PAF [9.3].

Summary Tables Weight, Cost, and Power Estimates Table 13.1 provides the current cost summary for the first ResuppLION mission when applicable. Once more is known about the development and production cost of each spacecraft component, an 85% learning curve during development of new technologies will be used to approximate an accurate total cost of the first mission. This number can then be used to approximate how much the total cost of all the ResuppLION missions will be, as each mission is expected to be very similar in nature.

Type Cost (Millions of Dollars) Development TBD Production TBD Launch 500 Total TBD Table 13.1 First Mission Cost Estimate

A summary of mass estimates, when applicable, for Asclepius were approximated by analyzing each subsystem. This summary is compiled in Table 13.2.

Component Mass (kg) Structures 2,318.34 Power 68.0 Engines TBD Propellant 10,000.0 C&DH 5.0 GNC 30.0 Communications 30.0 Thermal Control 71.62 Primary payload 64,917.0 S/C Gross 77,439.96 S/C Dry 67,439.96 Table 13.2 Summary Weight Estimate A summary of power estimates, when applicable, for Asclepius were approximated by analyzing each subsystem. This summary is compiled in Table 13.2. Subsystem Estimated Power GNC 270 W Thermal Control TBD Propulsion TBD Communication 30 W Command and Data Handling 35 W Total 330*(Not including Thermal Control and Propulsion ) Table 8.1: Asclepius Power Estimates

Recommendations for Future Work Structures and Mechanisms: In the future, the Structures and Mechanisms subsystem will finalize specific parameters relating to all six of Asclepius’ modules. This includes the thickness of the silicone rubber tank’s walls, the temperature at which each sub tank will be held at, the retracting mechanism the collapsible linear actuator will be using, and the specifications of the tank’s tubing. In addition, the details to how the command module will house all of the equipment and hardware required to operate Asclepius will be developed. From here, a more accurate mass estimate can be made of the subsystem. Furthermore, a detailed computer-aided design model of Asclepius will be made with a software that has yet to be determined. Structures and mechanisms is also currently working on the actuator mechanism that will be performing the tank collapse. A trade study is being performed on different designs. One interesting design that’s being analyzed is similar to a wind up tape measure. A half cylindrical tube would be rolled around a motor has the fuel is transferred. In it’s rolled out state, it will be rigid and support the tank. The tube will be flattened as it is rolled up into its compact state. Launch Vehicles: Looking forward, the Launch Vehicles subsystem will determine factors including the nature of loading the payload, vehicle scheduling availability, and vehicle reliability. These factors will then help finalize parameters such as the optimal launch date for each ResuppLION mission and the target translunar trajectory injection accuracy. Additionally, launch site preparations will be attended to, including an analysis of the launch site operation budget.

Propulsion: This report only contained estimates of the amount of DV required for the earth to moon transfer. Moving forward STK will be used to calculate a more exact DV. The type of transfers that were introduced are a good base that will aid in achieving the best DV. Once an exact DV is calculated, then an exact thrust can be generated. The thrust will indicate the amount of Super Draco engines that will be needed for the mission. This result may also cause the need to look into a different type of engine that may produce more thrust. Once all of that has been determined then an investigation into attitude thrusters will commence.

Space Payloads: Further developments in the Space Payloads subsystem include the communication payload parameters being finalized, such as the types and orientations of both the primary high gain antenna and the backup set of low gain antennas. Also, further research into the SLS’s payload adaptor system will be conducted in order to optimize Asclepius’ payload adaptor.

Ground Control: Further work must be done to see how challenging it will be to locate the Payload operations control center at NASA Goddard. Exact figures should be compiled for the on ground resources that would be required for the mission.

Communications: The location and mounting of the high gain antenna must be solidified as well as a coding scheme. Furthermore, the exact number and type of low gain antenna must be determined for backup and communications with the receiving spacecraft. A communications protocol must be devised between Asclepius and the receiving spacecraft. Furthermore the feasibility of using Goddard’s Space Network over the current plans to use the Deep Space network must also be researched for rendezvous.

C&DH Protocols for each system function will be mapped out to understand how Asclepius will react to malfunctions in the mission. This will provide a basis for the written-in mission software. Also, researching and designing a plan for how the command module will manage the collection of solar power and battery charging.

Power: In the future, an investigation into batteries and solar cells that are being developed today that could be available before 2029 will be researched. They could become viable options, especially if the mass of the battery or solar panel can be reduced. The type of solar panel, rigid or flexible panel will also be further investigated. The decision is reliant upon the exact dimensions of the spacecraft. Once that decision is made then an analysis will be conducted on the orientation of the solar panel and the expected amount of sun that the panels will receive during the mission.

Thermal Future work within the thermal control subsystem will require a more detailed analysis of the thermal condition of Asclepius throughout the whole mission. For example, analysis will need to be performed for when Asclepius is on Earth and LDRO. To obtain more accurate predictions of Asclepius’ thermal state, internal heat radiation from components will need to be determined, along with, more accurate spacecraft dimensions and material determination. When these factors are further defined, absorptivity and emissivity will be more accurately represented in the temperature calculations. With more accurate temperature analysis the next step will be to determine exact thermal control methods and their placement, so for example, on the command module the number of louvers and their location will be determined.

Conclusions The ResuppLION missions' objectives are to launch a resupplying propellant spacecraft, Asclepius, from the Earth, maneuver this vehicle into a Lunar Distant Retrograde Orbit, and then rendezvous with the University of Texas' Deep Space Transport vehicle for a successful fuel transfer. This line of missions will support one crewed orbital Mars mission every other opportunity beginning in 2033 and Mars surface mission cargo delivery beginning in 2037, while having the ability to expand. The first mission will involve the SLS Block 2 using a trans-lunar injection to send Asclepius on a trajectory toward the moon, where it can then use a combination of propulsive maneuvers and lunar gravity assists to enter into the LDRO needed to meet with the University of Texas' vehicle. Once in the LDRO, Asclepius will berth with the Deep Space Transport with the aid of a robotic arm attached to Asclepius' docking module. Following a successful berthing, Asclepius will collapse the silicone rubber tank that is storing the refueling propellants with structural panels attached to collapsing actuators. This will "squeeze" the liquid propellants through tubing that is routed to the Deep Space Transport's fuel tanks. Once the tanks are completely collapsed, Asclepius will undock from the Deep Space Transport and dispose of itself in a safe manner. Upon the feedback gathered from this first mission and the research that will be done up until the first mission's launch, the ten subsystems that are coordinating each mission phase will be iteratively updated to accomplish the ResuppLION mission objectives as efficiently as possible. As a lasting impact, the ResuppLION missions hope to begin paving the road for some future where deep space exploration vehicles have the ability to be refueled so they may continue their endeavors for longer periods of time.

Appendix 1.1 Link Budget Link equation: 퐸 푃퐿 퐺 퐿 퐿 퐺 푏 = 푙 푡 푠 푎 푟 푁0 푘푇푠푅 Transmitter power(P): P 20W

Transmitter to antenna line loss(푳풍)

퐿푙 0.8 (estimated) (Wertz et al)

Spacecraft Transmit and receive antenna gain(푮풕, 푮풓) = -159.59+20log(D)+20log(f)+10log(휂)

퐷푟 Variable (calculated to be 0.6m)

휂 (antenna efficiency) Estimated 0.6 (Wertz et al)

Name Frequency (GHz)

S 2.5-2.69

K 12.5-31

4휋퐴휂 Ground transmit and receive antenna gain(푮 , 푮 ) = 10log( ) 풕 풓 λ2 Deep Space Network Ka-band 85.87db (Nasa facts) S-band 76.06db

Space Loss(Ls) = 147.55-20log(405,696,000)-20logf frequency Max loss

S -213.2091dB

Ka -227.8414dB

Ku -234.4412dB

La (Transmission path loss): S 70^-1 dB

Ka 30^-1 dB

Ku 0.5 dB (Wertz et al)

푇0(1−퐿푟) 푇0(퐹−1) 푇푟 System Noise Temperature (푻풔) = 푇푎푛푡 + ( ) + ( ) where F = 1 + 퐿푟 퐿푟 푇0

퐿푟 (line loss antenna-receiver)

frequency Downlink Uplink

S 0.5 0.5

K 0.5 0.5

(Wertz et al)

푇푟 (receiver noise temperature)

frequency Downlink Uplink

S 75K 289K

K 289K 289-438K

(Wertz et al)

퐿0 (reference temperature) 290K (Wertz et al)

푇푎푛푡 (antenna temperature) Assumed small (20K)

Data rate(R)

Command 64bits once every second: 64bps

Health 100Kbps

푬 Energy per bit to noise power spectral density ratio ( 풃) 푵ퟎ With desired Bit Error Rate(BER) of10−5 QPSK (chosen since used by Lunar 9.6db(theoretical)+(1~3db(error))) reconnaissance orbiter)

Energy per bit to noise power with a 0.6m parabolic dish: S band 12.4043db Ka band 12.4043db

References Structures and GNC [1.1] NASA Glenn Research Center. (n.d.). Bernoulli's Equation. Retrieved from https://www.grc.nasa.gov/www/k-12/airplane/bern.html [1.2] International Docking System Standard (IDSS) Interface Definition Document (IDD) Revision E . (2016, October). Retrieved from http://www.internationaldockingstandard.com/download/IDSS_IDD_Revision_E_TAGGED.pdf [1.3] Canadian Space Agency. (n.d.). Canadarm2, the Canadian Robotic Arm of the International Space Station. Retrieved from http://www.asc- csa.gc.ca/eng/iss/canadarm2/default.asp [1.4] Zhang, H., “The Permeability Characteristics of Silicone Rubber,” Arlon Silicone Technologies Division, 302-834-2100, Bear, DE, 2006. [1.5] AZoM. (2001, September 25). Silicone Rubber. Retrieved November 08, 2017, from https://www.azom.com/article.aspx?ArticleID=920 [1.6] ProTech Composites. (2016). Carbon Fiber 101. Retrieved November 26, 2017, from http://www.protechcomposites.com/pages/About-Carbon-Fiber.html [1.7] Serapid. (2012, August). RollBeam telescopic linear actuator (Tech.). Retrieved November 26, 2017, from Serapid website: http://www.serapid.com/sites/default/files/public/product- documentation/en_rollbeam.pdf [1.8] Serapid. (2017). Push Pull Systems. Retrieved November 28, 2017, from http://www.serapid.com/en/industrial-equipment/horizontal-motion/horizontal-motion- systems/push-pull [1.9] Parma, G. (2011, May 20). Overview of the NASA Docking System and the International Docking System Standard (Symposium Presentation). Retrieved November 26, 2017, from NASA website: https://web.archive.org/web/20111015075220/http:/dockingstandard.nasa.gov/Document s/AIAA_ATS_NDS-IDSS_Overview_Draft1.pdf [1.10] NASA. (2015). SpaceX CRS-7: Seventh Commercial Resupply Services Flight to the International Space Station (Rep.). Retrieved November 26, 2017, from https://www.nasa.gov/sites/default/files/atoms/files/spacex_crs7_mission_overview.pdf [1.11] Boesinger, E. A., & Hakun, C. (2014, May 14). 42nd Aerospace Mechanism Symposium (Symposium Outline). Retrieved November 26, 2017, from NASA website: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20140008875.pdf

Launch Vehicle [2.1] Smith, D. A., “Space Launch System (SLS) Mission Planner’s Guide,” NASA M17- 6014, April 12, 2017. [2.1] Wertz, J. R., Everett, D. F., & Puschell, J. J. (2011). Overview of Payload Design. In Space Mission Engineering: The New SMAD (pp. 439-455). Hawthorne, CA: Microcosm Press. [2.2] Wertz, J. R., Everett, D. F., & Puschell, J. J. (2011). Communications Payloads. In Space Mission Engineering: The New SMAD (pp. 455-493). Hawthorne, CA: Microcosm Press. [2.3] Smith, D. A., “Space Launch System (SLS) Mission Planner’s Guide,” NASA M17- 6014, April 12, 2017. Power: [3.1] Chai, P. R., Merrill, R. G., & Qu, M. (2016). END-TO-END TRAJECTORY FOR CONJUNCTION CLASS MARS MISSIONS USING HYBRID SOLAR- ELECTRIC/CHEMICAL TRANSPORTATION SYSTEM. AAS,AAS(16), 255th ser. Retrieved September 22, 2017, from https://ntrs.nasa.gov/search.jsp?R=20160007700. [3.2] Belburno, E. A., & Carrico, J. P. (2000). Calculation of Weak Stability Boundary Ballistic Lunar Transfer Trajectories. AIAA Astrodynamics Specialist ,AIAA(2000), 4142nd ser., 1-11. [3.3] Merrill, R. G., Qu, M., & Chai, P. R. (n.d.). An Integrated Hybrid Transportation Architecture for Human Mars Expeditions. AIAA . Retrieved November 22, 2017, from https://ntrs.nasa.gov/search.jsp?R=20160006318. [3.4] James, M., Salton, A., & Downing, M. (2013). Noise Study in Support of the EA for Issuing an Experimental Permit to SpaceX for Operation of the DragonFly Vehicle at McGregor Test Site, Texas. Blue Ridge Research and Consulting,R(2014). Retrieved November 23, 2017. Ground Control [4.1] Garner, R. (2015, October 06). Goddard Missions - Present. Retrieved November 28, 2017, from http://www.nasa.gov/content/goddard-missions-present [4.2] NASA Facts- Deep Space Network [Pamphlet]. (n.d.). Pasadena , CA: Jet Propulsion Labratory.

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[5.2] Link Budget. (n.d.). Retrieved November 16, 2017, from http://propagation.ece.gatech.edu/ECE6390/project/Fall2010/Projects/group7/Project%20Web site_v3_files/Page550.htm [5.3] Slobin, S. D. (n.d.). 104 34-m BWG Stations Telecommunications Interfaces. Retrieved November 15, 2017, from deepspace.jpl.nasa.gov/dsndocs/810-005/104/104H.pdf [5.4] Slobin, S. D. (n.d.). 101 70-m Subnet Telecommunications Interfaces. Retrieved November 16, 2017, from deepspace.jpl.nasa.gov/dsndocs/810-005/101/101E.pdf Command and Data Handling [6.1] Radiation-hardened space electronics enter the multi-core era. (2017, June 21). Retrieved November 28, 2017, from http://www.militaryaerospace.com/articles/print/volume-28/issue- 6/technology-focus/radiation-hardened-space-electronics-enter-the-multi-core-era.html [6.2] UART vs SPI vs I2C | Difference between UART,SPI and I2C. (n.d.). Retrieved November 28, 2017, from http://www.rfwireless-world.com/Terminology/UART-vs-SPI-vs- I2C.html [6.3] BAE Systems moves into third generation rad-hard processors. (2002, May 1). Retrieved November 28, 2017, from http://www.militaryaerospace.com/index.html [6.4] Knapp, G. (n.d.). Radiation-hardened processors products. Retrieved November 28, 2017, from http://www.baesystems.com/en-us/our-company/inc-businesses/electronic- systems/product-sites/space-products-and-processing/processors [6.5] BAE Systems - RAD 5545 SpaceVPX single board computer. Retrieved November 28, 2017, from http://www.baesystems.com [6.6] Walls, C. (2014, September 12). Embedded Systems Programming Languages | EE Times. Retrieved November 28, 2017, from https://www.eetimes.com/author.asp?doc_id=1323907&page_number=2

[6.7] I2C - What's That? (n.d.). Retrieved November 28, 2017, from https://www.i2c-bus.org/ GNC [7.1] Wiktor, P., Chen, J. H., & DeBra, D. (1989). Optimal Thruster Configurations for the GP- B SPacecraft. Automatic Control in Aerospace, 203. Retrieved from https://books.google.com/books?id=3NbSBQAAQBAJ&pg=PA203&lpg=PA203&dq=O ptimal Thruster Configurations For The Gp-B Spacecraft.&source=bl&ots=2Y6JDzv1iJ&sig=DWtkvlTTAvrpl_JkOoZGxUbW7YQ&h l=en&sa=X&ved=0ahUKEwjqppubkdvXAhXH7YMKHVB0Cv8Q6AEIPDAD#v=onep age&q=Optimal%20Thruster%20Configurations%20For%20The%20Gp- B%20Spacecraft.&f=false. [7.2] Heacock, R. L. (1980). The Voyager Spacecraft. Proceedings of the Institution of Mechanical Engineers, 194. Retrieved from http://journals.sagepub.com/doi/pdf/10.1243/PIME_PROC_1980_194_026_02

[7.3] John Hopkins University Applied Physics Lab. (n.d.). New Horizons Spacecraft Systems and Component. Retrieved from http://pluto.jhuapl.edu/Mission/Spacecraft/Systems-and- Components.php [7.4] Neptec. (n.d.). Rendezvous and Docking Sensors. Retrieved from https://neptec.com/products/docking/ [7.5] Lemoine, F. G., & NASA GSFC, S. (n.d.). DORIS System Overview. Retrieved from https://space-geodesy.nasa.gov/docs/2012/DORISOverview_Lemoine_120607.pdf Power [8.1] Surampudi, R., Hamilton, T., & O. (2001). Solar Cell and Array Technology for Future Space Missions. NASA OSS,(D), 24454th ser., 1-94. [8.2] IEE Xplore (n.d.). Retrieved November 24, 2017, from http://ieeexplore.ieee.org/stamp/stamp.jsp?arnumber=4161575 [9.1] Thermal Systems - Mars Reconnaissance Orbiter. (n.d.). Retrieved November 15, 2017, from https://mars.nasa.gov/mro/mission/spacecraft/parts/thermal/ [9.2] Kolker, I., & Altemose, G. (2010, December 2). Space Health Monitoring Using Analog Multiplexers and Temperature Sensors. Retrieved November 11, 2017, from http://ams.aeroflex.com/pagesproduct/appnotes/mux/ApNote_an8500-4a.pdf [9.3] Sierra Nevada Corporation’s Space Systems. (2015). Retrieved from http://mediakit.sncorp.com/mediastore/document/Space-Technologies-Product- Catalog.pdf Space Technologies Product Catalog 2015, Rev 5 [9.4] Cobham. (2009). Xenon Propellant Tank. Retrieved from http://www.cobham.com/mission-systems/composite-pressure-solutions/space- systems/xenon-propellant-tank-datasheet/docview/ [9.5] NASA, Finkenor, M. M., & Dooling, D. (1999, April). Multilayer Insulation Material Guidelines. Retrieved from https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19990047691.pdf

General [10.1] Wertz, J. R., Everett, D. F., & Puschell, J. J. (2011). Space Mission Analysis and Design. Hawthorne, CA: Microcosm Press.