CranSpace

Idriss SISAÏD Edward ANASTASSACOS

Anees AHMAD Laia RAMIÓ

Kaitlyn MORRIS Enrique GARCIA BOURNE

Daniel PASTOR Mairéad BEVAN

Thomas AURIEL Jack LONGLEY

International Student Design Competition

INSPIRATION MARS

Cranspace.com Inspiration Mars Student Design Competition Section 0

Acknowledgments

We would like to thank Dr. Jennifer Kingston (Cranfield University) for her support through- out the project, particularly for the advice she gave us regarding cost estimation; Professor Michael J. Rycroft (University of Cambridge) for his keen interest in the project and con- tributions especially with regard to radiation shielding components. We also extend our thanks to Dr. Donald Rapp (former research professor, University of Southern California) for his professional insight and in-depth knowledge specifically regarding radiation related to Mars travel. Furthermore, Dr. Don Boroson (MIT Lincoln Laboratory) for sharing his ex- pertise on laser communication systems and Dr. Arturo Casadevall (Albert Einstein College of Yeshiva University) for his contribution to our understanding of the shielding proper- ties of Melanin Nano-Shells. Finally, we would like to thank Dr. Joseph Hamaker (Senior Cost Analyst - SAIC / The Millennium Group International; former Director of Cost Anal- ysis Division, NASA) for his invaluable contribution to the development of our cost analysis..

We would also like to acknowledge the contributions made to the project by Andrew McMil- lan and Daphne De Jong.

i Inspiration Mars Student Design Competition Section 0

Contents

1ReportOverview 1

2 Inflatable Space Habitat 2 2.1 General Assumptions and Considerations of Sundancer ...... 2 2.1.1 Specifications ...... 2 2.2 AdvantagesandDisadvantages...... 4

3 Radiation Shielding 5 3.1 Radiation and Water ...... 6 3.2 WasteManagementSystem ...... 11 3.2.1 Waste Collection and Delivery ...... 11 3.3 ApplicationofMelanin ...... 12 3.4 SPE Events ...... 13 3.5 MagneticFieldProbe...... 13 3.6 SPE Shelter ...... 13

4 Internal Housing and ECLSS 14 4.1 Bubble Configuration ...... 14 4.2 Food Management ...... 16 4.2.1 FoodRequirements ...... 16 4.2.2 Food Storage ...... 17 4.3 WasteManagementConfiguration...... 17

5 Communications 18 5.1 Traditional Communications System ...... 18 5.2 Remote Communications System ...... 19 5.3 CommunicationSatelliteDesign ...... 20 5.3.1 Design Drivers ...... 20 5.3.2 Design ...... 20 5.3.3 AddedPayloads...... 21 5.3.4 Structure ...... 22 5.3.5 Power ...... 22

6 Science 24 6.1 Extended Boom and General Sensors ...... 24 6.2 Analysis of the e↵ectivenessofMelaninNano-shells ...... 24

ii Inspiration Mars Student Design Competition Section 0

6.3 The E↵ects of Deep Space Radiation on the Growth of Plants ...... 24 6.4 CrewActivity...... 25 6.4.1 Psychological evaluation ...... 25 6.5 Laser-based Communication System Demonstration and Mars Reconaissance 25 6.5.1 Technology Demonstration ...... 25 6.5.2 Scientific Observations ...... 25 6.5.3 Earth Re-entry ...... 26

7 Launch Vehicles 26 7.1 Launcher Comparison ...... 26 7.2 MaximumNumberofLaunchesandLEOLaunchMass ...... 27 7.3 In-Orbit Propulsion ...... 27 7.3.1 Delta-V Budget ...... 27 7.3.2 In-SpacePropulsionSystem ...... 28

8Reentry 29 8.1 Reentry Vehicle Configurations ...... 29 8.2 Reentry Trajectory ...... 30 8.3 HeatFlux ...... 31 8.4 HeatLoad...... 32 8.5 TPS material ...... 32

9 Electrical Power Systems 33 9.1 Dragon ...... 33 9.1.1 Power Generation ...... 33 9.1.2 Energy Storage ...... 33 9.2 Sundancer ...... 34 9.2.1 Power Consumption ...... 34 9.2.2 Life Support ...... 34 9.2.3 Waste Management ...... 34 9.2.4 Communications ...... 35 9.2.5 CrewActivities/Science...... 35 9.3 PowerGeneration...... 35 9.3.1 Case A ...... 36 9.3.2 Case B ...... 36 9.4 Energy Storage ...... 36

iii Inspiration Mars Student Design Competition Section 0

10 Cost Evaluation 37 10.1Firstapproach...... 37 10.2 Detailed Cost Study - QuickCost 5.0 Model ...... 37 10.2.1 Main Spacecraft ...... 37 10.2.2 CommunicationsSatellite ...... 38 10.3 QuickCost Results ...... 39 10.4 Cost Scheduling ...... 39 10.5 Validation ...... 39 10.5.1 Communicationssatellite...... 39 10.6 Mainspacecraft-NAFCOMWeight-BasedModel ...... 40

11 Mission Phases and Scheduling 42 11.1 Development Schedule ...... 42 11.2 Launch and Docking Schedule ...... 43 11.2.1 Launch and mass distribution ...... 43 11.2.2 Launch sequence ...... 44 11.2.3 Docking and in-orbit assembly ...... 44 11.3 Interplanetary Events ...... 45 11.4 Risk mitigation and 2021 launch window ...... 47

12 Conclusion 48

iv Inspiration Mars Student Design Competition Section 1

1ReportOverview

The following report proposes a design for a two-person Mars fly-by mission to be launched in 2018. Building on the foundation laid out by Inspiration Mars, Cranspace has sought to resolve areas of weakness and propose the designs outlined in this report. The key ar- eas addressed include the provision of a suitable environment for the crew, ensuring safety from radiation, the identification of science experiments that can be carried on board, and an analysis of the re-entry phase, whilst ensuring that the cost of the mission is as low as possible. Other areas such as launch vehicles and propulsion systems are discussed as they need to accommodate these systems.

As minimising the cost of the mission is one of the main overall objective, o↵-the-shelf technologies and structures have been incorporated wherever possible. One such structure is the Sundancer Inflatable Habitat by Bigelow Aerospace. As Bigelow has ceased development of the Sundancer in favour of developing the larger BA 330 module, the cost analysis in this report has allocated funding for the completion of its development []. However, this cost will be greatly reduced as Bigelow has used similar technology to develop the BA 330. The Sundancer will contain six bubble rooms that will accomodate the astronauts. The purpose of these bubbles is primarily radiation protection. They will be surrounded by water, which will be transferred between one bubble and the next as the crew moves between them. This system will aim to minimise the amount of water needed for shielding on the spacecraft. For further shielding, the waste management system will treat solid waste and store within tubes around the spacecraft. The radiation shielding and waste management systems are areas for which development costs will be higher.

The mission will also carry a small satellite carrying a laser communication payload that will enable a high data transmission rate during the interplanetary trajectory. This satellite will be left in Martian orbit as a future communications relay from the Red Planet.

SpaceXs Dragon capsule was chosen as a re-entry vehicle. Minor modifications will be required to resist the high re-entry velocity. A literary survey of materials used in re-entry has been made.

The mission will employ four Falcon Heavy launches, and will cost an estimated 2.1 to 2.85 billion FY14 U.S. dollars. There are a number of constraints that have, so far, limited man’s reach in the solar system and a great many considerations that need to be made in

1 Inspiration Mars Student Design Competition Section 2 order to ensure the well-being of those that take the risk to travel out there. This mission is intended to act as a technological stepping stone, demonstrating the necessary techniques and providing the technologies as the next phase for the human race to finally fulfil the universal dream of establishing an extraterrestrial human colony on Mars.

This report considers how to overcome the constraints of interplanetary travel such as galac- tic cosmic rays, solar particle events, waste management problems and the associated cost with a mission of this magnitude. Configurations of primary subsystems are proposed to suit the requirements of the astronauts. Food management is investigated and scientific experiments are suggested to take advantage of the spacecraft’s unique position. Following a comprehensive literary survey of Launch Vehicles, this report outlines the requirements and parameters of the launch configurations, engine specifications, capabilities and payload constraints.

For additional information, simulations and videos, please visit our website: http://www.cranspace.com

2 Inflatable Space Habitat

CranSpace design proposes the implementation of the Sundancer inflatable module [1] de- veloped at Bigelow Aerospace [2]. Bigelow Aerospace are currently preparing for the first rendezvous of an inflatable spacecraft with the International Space Station [3] providing a system demonstration of the technology to be implemented on-board the Bigelow BA 330 [4] and the Sundancer [1]. The Sundancer has been designed to accommodate up to 3 people in Earth Orbit. This section considers the advantages and disadvantages of implementing the Bigelow Sundancer module and possible modifications to ensure that it is suitable for long-term deep space flight.

2.1 General Assumptions and Considerations of Sundancer

The assumptions made of the Sundancer’s capabilities are presented below and are based on the specifications outlined by Bigelow Aerospace [2] (see table 1).

2.1.1 Specifications

The assumptions made of the Sundancers capabilities are presented below and are based on the specifications outlined by Bigelow Aerospace (see table 1).

2 Inspiration Mars Student Design Competition Section 2

Occupancy Up to three on a long-term basis Volume 180 m3 Radiation Protection Bigelow Aerospaces shielding is equivalent to or better than the Inter- national Space Station and substantially reduces the dangerous impact of secondary radiation. Ballistic Protection The Sundancer utilises an innovative Micrometeorite and Orbital De- bris Shield. Hyper-velocity tests conducted by Bigelow Aerospace have demonstrated that this shielding structure provides protection superior to that of the traditional aluminium can designs. Propulsion The Sundancer utilises two propulsion systems on the fore and aft of the spacecraft. The aft propulsion system can be refuelled and reused. Electric Power Every Sundancer habitat will include an independent power system comprised of solar arrays and batteries. Avionics Each module will contain an independent avionics system to support navigation, re-boost, docking, and other manoeuvring activities. Environment Control and Each Sundancer will contain its own independent ECLS system includ- Life Support System ing lavatory and hygiene facilities. Windows Sundancer will boast four large windows coated with a film for UV protection, providing an unparalleled opportunity for both celestial and terrestrial viewing. Table 1: Specification of the Sundancer module [1]

Radiation Protection The radiation shielding specifications provided by Bigelow Aerospace are assumed to be limited as there is no publicly available documentation regarding its capa- bilities. Whilst it is true that testing on board the ISS will occur, radiation levels within LEO are considerably lower than those experienced beyond the natural shielding of the Earth’s magnetic field [5]. It is therefore necessary to consider additional passive and possibly active radiation shielding in order to provide a degree of protection in the case of Solar Particle Events (SPE) (see section 3).

Ballistic Protection It is assumed that the ballistic protection provided by the Sundancer is capable of shielding the spacecraft from micro-meteoroids whilst in deep space. The only additional ballistic protection methods that will be considered are collision avoidance manoeuvres, rather than additional passive ballistic shielding.

Propulsion It is assumed that the in-built systems are sucient for manoeuvres such as collision avoidance, as well as emergency procedures that are initiated in the case of solar flares and coronal mass ejections (see section 7).

Electrical Power The electric power is sourced from solar arrays and batteries. It has been assumed that the solar arrays and batteries can be configured to satisfy the electrical power requirements of the mission. A assessment into the required sizing of the solar arrays and batteries is made and presented (see section 9).

Life Support While the life-support systems have limited available specifications, they are undergoing initial human trials [6]. Life-support systems are therefore expected to be

3 Inspiration Mars Student Design Competition Section 2 ready by 2018. Modifications are expected to occur in order to accommodate additional life sustaining measures (see section 4).

Figure 1: Design outline of the Sundancer Inflatable module. Developed with Sketchup.

2.2 Advantages and Disadvantages

Volume The Sundancer inflatable module provides the astronauts with a dramatically larger available volume that can be used for both storage and living space. This spatial advantage provided by Inflatable Habitats over more traditional ”can-like” structures is an important one, especially considering the long flight duration. Cramped conditions similar to the ATV [7] or Orion [8] would dramatically reduce moral on the spacecraft and such continuous close proximity contact between astronauts may induce conflict over time in what will be an already stressful environment. Additionally, Bigelow Aerospace inflatable spacecraft have been designed to be inflated in-orbit after being stored in low volume rocket fairings.

Developed Sub-Systems By incorporating the Sundancer, a large number of the sub- systems, including the critical life support system, are being independently developed by Bigelow Aerospace, relieving development costs. Modifications are expected, however, and coordination with Bigelow Aerospace is needed to tailor aspects of various system designs.

4 Inspiration Mars Student Design Competition Section 3

Multilayer Structure The multilayer structure nature of the Sundancer’s exterior acts as an external thermal blanket, helping to keep the spin-stabilised vehicle’s interior at a constant temperature while shielding the spacecraft’s systems and inhabitants from micro-meteoroids and space debris. Studies suggest that inflatable structures provide better shielding than is available on-board the ISS and also provides significant protection against radiation [9]. A survey of these studies indicate that the exterior structure is to be made from kevlar-like materials, Vectran, Nexttel and polyethylene.

Development Status The Sundancer’s development has been frozen as Bigelow Aerospace have focused their attention to larger vehicles. However, there have been continuous advances of the sister BA 330 vehicle which is expected to be launched as early as 2015 [3]. The Sun- dancer is expected to use near identical systems implemented by the BA 330 and so, if Bigelow Aerospace could be convinced (financially)intoresumingdevelopmentprocedures- the design could be fast-tracked using technology demonstrated by the BA 330.

The overall layout of the internal structure of the Sundancer is expected to be modified in order to address mission constraints such as space and shielding requirements. Given the nature of the inflatable of the mission, much of the assembly of the internal components need to be made in-orbit. This is explained in detail in the sections related to the individual sub-systems.

3 Radiation Shielding

One of the most dicult obstacles to overcome in interplanetary travel is the need to pro- vide adequate shielding to protect crew and equipment from radiation. The main radiation sources of radiation in deep space are heavy ions fully ionised atomic nuclei of galactic cosmic rays (GCR) and energetic protons originating from SPE. This secondary radiation e↵ects occuring after initial exposure. Without adequate protection, astronauts can su↵er from acute radiation syndrome (ARS) or radiation sickness, damage to the central nervous system, cancer and death. In practice, such shielding often requires a large mass and high development costs. NASA regulations state that astronauts can be exposed to no more than 3% of added Risk of Exposure-Induced Death (REID) [10].

5 Inspiration Mars Student Design Competition Section 3

Figure 2: Image illustrating approximate values of radiation levels with increasing distance from the Earth’s surface. Units are displayed as a log scale of annual levels of rems. Image taken from Scientific America.

The spacecraft will traverse a number of di↵erent radiation environments through the du- ration of the mission, most notably the Van Allen Radiation Belts (see Figure 2). These consist of a large number of highly energetic charged particles [11]. In the Apollo missions, astronauts who traversed the belt regions were largely unprotected from this radiation. De- spite their limited exposure time, they received the majority of their total mission dose while passing through the inner belts. This dose varied through a range of 1.6 to 11.4 mGy less than the limit set by the United States Atomic Energy Commission for people working with radioactive materials [12]. A similar approach is used in this mission design, whereby the spacecraft performs its delta-v before passing through the belts, so as to minimise the time spent within them.

3.1 Radiation and Water

The level of radiation protection provided by a material is measured by the extent to which it is penetrated by energetic particles [10]. Other important factors that must be taken into account include the chemical properties of the material and the way in which the human inhabitants must manage and interact with it. Aluminium is often considered as the material of choice, but has a high mass and still produces secondary neutrons following exposure to radiation. Polyethylene, a hydrogen rich material, has been successfully tested on board the ISS [13] and is therefore a candidate for radiation protection. Water is also capable of absorbing a high amount of energy while emitting a low number of neutrons, although it has not been used in previous missions due to the large amount that would be required. It was also deemed unsuitable for exterior mounting on a spacecraft due to its structural and thermal properties [14]. There are, however, a number of advantages to using water to shield

6 Inspiration Mars Student Design Competition Section 3 against both GCRs and SPEs. Figure 4 compares waters shielding properties to those of other materials.

Figure 3: Dose equivalent as a function of wa- Figure 4: Point estimates of 5 cm depth dose ter shield thickness resulting from galactic cos- for GCR at Solar Minimum as a function of mic rays [15]. areal density for various materials [16] .

Water has the added advantage of being in a liquid state at room temperature and atmo- spheric pressure. It can therefore easily be transported around the spacecraft in order to provide continuous and varying levels of shielding thickness in di↵erent parts of the habitat, depending on the local radiation levels. For future designs, this concept could be used as an adaptive shielding mechanism that could provide a thin level of shielding for regions of space with low incident flux, and thicker shielding for regions with higher flux. In order to reduce complexity, a simple approach has been used within this report. It is worth noting that whilst moderate amounts of shielding reduce the dose of GCRs due to the removal of lower energy components, the e↵ectiveness of shielding approaches diminisihing returns be- yond 25 cm2 (see Figures 3 and 4). This design uses water as a shielding medium within the Sundancer - essentially acting as a secondary shielding layer for the entire habitat module beneath its external structure. It has been previously mentioned that the radiation shielding on board the inflatable modules provided by Bigelow Aerospace is considered as incapable of suciently shielding astronauts in deep space. Whilst the ballistic and radiation shielding properties of the materials used are to be demonstrated on the ISS by 2015 [3], it will not be subjected to the same levels of radiation as experienced in deep space. Even so, information regarding the structural properties of the Sundancer is limited, correlations have been made with previous studies and reports concerning inflatable structures in space. The European Space Agency’s (ESA) Aurora Exploration Program has provided a considerable amount of information regarding the comparison between inflatable and rigid vehicle structures, and

7 Inspiration Mars Student Design Competition Section 3 states that the material structures of inflatable vehicles is expected to be based on polyethy- lene, polyacrylate, ethylene vinyl alcohol (EVOH), Kevlar and Nomex [9]. The thickness of the Sundancer’s external layer is expected to be around 40-50 cm when fully inflated. According to both the Hadronic-Bertini and the Hadronic-Binary physics models, the level of shielding provided by this structure would be comparable to a similar thickness of water, as shown in Figure 4. Accounting for the various considerations listed above, an internal radiation-resistant housing system has been designed, as shown in Figure 5.

Figure 5: Simplistic design of the encased Bubble Structure within the Sundancer Inflatable Module. Image developed using Sketchup.

Radiation can approach the spacecraft from any direction. The system’s additional shielding comes in the form of dome-shaped ”bubble” rooms that are surrounded by water to protect the astronauts from incoming GCRs. As the astronauts move into neighbouring rooms, the water is flushed into the relevant bubble. This method will provide the astronauts with adequate shielding in all directions at all times, while still taking advantage of the space available inside the Sundancer. The general description of this system can be separated into two sections: the cylindrical core structure and the bubble modules.

I. Cylindrical Core This is a component of the Sundancer’s internal structure which must be modified prior to launch and adapted in order to accommodate mission critical components such as the life-support and central computing systems. The modified core acts to support the network of tubes which provide a connection between the bubble modules, the waste management system and the water systems. The core itself is segmented into 3 sections separated by a hatch. Each section is connected in parallel to the life support and central computer systems so that all bubble structures and core sections have access to the on board data handling system as well as a continuous supply of oxygen. At each end of the core, any docked structures (e.g. the Dragon capsule) can be accessed via a pressurised hatch. A comparison between the original and modified core structure is presented in Figure

8 Inspiration Mars Student Design Competition Section 3

6. The modified core is structured so that the original four separate pillars are combined to provide two larger support pillars in order to allow for easy deployment of the bubble modules (see Figure 7). The gaps in the core allow the astronauts to enter and exit the ’bubbles’ while providing the same, continuous level of protection.

Figure 6: A simplified design of the original core (left) compared with the modified core structure (right). This figure illustrates the gain in space available to the astronauts when transitioning between bubble structures.

II. Bubble Modules: General Configuration Each pair of bubbles is assembled so that they are oriented on opposite sides of the central core. Two larger sets of bubble mod- ules will have external and internal diameters of 2.9 m and 2.4 m respectively, providing a shielding thickness of 25 cm (or 25 g/cm2) and a smaller bubble set will have an external and internal diameter of 2.56 m and 2 m respectively, providing a shielding thickness of nearly 30 cm. The overall configuration of the bubble pairs is described in Section 4.1. The bubbles have been designed with the following considerations in mind: robustness, ease of repair, ease of construction, ease of stowing, ease of water transportation, simplicity and cost. The design is based on the truncated icosahedron structure - essentially two tiled buckyballs, the larger one making up the exterior surface and the smaller one providing the interior surface. The final configuration is depicted in Figure 7. The surface tiles are composed of 3mm thick composite panels. The opening of the structure (the bottom face in Figure 7) is directly connected to the core of the Sundancer. The tiles of adjacent bubble pairs are connected to allow water to flow between them. The arrangement of the tiles is shown in 8. The edges of the tiles are lined with solid poles, which will provide the structure with a level of rigidity once mounted. To ensure adequate ventilation, small gaps in the surface are provided, with fans to allow for the circulation of air.

This design allows the structure to be folded so that it is compact. These bubbles will then be assembled onto the core before launch. The dimensions of the plates have been

9 Inspiration Mars Student Design Competition Section 3 designed so that the overall width of the system is less than 40 cm when stowed. Each group of bubbles is then folded in order to minimise the deformation of the flexible pipes between the blocks. The assembly procedure is as follows: following the docking of the Dragon to the Sundancer, and inflation of the latter, the astronauts will enter. They will then connect each set of bubbles to the core. Their edges will then be connected as described above.

Figure 7: Illustration of the bubble’s filled and empty constituent blocks.

The walls of each bubble can be compressed in order to expel its water contents. This controlled compression will be performed in a predefined sequence, with the furthest away from the core being the first. This will ensure that no trapped water remains in the structure. Once the water has been flushed and all the structures emptied, the valves will be shut.

Figure 8: The illustration on on the left provides the mechanims for directing water within a tile group. The arrows represent a force ontho the surface compressing the water and forcing a displacement out of the tile group. With the valve open, water is confined to flow in the direction shown by the arrow. These groups are laid out as shown on the right.

10 Inspiration Mars Student Design Competition Section 3

3.2 Waste Management System

Human waste produced during the journey must be stored and managed carefully. Current methods used on board the ISS involve disposing of solid waste via a freighter spacecraft. Paragon Space Development Corporation has proposed a system that uses ionomer mem- brane ”bags” [17] to extract water from human waste through reverse osmosis. These bags can be used to provide a source of radiation shielding by lining the spacecraft walls with them. However, the need for astronauts to manually interact with faecal matter raises a number of health issues. The following solution builds on this idea while employing a simple yet innovative system that provides both radiation shielding in an automated and ecient manner to manage waste.

3.2.1 Waste Collection and Delivery

The flow mechanism of the delivery system relies on a peristaltic motion to push waste through flexible tube structures. This method was inspired by mammalian intestines. The peristaltic motion is generated by the activation of concentric muscles, and drives solid waste through the digestive system. The same principle will be applied to the waste delivery system.

This system utilises the same material that makes up the ”bags” developed by Paragon. However, rather than a bag-like container, the material is extended so that it forms a tube- like structure with an approximate length of 1.5 km and a diameter of 2 cm (see table 2). The ionomer membrane slowly extracts water from the system through reverse osmosis. This water is then collected for filtration and reuse. The tube is contained in an external vessel designed to insulate the waste-containing inner tube. The peristaltic motion of the waste management system is induced by inflatable rings positioned at intervals along the tube. Positioned at intervals along the tube the synchronised contractions and relaxations along the conduit induce the forward motion of the solid waste (see Fig. 9). The total eciency of the water extraction procedure increases as the waste travels further along the tube, as the contact surface increases allowing for more water to be absorbed.

Inner pipe diameter 2cm Outer pipe diameter 5cm Dry mass 570 g/m Loading mass 2500 g/m System Length 1.5 km Waste volume 0.5 m3 Table 2: Dimensions of the waste management delivery system minus integration with space toilet.

11 Inspiration Mars Student Design Competition Section 3

Figure 9: Illustration of the artificial intestine. Schematic representation of the waste management system On this schematic, only the inner pipe, toric ring and bracing to maintain in position rings are display. Pipe which delivers the pressure, micro valves and external pipe are not displayed.

Solid waste delivery requires water input into the system in order to reduce static friction along the surface of the inner tubes. The water supply for waste management is isolated from all other systems.

An additional development cost needs to be taken into account in order to test the integrity of the ionomer membrane within our proposed system. However, the design strengths fulfil criteria regarding hygeine and storage, whilst additionally contributing to radiation shielding.

3.3 Application of Melanin

Melanin is an organic molecule capable of absorbing radiation (UV to gamma-rays). Dr Dadachova at the Albert Einstein College of Medicine, University of Yeshiva, in New York, demonstrated that it is possible to use Melanin for radiation shielding [18]. The outermost tube will be coated with a layer approximately 3.5mm thick. This will provide additional protection to the crew. This total 7mm thickness is intended to reduce penetration by inci- dent gamma rays by up to 50%.

The Linear Attenuation Coecient of Melanin µ, which quantifies the quality of a radiation shield, is 1.01 cm1. From this, the Half Value Layer thickness is given by the following equation:

1 HLV = ln2=0.68cm (1) µ

12 Inspiration Mars Student Design Competition Section 3

Name µ (cm-1) HLV (mm) Density (g. cm-3) Weight (g. cm-2) Melanin 1.01 7 1.33 0.931 Lead 27.1 2.5 11.34 2.835 Aluminum 0.384 18 2.7 4.87 Table 3: A comparisson of HVL values between various materials.

3.4 SPE Events

Solar Proton Events produce extremely high radiation levels, and can result in instant radi- ation induced death. The mission will take place during a solar minimum (year 2018), and hence very few SPEs are expected. However, precautions should be taken to prevent poten- tially catastrophic results. In the case of an SPE, the crew would be alerted by the ground station, who would continue to provide observations and analyses. This would provide the crew with approximately a day to prepare the spacecraft’s emergency orientation (see 3.5) and begin preparations of the SPE shelter (see 3.6).

3.5 Magnetic Field Probe

A magnetic field probe has been designed to be deployed on a telescopic boom of approxi- mately 8 m during the beginning of the interplanetary travel phases to monitor the environ- ment’s interplanetary magnetic field (IMF). This would cheaply provide detailed information of the magnetic field lines and strength. In the event of an SPE, the spacecraft would then be able to properly align the SPE shelter (see Section 4.7) to counter the incident flux. Data col- lected throughout the mission would also provide useful scientific data to potentially develop or contribute to existing IMF models.

3.6 SPE Shelter

Small modifications are required to convert the cylindrical core into a SPE shelter. The procedure in the event of an SPE is as follows: the onboard water used for radiation shielding is to be evacuated from the bubbles to fill an enclosed segment of the core (see section 3.1).

13 Inspiration Mars Student Design Competition Section 4

Figure 10: Flushing water from the bubble to the core to provide a water shielding thickness of approximately 3 m in length oriented in the direction of the incident radiation. The crew are kept inside a hatch which itself adds material protection in all directions. A full design for this hatch would need to be optimised in a way to minimise radiation dose.

4 Internal Housing and ECLSS

The housing structure has been designed to keep the inhabitants continuously protected from GCRs. As such, the overall housing which is encased within the Bigelow Sundancer module is centered around the two radiation shielding components described in sections 3.1 and 3.2. This section describes the general configuration of the internal housing system and discusses how it is suited for long-duration deep space flight.

4.1 Bubble Configuration

The configurations of the bubble structures are dependent on the requirements of the inhab- itants. The bubbles provide what are essentially ”rooms” and each can be customised to suit the room requirements, for example the sleeping space requirements are di↵erent from the exercise space requirements. The total dimensions of the bubbles must fit within the Sundancer with some margin allowed for other sub-systems, such as the waste management system. The final configuration (see Figure 11) consists of three sets of rooms: two large sets and one slightly smaller. The smaller rooms are allocated to the the sleeping quarters and lounge or leisure areas as they do not require as much space and benefit from thicker shielding. Each room pair is separated by a hatch that allows access to the next room as water is flushed between them.

14 Inspiration Mars Student Design Competition Section 4

Figure 11: Illustrates the configuration of the bubble housing structure together with the specification for the two di↵erent types of bubbles .

Rooms The following room allocation is provided to accommodate the various activities of the astronauts during the trip to Mars. The largest choice for the bubble dimensions are 2.5 m. The design also makes use of a smaller bubble set at just 2 m long. This is to be used as a dormitory which does not require as much volume while allowing thicker shielding. With the chosen bubble sizing and accommodation requirements, the next step is to distribute the space for all the activities that will be conducted during the mission. The main areas include:

1. Hygiene and WC: This includes a shower approximately 2x1.5x1 m. Toilet dimensions are 0.5x0.5x1 m, with extra space to provide integration with the water and waste management system.

2. Medical and Exercising: It is important that astronauts do enough exercise to main- tain acceptable bone mass. In addition to exercise equipment, medical equipment is provided on board, not only for possible contingencies but also to take regular tests and provide health monitoring.

3. Kitchen and Dining: This area are used for food preparation and dining. It would also provide space for food storage, e↵ectively acting as a pantry(see section 4.2.2).

4. Science: Carrying out scientific experiments remains a goals of the mission. A dedicated area has been assigned to accommodate computer systems and a small space laboratory for technology development and scientific research (see section 6).

5. Sleeping quarters: Contains the beds and most of the personal belongings.

6. Lounge/Communications: Provides an area of comfort to the astronauts and a place to relax, where entertainment systems could be included. Additionally, this room acts as

15 Inspiration Mars Student Design Competition Section 4

a communications post from which astronauts would transmit back to ground stations on Earth. This post can be accessed remotely from all parts of the spacecraft.

4.2 Food Management

4.2.1 Food Requirements

For a long duration mission to Mars, managing the astronauts food is essential to their survival and the success of the mission. It is important to consider the nutritional content, the variety of food and ease of preparation in zero gravity. In space, the average persons intake of solid foods is 0.62 kg per day; hence, a minimum of 621 kg of solid food must be taken with the two astronauts [19]. The total mass estimate of the entire food management system, including 650 kg of dehydrated food, packaging, and preparation equipment, is 1458 kg [20]. This food will consist of a large variety of foods to maintain morale throughout the mission. It must also ensure that the astronauts get the required nutrients, minerals, and other compounds to keep them strong and healthy. The nutritional value required daily for a male astronaut of 98.5 kg and a female astronaut of 61.7 kg is shown in Table 4 [21].

The body masses of the astronauts were assumed using the 95th percentile male and fe- male from [21], from which the daily intake values were calculated assuming a requirement of nutrients and energy on a per body mass basis. The 95th percentile mass values were used to ensure that sucient food was on board to account for any delays in the mission or di↵erences in astronaut metabolism. Using these values, large bags of dehydrated food can be arranged to consist of the weekly amount of food required for the astronauts. The food will be carefully planned to provide the required weekly minerals and energy.

Daily Male Req. Daily Female Req. Total Weekly Req. Energy (kJ) 12950 9600 22550 Protein (g) 78.8 49.36 897.12 Fat (g) 128.05 80.21 1457.82 Carbohydrates (g) 472.8 296.16 5382.72 Phosphorus (g) 0.8 0.8 11.2 Sodium (g) 3.5 3.5 49 Iron (g) 0.018 0.018 0.252 Calcium (g) 0.8 0.8 11.2 Magnesium (g) 0.35 0.35 4.9 Potassium (g) 2.7 2.7 37.8 Table 4: The recommended daily nutrient intake for a male and female astronaut [19], as well as the total weekly nutrient intake of both astronauts.

16 Inspiration Mars Student Design Competition Section 4

4.2.2 Food Storage

Food storage was a closely assessed issue as food is one of the few resources that cannot be recycled. The total dehydrated food volume for the mission’s 501 days, plus contingencies, is calculated to be 14 m3.This14m3 is approximately the volume provided by a single small bubble structure and would take up a considerable amount of useful space. Instead, approx- imately two thirds of the food is kept in the Dragon capsule with the remaining third kept in the kitchen bubble. This would mean access to the Dragon capsule would only be required two or three times throughout the mission’s interplanetary phase and poses minimal risk of damaging the capsule. An analysis of the food volume for 8 weeks worth of dehydrated food showed that less than 35% of the kitchen bubble would be filled while requiring only 2 added entries into the dragon. Despite the consumed space of the bubble, Figure 12 illustrates that there is sucient space within the module to still cook and eat.

The option of storing food in the surrounding volume of the bubble structure to be used as an additional layer of shielding has been discarded. This would require the astronauts to rely on mechanical procedures in order to reach the food. Should something occur to the retrieval mechanisms, exposure to harmful radiation levels would occur as the astronauts would be required to exit the internal housing structure.

Figure 12: Kitchen Illustration. Despite the food storage taking considerable space within the mod- ule, this illustration shows that there is sucient space to accomodate both individuals at the same time.

4.3 Waste Management Configuration

The waste management system was outline in Section 3. The tube system is shipped within the Sundancers structure. Post inflation, the tubes must be position manually on a face of the Sundancer. This would eciently be handled by providing a layer of velcro along the wall and simply strapping in the system of tubes in order to provide a completely covered walled

17 Inspiration Mars Student Design Competition Section 5 that would provide additional protection for CGRs (shielding from one direction only) and SPEs (see Figure 10 - left).

5 Communications

5.1 Traditional Communications System

The spacecraft is equipped with standard X and Ka Band transceivers for TT&C and low bandwidth video respectively. A preliminary link budget is shown below, with the prop- agation path length taken to be 1.5AU, approximately the maximum distance from the spacecraft to Earth. Low gain antennae will also be included for redundancy. Note that if deemed necessary, the required data rate could be dropped to allow for lower transmitter power or antenna size. A link budget for the laser system can be found in [22].

XBandLinkFrequency(GHz) 8 Ka Band Link Frequency (GHz) 30 Transmitter Power (W) 550 Transmitter Power (W) 300 Transmitter Power (dBW) 27.40 Transmitter Power (dBW) 24.77 Transmitter Gain (dB) 42.71 Transmitter Gain (dB) 54.20 Spacecraft Antenna Diameter (m) 2.2 Spacecraft Antenna Diameter (m) 2.2 Propagation Path Length (km) 2.24 108 Propagation Path Length (km) 2.24 108 ⇥ ⇥ Free Space Loss (dB) 2.78 102 Free Space Loss (dB) 2.89 102 ⇥ ⇥ System Losses (dB) -5 System Losses (dB) -5 Receiver Gain (dB) 72.77 Receiver Gain (dB) 84.25 System Noise Temperature (K) 200 System Noise Temperature (K) 200 Minimum Data Rate (Mbps) 1 Minimum Data Rate (Mbps) 7.5 Eb/N0 (dB) 5.95 Eb/N0 (dB) 6.05 Required Eb/N0 (dB) 3 Required Eb/N0 (dB) 3 Link Margin (dB) 2.95 Link Margin (dB) 3.05 Table 5: Communications budget for tradi- Table 6: Equipment Power Consumption tional communications systems.

Constraints of Traditional Communications System There are a number of con- straints in using conventional RF antennas for communication over such long distances. RF systems su↵er from increasing propagation losses with increasing separation between the transmitter and receiver. While RF systems can benefit from high gain antennas, the power requirements are still considerable, especially when transmitting across the distances con- cerned here. With conventional systems, high data rates for audio and video transmission (e.g. video feed of Mars) is extremely dicult. However, free-space optical communications using infra-red lasers have been identified by NASA as a technology that would provide significantly higher performance data return rates [23]. A proposed laser communication system is provided in Section 5.2.

18 Inspiration Mars Student Design Competition Section 5

5.2 Remote Communications System

It can be dicult to justify the costs, financial and otherwise, associated with such a high profile mission when the only objectives are to perform a flyby of Mars and return to Earth. Without spending any significant period of time in Martian orbit, the scientific opportunities o↵ered by such a mission are restricted to the study of human health and biology in space. Although this is certainly a worthwhile area of research, the value of the mission could be significantly increased by supplementing it with a smaller mission that would take advantage of the unique opportunities on o↵er. It would be additionally beneficial if this secondary mission could be carried out in conjunction with other organisations - this might allow both the costs and benefits of such a mission to be shared. One such organisation is Mars One [24] who have announced plans to place a telecommunications satellite in areostationary orbit by 2018 [25], therefore, a successful mission would certainly be of interest to them.

In recent years, a number of missions with the aim of demonstrating free-space laser com- munications have been proposed by NASA - most notably the Lunar Laser Communications Demonstration (LLCD), which flew on board the LADEE spacecraft in 2013 and successfully demonstrated a downlink data rate of 622 Mbps from lunar orbit [26]. In 2009 the design of a spacecraft that would demonstrate laser-based communications between Earth and Mars [25] was cancelled due to budget constraints. Its payload, known as the Mars Laser Communi- cation Demonstration system (MLCD) was expected to achieve a downlink data rate of up to 30 Mbps [27]. Subsequently, a modified design - the Deep space Optical Terminal (DOT) was designed to include the laser communications system along with a stabilisation platform and was designed to be adaptable and fit onto generic Mars orbit vehicles [22]. While the design has been shown to be feasible, there has not yet been a mission proposed that would make use of the system.

Laser Communications Satellite The design proposes building a small satellite incor- porating the DOT system that would be deployed in space to fly alongside the primary spacecraft operating as a high performance data relay between the spacecraft and Earth when conditions are favourable. This tandem flight was selected due to the high pointing re- quirements incurred by the DOT system, which would be extremely dicult to achieve with the main spacecraft. During the mission’s flyby stage, the satellite will manoeuvre itself into a 500 km altitude orbit about Mars to carry out a Mars observational mission while contin- uously demonstrating DOT technology. Additionally, the spacecraft would also be equipped with a standard Electra telecommunications package [28], allowing it to function as a data relay for existing and future Mars orbiters and landers. Section 5.3 presents a baseline design

19 Inspiration Mars Student Design Competition Section 5 for this secondary spacecraft, detailing the main requirements and constraints that influence the spacecraft design. This is followed by a brief overview of the di↵erent subsystems and mission phases. The technical details of the laser communication payload are provided by Biswas [22]. The overall mass budget is included in Section 10.6.

5.3 Communication Satellite Design

5.3.1 Design Drivers

Pointing Accuracy Constraint The primary constraint of using a laser system with atransmitted-receiverseparationofthisscaleisthenecessaryAttitudeandOrbitControl Systems (AOCS) to keep the spacecraft pointing at the Earth. A laser communications link between Mars and Earth would require transmitter pointing accuracies on the sub- microradian level [29]. This combined with an additional need for a milliradian level point- ahead angle to compensate for the relative movement of Mars and Earth, as well as the diculty of implementing closed loop control due to the round-trip time delay places a demand for extremely high pointing accuracy and stability on the entire system. While the DOT system includes a stabilisation platform that shoulders much of this responsibility, the spacecraft must still be able to point the laser terminal towards Earth with around 3 mrad accuracy [22].

Orbit and Lifetime The mission lifetime chosen is 2.5 years, which should ensure a high volume of scientific data return and justify development costs. The selected circular orbit has an altitude of 500 km with an arbitrary inclination. While areostationary orbits are preferred for communication satellites around Mars, it was decided that an orbit that saw more of the planets surface would provide more useful data return, particularly from the cameras on- board. 500 km was chosen because it was low enough to allow for high-resolution imaging from the cameras without requiring large amounts of propellant to maintain altitude.

5.3.2 Design

As with the main mission, cost and technology readiness level are key factors. For this reason, existing technology has been used throughout the design. As mentioned earlier, the laser system is based on the design of the DOT, which consists almost entirely of components already used in terrestrial optical communications (see [22] and [30] for information on the system and stabilisation platform respectively). The rest of the spacecraft design uses many of the components previously used on the Mars and Lunar Reconnaissance Orbiters, since they are recently built spacecraft of a similar mass and performance level to what is required

20 Inspiration Mars Student Design Competition Section 5 of this system. The spacecraft is to be stored in the unpressurised trunk of the Dragon capsule before deployment.

5.3.3 Added Payloads

As mentioned previously, the payload consists of the DOT system, along with two narrow angle cameras and a wide angle camera based on those of the Lunar Reconnaissance Orbiter [31], and the Electra package [29]. Although the DOT design places the laser terminal mounted to the body of the spacecraft, the terminal can also be mounted on a two- axis gimbal for improved accuracy. This explains the slightly higher mass and power estimates in the budgets below, compared to those given in the DOT design.

Figure 13: Schematic of Laser Flight Terminal from DOT Design. [32]

Cameras The two narrow-angle cameras on the LRO have been able to capture extremely high quality (sub-metre resolution [31]) images of the Moon, and the high pointing accuracy of this spacecraft will allow for unprecedented imaging opportunities of the surface of Mars. The cameras will likely require some additional development in order to be used for this mission, which has been accounted for in the cost analysis. The cameras are mounted to the exterior of the satellite, pointing in the nadir direction. A wide-angle camera similar to the one carried on the LRO mission is also to be adapted for this mission at little extra cost. This is mounted in the same way as the narrow angle cameras.

Electra The Electra Proximity Link Payload has flight heritage, having already been car- ried on the Mars Reconnaissance Orbiter and MAVEN spacecraft to date. Electra allows for UHF communications links with data rates of up to 1 Mbps to be established between spacecraft around Mars, enabling data relay and ranging capabilities throughout the network. A detailed description of the package can be found in [28] The main physical components

21 Inspiration Mars Student Design Competition Section 5 of the system are a transceiver unit with dimensions of 17.2 x 21.9 x 14 cm and a small nadir-pointing UHF antenna.

5.3.4 Structure

As with the MRO, the structure will consist mainly of carbon composite and aluminium honeycombed plates. The structural mass was estimated based on the total spacecraft mass using techniques given in [33].

5.3.5 Power

This spacecraft will use Gallium Arsenide solar cells and Lithium-ion batteries. A preliminary power analysis has been provided in Figure 7 and it has been determined that the maximum solar array area required is 13 m2. This value is extremely large, but this is primarily because it has been assumed that all equipment is constantly running at full power- that is, duty cycling has not been taken into account. Another reason for this high value is the value for solar flux in the calculation is that experienced when Mars is at aphelion. The power subsystem requirements could be dramatically lowered by reducing usage of high power consumption equipment at such distances.”

Maximum night-time load (W) 662.5 Maximum Battery Discharge (Wh) 477.28 Acceptable Depth of Discharge (%) 40 Required Capacity (Wh) 1193.19 Power needed to charge batteries (W) 501.06 Maximum daytime power requirement (W) 736.1111 Power needed from solar arrays (W) 1237.18 Solar flux at maximum Mars-Sun distance (W/m2) 470 End-of-Life solar array eciency (%) 25.76 Required solar array area (m2) 13.48 Battery mass (kg) 9.55 Solar cell mass (kg) 37.06 Table 7: Power Sub-system Sizing

AOCS This subsystem was based on that of the MRO, which is of a similar size and can achieve the pointing accuracy and stability required for this spacecraft. It uses 4 Honeywell Constellation Series HR16 reaction wheels for attitude control, along with thrusters for un- loading. Attitude determination is performed by the Galileo Avionica A-STR Star Tracker, along with SST US 2-axis Sun Sensors and a Honeywell Miniature Inertial Measurement Unit for inertial navigation. If this combination is deemed to be inadequate for the precise attitude knowledge required, the SSBV Earth Horizon Sensor has also been identified as an additional attitude determination sensor. Extra units of each of these sensors will be carried for redundancy. Technical specifications for each of these components can be found

22 Inspiration Mars Student Design Competition Section 6 in [34] [35]. For a more detailed analysis of the AOCS and propulsion systems, please see www.cranspace.com.

Ground Stations Although there are no currently operational ground stations that are capable of supporting a data link for this mission, it is extremely unlikely that a ground station would have to be constructed from scratch. The LLCD uses a ground terminal which includes a transmitter, consisting of an array of four 15 cm telescopes, and a receiver, which consists of four 40 cm telescopes as pictured below. Although it is based primarily in White Sands, New Mexico, the terminal is designed to be transportable. Although this terminal does not have the capability of supporting a link from Mars (the DOT report specifies aminimumreceiverapertureareaof3.8m2 [22], building an upgraded version with the required capabilities would not be likely to require huge research and development costs. If this secondary mission is supported by another organisation such as NASA or Mars One, the cost of doing so would not fall entirely on the Inspiration Mars Foundation. However, due to the extreme sensitivity of a laser communications link to factors such as cloud cover and other atmospheric conditions, it is likely that the system would require multiple ground stations in carefully chosen locations around the world, to ensure that at least one ground station is reachable when links are to be established. The DOT report explored the possibility of retrofitting existing ground-based astronomical telescopes to operate as ground stations [22] [32].

Figure 14: The Lunar Laser Figure 15: Concept of Operations for the DOT [32] Ground Terminal [22]

23 Inspiration Mars Student Design Competition Section 6

6 Science

The variety of space environments encountered, as well as the numerous challenges faced by the mission and its crew, will provide a unique opportunity for scientific research. As no spacecraft of this size and mass has ever been sent to Mars, the Inspiration Mars mission has the chance to carry a multitude of science payloads.

6.1 Extended Boom and General Sensors

Anumberofmagnetometersandspectrometersexternallymountedonthespacecraftwill provide valuable science data for the mission. It was stated in Section 3.5 that a magnetic field probe would be placed at the end of a telescopic boom approximately 8 m long when fully extended. A record of the magnetic field variations in deep space may be of scientific interest. Additional sensors may also be included to perform brief experiments when passing through the Van Allen belts.

6.2 Analysis of the e↵ectiveness of Melanin Nano-shells

The Melanin on board the spacecraft (coating the waste management system) will provide the crew with high levels of electromagnetic radiation protection. A sensor to determine the X-ray and gamma-ray flux will be located outside of the spacecraft in order to allow for comparison with values measured within the bubble structures. A successful demonstra- tion of the technology could considerably help the development of highly radiation resistant structures.

6.3 The E↵ects of Deep Space Radiation on the Growth of Plants

For future long term space exploration missions, it is envisioned that plants grown in space could play a major role in the biological life support system (BLSS) providing both clean air and food [36]. Furthermore, the presence of plants has been noted to have a positive e↵ect on the well-being of crew members [37]. However, all such experiments have been performed in LEO under significantly di↵erent radiation conditions than those experienced on a deep space mission. The Inspiration Mars mission is therefore an excellent opportunity to examine the growth of plants and to assess whether plants grown in such conditions will be suitable for human consumption.

[36] identifies the selection of the perfect plant material for BLSS by considering the rate of oxygen and biomass production and the radiation resistance of various plant species as an

24 Inspiration Mars Student Design Competition Section 6 important step in the development of bio-regenerative life support systems. This is therefore an important theme that can be studied on board. There are several types of equipment currently being used on the ISS for the growth of plants in space, therefore no further de- velopment costs will be incurred.

6.4 Crew Activity

The monitoring of crew activity on board can be achieved using widely used devices such as the Fitbit [38] and Google Glass [39]). These wearable electronics, when integrated with specifically developed applications could provide benefits such as Augmented-Reality capabilities, remote voice control, imaging, audio and video correspondence with Earth and entertainment. The development of applications for space for Google Glass hardware is not expected to require considerable investment. The Fitbit could provide physical activity monitoring and it could be of scientific interest to observe how this varies throughout the duration of the mission.

6.4.1 Psychological evaluation

On board cameras with facial recognition software [40] to monitor the mental state of the crew could also be utilised. This software evaluates the astronauts concentration, which aids in avoiding potential human errors.

6.5 Laser-based Communication System Demonstration and Mars Reconaissance

6.5.1 Technology Demonstration

In Section 5.3, an outline for the small laser-based communication relay satellite was pro- vided. The satellite will demonstrate laser-based communications for deep space applications. While a number of missions making use of optical communications have been proposed, the application of the technology is not yet widespread. By demonstrating the technology over distances far greater than previously tested, broad implementation in the space industry could be achieved.

6.5.2 Scientific Observations

In addition to the demonstration, the satellite can be adapted to give additional benefits at little extra cost. Outfitting it with cameras similar to those used on the LRO [31] could

25 Inspiration Mars Student Design Competition Section 7 allow for extremely high quality images to be returned from Mars. As was done by the Reconnaissance Satellites, stereo images will be returned and will allow the calculation of topological data [41].

6.5.3 Earth Re-entry

With this being the first time a manned mission has attempted re-entry to Earth at such a velocity, data on temperature and pressure distributions experienced by the capsule during re-entry could provide valuable information that will aid future re-entry missions. This data can be collected in a similar way as for the Mars Science Laboratory, by using atmospheric data sensors such as MELDI and MEADS integrated into the re-entry capsule heat shield [42].

7 Launch Vehicles

7.1 Launcher Comparison

Due to the substantial payload mass, the evaluation of launch vehicles was confined to only the heaviest launchers. As Inspiration Mars has been put forward as an American mission, all Russian and Chinese launchers were excluded as possibilities. Table 8 provides a list of the considered launch vehicles, as well as their payloads to LEO and launch cost. The Ariane 5, Delta IV and Atlas V all have high per-mass costs to deliver a payload into orbit, when compared with either the Falcon Heavy or the Falcon 9 launchers. While the Space Launcher System (SLS) Block 2 would allow for the largest payload to be lifted to LEO, its price per kilogram in orbit remains higher than either of the Falcon launchers. Additionally, the SLS is only expected to perform its first test flights in 2017 and is therefore unlikely to be ready for use in time by 2018. In comparison, the Falcon Heavy launcher should perform its first flight later in 2014 [43]. The Falcon Heavy launcher provides the most cost-ecient means of access to LEO, as well as a high payload capacity of 53 metric tonnes. It can also be fitted with SpaceX’s own Dragon capsule, which is to be modified and demonstrated for human occupation.

Launcher Ariane 5 Ariane 5 ME Atlas V 551 Altas V 541 Delta IV-H SLS Block II Falcon 9 v1.1 Falcon Heavy Payload to LEO (t) 21 25.2 18.5 17.1 22.95 105 13.15 53 Cost of launch (M$) 120 150 190 223 380 750 56.5 135 Cost per kg (k$/kg) 5.714 5.952 10.270 13.041 16.558 7.143 4.297 2.547 Table 8: Total payload to LEO and associated costs for various launch vehicles.

26 Inspiration Mars Student Design Competition Section 7

7.2 Maximum Number of Launches and LEO Launch Mass

Due to schedule, planning and in-orbit docking procedures, as well as cost considerations, the number of launches available for the mission was limited to an absolute maximum of four launches. It was decided that the final launch sequence would consist of four Falcon Heavy launches. This would allow a maximum payload of 172 metric tonnes in a 200km altitude circular orbit. By extrapolating the loss of payload for higher orbits using SpaceXs Falcon 9 user manual [44], Table 9 was produced showing the total payload to 700 km per Falcon Heavy would be reduced from 53 metric tonnes to 47.2 tonnes. This corresponds to a total mass in orbit of 188.8 tonnes using four Falcon Heavy launches. This total would have to include the dry mass of the craft, the propellant needed in order to inject it into its orbit, as well as any additional fuel tanks.

Circular Orbital Altitude 200 km 300 km 400 km 500 km 600 km 700 km 800 km 900 km 1000 km Falcon 9, v1.0 10450 10250 9950 9750 9550 9350 9250 9150 8950 Falcon 9, v1.1 Payload (kg) 13000 12751 12378 12129 11880 11632 11507 11383 11134 Falcon Heavy 53000 51986 50464 49450 48435 47421 46914 46407 45392 Table 9: Table shows the payload mass limit for a range of altitudes for the Falcon 9 and Heavy launch vehicles.

7.3 In-Orbit Propulsion

7.3.1 Delta-V Budget

Earth Escape The velocities at infinity required for the desired trajectory to Mars were calculated and detailed in the Inspiration Mars feasibility study [45], and can be found in Table 10.

Departure V Infinity (km/s) Arrival V Infinity (km/s) Mars-bound 6.232 5.417 Earth-bound 5.417 8.837 Table 10: Velocity at infinity of Earth’s departure and the return trip’s arrival.

These velocities correspond to an orbit optimised for free return. Using the velocity at infinity upon departure from Earth, the following equation was used in order to calculate the velocity required at the perigee of the escape hyperbola.

2 v µE µE v = vp vc = v2 1 + u 2 rp ! s rp u t where µE is the gravitational parameter of Earth, vp is the velocity at the perigee of the hyperbolic trajectory, vc is the velocity in the circular parking orbit and rp is the radius of

27 Inspiration Mars Student Design Competition Section 7 the circular parking orbit.

As can be seen from the equation, the necessary velocity is dependent on the radius of the parking orbit. This design implements a circular parking orbit at an altitude of 800 km in order to avoid manoeuvre complications with LEO orbiting spacecraft (ISS) while ensuring the altitude is great enough to minimise atmospheric drag. This leads to a required delta-v of 4800 m/s.

Gravity Loss Gravity losses are also dependent on the initial acceleration provided by the engine and have been assumed to be approximately 200 m/s as a contingency, hence bringing the total delta-v to provided 5000 m/s. The validity of this contingency is discussed and approved below, using the selected orbital engine.

7.3.2 In-Space Propulsion System

Merlin 1D Engine The simplest launch option involves using the Falcon upper stage engine, currently the Merlin 1D Kerosene/LOx engine. This would mean increasing the size of the propellant tanks carried by the upper stage, or carrying separate tanks as payload. This would be advantageous, as the only cost would be that of the new propellant tanks, but would not represent a large investment in terms of development costs, nor the purchase of other expensive engines. However, with an Isp quoted at 342s [43], the Merlin 1D would require a large quantity of fuel in order to reach the required delta-v, as shown in Table 11 below. Other chemical propulsion systems were also considered. While the Merlin 1D engine uses a combination of kerosene and liquid oxygen as propellants (responsible for the low Isp), LOx/LH2 engines, like the RL-10B-2 [46] engine employed by the Delta IV launchers upper stage, as well as the Space Shuttle Main Engine generally produce higher Isps. However, they come with the disadvantage of requiring heavier and more complex systems for propellant storage. Furthermore, these engines are more complex and more costly than the Merlin 1D. It is clear that, in already using SpaceX’s Falcon Heavy launch vehicle, there is the distinct cost-benefit and simplicity in using SpaceX’s Merlin 1D engine.

Raptor Engine SpaceX is currently developing a new orbital engine - the Raptor engine, fuelled with a mixture of methane and liquid oxygen. These new engines should equip SpaceXs vehicles in the future, but as of now remain in development. These new engines are included in Table 11. As there is only limited data available for methane-fuelled engines, two values of specific impulse were employed in this table: the pessimistic Isp value of 350s

28 Inspiration Mars Student Design Competition Section 8 represents a small increase from the Merlin 1D engine, whereas the optimistic value of 380s is reportedly the target that SpaceX aims to achieve.

Burn 1 Burn 2 Engine Isp (s) Init. Mass (t) Delta-V (m/s) F. Mass (t) Init. Mass (t) Delta-V (m/s) F. mass (t) Merlin 1D 342 196.7 825 154 147.5 1152 105 Raptor 350 196.7 862 153 147.5 1206 104 Raptor 380 196.7 936 153 147.5 1309 104 Burn 3 Burn 4 Total S/C Engine Isp (s) Init. Mass (t) Delta-V (m/s) F. Mass (t) Init. Mass (t) Delta-V (m/s) F. Mass (t) Dry Mass (t) Merlin 1D 342 98.3 1340 66 59.7 1682 36 31.8 Raptor (p.) 350 98.3 1413 65 59.7 1519 38 36.6 Raptor (o.) 380 98.3 1534 65 59.7 1221 43 41.7 Table 11: Comparison between the Merlin 1D values and the Raptor estimated values.

As the future of the Raptor is yet very uncertain, it was decided that the Merlin 1D engine would be used in this design. If the Raptor engine were to become available by the onset of the mission, the potential mass gains shown in Table 11 could be applied as an enhancement to this design with the additional fuel savings potential being used to reduce reentry velocity or to reduce the number of total required launches.

8Reentry

The purpose of the thermal protection system (TPS) is to slow down the spacecraft, thus providing a human tolerant temperature inside the spacecraft and avoiding high acceleration loads. This is achieved by using drag to counteract excess speed. However, the interaction between the air and spacecraft produces a large amount of heat, which the material must be capable of withstanding. The main requirements driving the design of the heat shield mate- rial are its maximum load factor, the maximum heat flux that the material can withstand, the total mass of the TPS and re-usability if aerocapture is the method used.

8.1 Reentry Vehicle Configurations

Develop a new vehicle is very expensive and time consuming. Due to the available time, a available spacecraft will be selected. It will focus on the TPS subsystem on each vehicle.

NASA proposal NASA’s Orion MPCV (Multi-Purpose Crew Vehicle) is being designed for travelling beyond low Earth orbit (LEO). However, the capsule was not considered for this mission due to scheduling uncertainties- it is unlikely that the capsule would be ready for use by 2018. The extra mass of the capsule, compared to that of the Dragon, is another reason why the Dragon is preferred.

29 Inspiration Mars Student Design Competition Section 8

SpaceX Proposal SpaceX’s Dragon capsule is designed to deliver both cargo and eventu- ally humans to destinations in LEO. It is the only spacecraft currently in production that is capable of returning significant amounts of cargo to Earth. Under an agreement with NASA, SpaceX is currently adapting the capsule to enable it to transport humans [47].

8.2 Reentry Trajectory

The main problem encountered when designing a spacecraft to re-enter from Mars is the high approach velocity on returning to Earth, which is around 14km/s. To date, no manned spacecraft has achieved re-entry at such a velocity- the highest that has been achieved is by the Apollo missions, which at 11km/s [48]. The highest successful re-entry velocity for an unmanned spacecraft was experienced by the Stardust probe, which returned at 12km/s. The maximum acceleration load that is permissible for astronauts returning from long du- ration missions to experience has been limited by NASA to 5gs.

The re-entry path angle can be altered to limit these loads- as this angle increases, the acceleration also increases. In order to be captured into Earth orbit, the speed must be re- duced to around 11km/s. The lifting body configuration of the Apollo or the Soyuz capsules is also used to to reduce acceleration in the case of a purely ballistic re- entry. If the path angle is too low, the spacecraft is unable to re-enter and returns to space. This e↵ect can be used to reduce the acceleration load and heat flux experienced, although it comes with the cost of increasing the mission duration by a few days. After one missed re-entry attempt, the capsule can reduce its speed to a value low enough to allow for re-entry. In this case, the heat shield must be designed so that it can be used for multiple entries into the atmosphere. The aerocapture method also imposes restrictions on other elements of the mission design- since the service module and inflatable habitat must be jettisoned before the first re-entry, the subsystems on board the Dragon capsule must be capable of supporting the astronauts throughout the duration of the re-entry period. The electrical power subsystem is of partic- ular importance- see section 7.1.2 for a sizing analysis of the hydrogen fuel cells.

The model used for the Apollo missions [48] has been implemented using the properties of the Dragon capsule. Fig. 16 shows the trajectory for the first re-entry from Mars. Once the altitude is high enough the new orbit is calculated. A second re-entry is used to decrease the speed to 8.1 km/s resulting in a 1.5 hour orbit. The third and final orbit eliminates the remaining speed in order to return the astronauts safely to Earth.

30 Inspiration Mars Student Design Competition Section 8

5 Reentry Acceleration x 10 Reentry altitude Reentry Speed 14000 5 3

4.5 12000 2.5 4 10000 3.5 2

3 8000

2.5 1.5 Altitude, m

Velocity, m/s 6000

Acceleration, g 2 1 1.5 4000

1 0.5 2000 0.5

0 0 0 50 100 150 200 250 300 0 50 100 150 200 250 0 50 100 150 200 250 time, s time, s time, s

Figure 16: First re-entry. The aerocapture reduces the speed from the original 14.2 km/ to a suitable 10.95 km/s that has 3.5 days as orbit period. The initial fligh path is 6.1 deg

4 Reentry acceleration x 10 Reentry altitude Reentry velocity 12 2.5 8000 11 7000 2 10 6000 9

1.5 5000 8

4000

Altitude, m 7 Velocity, m/s

Acceleration, g 1 3000 6

2000 0.5 5

4 1000

0 100 200 300 400 500 600 700 0 100 200 300 400 500 600 700 0 100 200 300 400 500 600 time, s time, s time, s

Figure 17: Simulations of the last re-entry.

8.3 Heat Flux

This is the key parameter in the TPS selection. Fig. 19 shows the maximum heat flux as afunctionofthere-entryspeed[49].Inthiscase,theheatfluxisexpectedtobearound 3300 W/cm2. This graph also shows results for two di↵erent mass loadings. Fig. 18 shows this heat flux against time for each trajectory. The peak heat flux is experienced at the beginning of the trajectory.

Figure 18: Heat flux throughout the reentry. Figure 19: Heat Flux for di↵erent configura- For the expected reentry at 14 km/s the max- tions imum heat flux is around 3300 W/cm2

31 Inspiration Mars Student Design Competition Section 8

8.4 Heat Load

The heat load defines the required thickness of the TPS. For example, the Dragon capsule has 3 inches of PICA-X and the Apollo mission had 7 inches of Avco 5026-39G [50] With lifting bodies the heat flux peak reduces but the heat load increases. For the Mars re-entry this parameter is not the key parameter to select in the TPS selection.

Figure 20: Heat load for several configuration, from [49].

8.5 TPS material

Considering all possible materials on the market, there are few possibilities to use for the TPS:

AVRCOAT Used in the Apollo missions [50]. It was the standard TPS material until the development of PICA.

PICA Developed by NASA. It has been used in the Stardust probe and the MSL.

PICA-X Recently, Space-X has developed a new material called PICA-X based on the original PICA. It is ten times more cost e↵ect and more reusable [51]. They have tested it and it has been shown to withstand temperatures of up to 1850 degrees Celsius. However, it cannot withstand the 3500 W/cm2 that are expected [49]. According to the CEO of Space- X, Elon Musk, the spacecraft can re-enter from Mars [52] but no further references has been found.

High Density Carbonic Phenolic A material developed by NASA. The document [49] shows that it is valid for the conditions of the re-entry from Mars. It has been used in the Phobos Grunt [53]. This mission was designed to travel to one of the Mars moons, Phobos, and return with samples. The spacecraft had problems in low Earth orbit and it re-entered

32 Inspiration Mars Student Design Competition Section 9 without leaving the Earth. The material has also been proposed for the Mars Sample Return mission. [54]

Hence, only the High Density Carbonic Phenolic can perform a re-entry at 14 km/s, so this material will be the one used for the TPS. For testing the material, new facilities will be necessary. This material will be adapted to be used on the Dragon body instead of PICA-X.

9 Electrical Power Systems

This design only considers traditional electricity generation methods and storage systems - solar arrays and batteries. Bigelow’s Sundancer module and SpaceX’s Dragon capsule are expected to be equipped with solar arrays large enough to sustain full functionality of their original equipment systems at a distance of 1 AU from the Sun, but not at the 1.39 AU distance of Mars.

General Assumptions The Dragon will be equipped with solar arrays and batteries as standard. A regenerative hydrogen fuel cell will be added inside the crew compartment. The Sundancer will be modified considerably to accommodate the radiation shielding systems. Both of these are expected to have significant additional power requirements. Additional solar panels and batteries will therefore be required.

9.1 Dragon

9.1.1 Power Generation

During the interplanetary phase of the mission the Dragon will not be used as a primary habitat and therefore will require minimal power. Even at the distance of Mars, this will be provided by the Dragon’s own solar arrays. SpaceX states the mean power provided by the solar arrays in LEO is 2kW [55]. At the Mars flyby this will be reduced but will remain greater than 1kW [47].

9.1.2 Energy Storage

The batteries of the Dragon are capable of powering the capsule through LEO eclipse and normal re-entry operations. The two-pass re-entry method used will require additional energy storage. Assuming a continuous power drain of 2 kW for a maximum of 3 days (72 hours), 144 kWh of energy will be needed. If Li-ion batteries with an energy density of 100 Wh/kg and a charging eciency of 0.96 were used, the battery mass would be of 1500kg. Due to

33 Inspiration Mars Student Design Competition Section 9 this large mass, these batteries will be replaced by hydrogen fuel cells. These have extensive space heritage from the Apollo and Space Shuttle programs. The typical energy eciency for a proton exchange membrane fuel cell is 50%. This leads to approximately 20kg of reactants (hydrogen and oxygen) for a total mass of 90kg including the fuel cell and tankage.

Re-entry electrical energy required 144 kWh Chemical energy required 259.2 MJ Mass of hydrogen [56] 2.16 kg Mass of hydrogen tanks [57] 40 kh Mass of Oxygen 17.28 kg Mass of Oxygen Tankage 24.1 kg [58] Fuel cell dry mass at (275 W/kg) 7.27 kg [33] Table 12: Fuel cell requirements

9.2 Sundancer

9.2.1 Power Consumption

During the interplanetary phase the major power consumption will come from life support, waste management, communications and crew activities/science systems. The total load of these systems will be 8188W, as detailed below.

9.2.2 Life Support

Life support systems are nominally provided by the Sundancer module, but their power consumption must be estimated in order to modify the spacecraft’s solar arrays. Assuming a four bed molecular sieve is in use on board the ISS for CO2 removal, and a Sabatier process

O2 generation system were assumed, air management would total 600W of power. For water management 140W were assumed based on filtration and distillation needs. This led to a total system power requirement of 740W.

9.2.3 Waste Management

The peristaltic action of the waste management system is provided by pumping water. The power required to do this was calculated by Bernoullis equation.

M˙ P P = ⇥ ✏ where P is the power; M˙ is the mass flow rate; P is the drop in pressure; ✏ is the eciency of the whole system.

Atypicalcentrifugalpump[59]isusedandamassflowrateof50g/sisassumed.The

34 Inspiration Mars Student Design Competition Section 9 pressure drop is considered to be 5 bar. This would give a power requirement of approxi- mately 333 W.

9.2.4 Communications

In addition to the ComSat’s laser communications system, a conventional radio system will provide redundancy and primary communications on the return to Earth. An X/Ka-band system was sized using the link budget equation to estimate the power required. The X- band link will provide telemetry, data transfer and some communication capability. The Ka-band link will provide video and image transfers on the return travel. A 2.2 m antenna and power of 550 W for the X-band was determined to allow a data rate of 2Mbps. The Ka-band required the same antenna and 300 W of power for a 7.5Mbps data rate. Both systems allowed for a 3 dB margin. All assumptions made in this calculation were the same as for the ComSat. Maximum power for simultaneous communications is therefore 850 W. This also assumes that the radio links to the sub satellite will only require a small amount of power.

9.2.5 Crew Activities / Science

The power required for the crew activities and equipment and the mission science is shown in Table 13.

Activity Required Power (W) Crew Electronics and OBDH 3000 Human Research 1000 Plant growth experiments 2 x 200 W 400 Experiment Freezer 500 Total 4900 Table 13: Power requirements for Crew Activity

9.3 Power Generation

Two limiting cases were identified on the Sundancer’s solar arrays’ power production:

Case A: at Mars flyby, the solar flux will be at its lowest. • Case B: when returning to Earth, radiation degradation of the solar cells will reduce • the power generated.

The lowest power raised from these two cases is used to size the array. Standard assump- tions of inherent degradation factor of 0.77, electrical conversion eciency 0.9 and pointing accuracy of 5o were made.

35 Inspiration Mars Student Design Competition Section 10

9.3.1 Case A

Mars will be close to perihelion during the flyby, at 1.397 AU from the sun. The solar flux at the point of the Mars Flyby is calculated to be approximately 700 W/m2 [60]. Therefore the electrical power output will be 131 W/m2,accountingforradiationdegradation.

9.3.2 Case B

Assuming GaAs based solar cells, the power conversion eciency will be 28 % at beginning of life. This is degraded by radiation exposure by a factor of about 2.75 %/yr in LEO [33]. As the spacecraft is travelling through a harsher radiation environments than LEO, this factor was increased to 10 %/yr. Therefore, for a 501 day mission the eciency is reduced to 26.1%, leading to an electrical power output of 245 W/m2 at the end of life.

Case A is therefore the limiting case for the power generation system. The arrays are sized to this case with an area of 66m2, generating 8360W. This can be divided into two 11m x 3 m arrays deployed. These can be stored along the sides of the Sundancer module during launch.

9.4 Energy Storage

As only 1 eclipse will be encountered outside LEO this does not drive the battery require- ments of this module. Peak loads such as pumping water between radiation shield bubbles may require more power than instantaneously generated. It is dicult to estimate this without further as the power depends on the pressure drop which would have to be found by prototyping or specialist analytical methods. In this case the hydrogen fuel cell on the Dragon should be employed. No additional batteries will be used on the Sundancer other than those installed as standard for LEO.

Energy Considerations during Reentry During the aerobraking and re-entry manoeu- vre the Sundancer and Dragon’s trunk will be jettisoned, hence eliminating the primary source of power. The Dragon capsule will therefore rely on internal power between the initial separation from the other components until the crew is recovered. As only one aer- obraking orbit will be performed before the final re-entry, this length of time will be of 3 days.

36 Inspiration Mars Student Design Competition Section 10

10 Cost Evaluation

Most of the components used in this design are either o↵-the shelf components or modifica- tions of existing technologies. These modifications will need some development investment, but the cost of doing so will still be considerably lower than that of developing a new tech- nology.

10.1 First approach

The paper ‘Manned Mars Mission Cost Estimate’ [61] was used as a reference to aid in estimating the costs of the mission, although the values given were not directly applicable since the estimate was based on values and technology from 1985. As the mission described in the paper included a landing on Mars while this mission does not, the values produced had to be modified in order to account for this. An adjusted comparison yielded an initial estimated average mission cost of $1.91B. This is below the stated maximum of $2B. However, the maximum estimate obtained is $4B, which largely exceeds the requirements. For this reason, further studies were conducted.

10.2 Detailed Cost Study - QuickCost 5.0 Model

One of the authors of the previously mentioned paper, Dr Hamaker, was contacted for further discussion about cost estimation for such a mission. He provided his QuickCost 5.0 model to assist in this task. The model calculates Cost Estimate Relationships (CERs) for a number of mission types based on a large database of information on previous missions. The models that were used to perform the estimation for the mission were the Module & Transfer Vehicle Model for the main spacecraft and the Satellites Model for the ComSat.

10.2.1 Main Spacecraft

The parameters used by the QuickCost 5.0 model included the spacecraft dry mass, design life, whether or not the mission was manned, whether the vehicle was to be used as a transfer vehicle, whether it required a propulsion system and its Initial Operating Capability. The percentage of the spacecraft that consists of new designs is also an input parameter - the nominal value for this is taken to be 50%. However, the dependence of cost on this parameter was also studied. With these parameters it was possible to obtain the median estimated cost for DDT&E (Design, Development, Test and Evaluation) and TFU (Theoretical Flight Unit). The results obtained are shown in the figure below. The confidence level refers to the confidence in meeting the expected budget.

37 Inspiration Mars Student Design Competition Section 10

Figure 21: Main spacecraft cost of the mission as the function of the desired confidence level in percentile

The highlighted region in the graph was seen as the most relevant to this mission. With more than 50% new components, and with a confidence level of 75% a total cost estimate of $2.18B was obtained. If the development costs were reduced to a smaller percentage (30%) of new components, this cost could be reduced to $1.63B, with a confidence level of 75%.

The following table shows the QuickCost inputs and results for the nominal case.

Total Dry Mass (kg) 31,800 New Design (Percentile) 50% Design Life (months) 18 Human Rated? 1 Transfer Vehicle Capability (i.e. Propulsion) 1 Initial Operating Capability (IOC) 2018 Non-Market Economy (i.e. Russian) 0.0 Year Dollar of Output (20XX) 2014 Median Estimated Cost $1,200.3 Median Estimated Schedule (months) 54.7 Desired Confidence Level (percentile) 75% Median Cost + Reserve to Achieve Desired Confi- $1,474.4 dence Level Median Schedule + Reserve to Achieve Desired Con- 67.2 fidence Level Calculated MCPLX for PRICE Users 10.82 Calculated Peak Size of the Government Project Of- 143.1 fice Ground System Cost $88.5 Launch Services Cost $474.4 Annual MO&DA Cost $73.7 Life Cycle Cost (Over Design Life) $2,184.6 Table 14: QuickCost analysis for main spacecraft. All cost values are given in Millions of Dollars.

10.2.2 Communications Satellite

A similar analysis has been performed for the laser communications satellite, using the Satel- lite Model from QuickCost 5.0. Parameters include the total instrument complexity, assumed to be 65% doe to high AOCS pointing and stability requirements, and the percentage of the

38 Inspiration Mars Student Design Competition Section 10 bus and instruments which consist of new designs, assumed to be 40% for both as existing designs are being modified. With these parameters, the total life cycle cost of the ComSat, including development, production, ground system, launch and operations is estimated to be around $303M.

10.3 QuickCost Results

From the QuickCost model, the total development and production costs of the mission, including the ComSat, through phases B/C/D will be around $1.45B, and the total life cycle cost is $2.49B with the assumptions stated above.

10.4 Cost Scheduling

An analysis was performed to determine how costs will be spread over the mission lifetime. A Beta function showed that 60% of the total funds were spent in the first half of the mission life cycle. Annual MO&DA Costs were considered at an additional 5%/year on top of these. This was then adjusted for inflation using the NASA New Start Inflation Index [62].

Figure 22: Time Cost Distribution

10.5 Validation

To validate these cost estimates di↵erent models were used and their results were compared.

10.5.1 Communications satellite

A second cost estimation for the ComSat was performed using the USCM8 method, outlined in SMAD [33] and adjusted for inflation. Due to schedule constraints and the fact that

39 Inspiration Mars Student Design Competition Section 10 most of the components have considerable flight heritage, it was assumed that a protoflight method will be used. For this reason, the non-recurring cost CERs were used to calculate subsystem costs and a 30% margin was included to cover refurbishment costs. The results of this analysis are shown below.

Table 15: USCMD8 Costing of the ComSat

Element Cost (FY14$K) Standard Error Min Cost (FY14$K) Max cost (FY14$K) Structure/Thermal Control 18317.40 22% 14287.57 22347.23 ADCS 37536.36 44% 21020.36 54052.36 EPS 3857.52 41% 2275.93 5439.10 Insertion Engine 5860.35 22% 4571.07 7149.63 Orbit Acquisition Engine 3461.72 22% 2700.14 4223.30 TT&C 31768.58 n/a 31768.58 31768.58 Laser 21453.66 Electra 3218.05 Narrow Angle Cameras 11197.74 Wide Angle Camera 2356.37 Payload Total 38225.82 39% 23317.75 53133.89 Integration, 26074.77 42% 15123.37 37026.18 Assembly and Testing Program Level 158306.96 37710.81 28% 27151.78 Ground Equipment 31258.24 37% 19692.69 42823.79 Reaction Control System 18627 2526.51 35% 1642.23 Total Cost 307577.49 Minimum Cost 212616.92 Maximum Cost 402538.06

As can be seen, the estimate given by this model ($288M) is similar to that given by the QuickCost model ($303M). Note that since the USCM8 model only provides CERs for com- munications payloads, the NASA Instrument Cost Model, also given in [33] was used to estimate the costs of the Narrow Angle Cameras and Wide Angle Camera. The costs for the Electra package and the laser subsystem were estimated based on the relative masses and costs of similar missions carrying these payloads.

10.6 Main spacecraft - NAFCOM Weight-Based Model

The second estimation of the main spacecraft cost used the NAFCOM method which ex- trapolates from other space projects, with mass as a scaling parameter. In 2011 a NAFCOM study was carried out on the Falcon 9 launch vehicle with a total cost estimated between $4B (traditional approach) and $1.7B (commercial approach). SpaceX had only spent $390 M, highlighting the radical cost savings of their methodology. If our mission can be more cost ecient than traditional aerospace projects the NAFCOM model may be overly conservative.

NAFCOM looks at the di↵erence between manned and unmanned missions and to iden- tifies the main characteristics which can be compared to the mission at hand. The manned mission data was taken mainly from 60 NASA missions (Apollo, Gemini, Space Shuttle Or- biter, Skylab and SpaceLab).

Acostapproximationcanthenbemadeforthismissionusingthecomplexityfactor,↵,

40 Inspiration Mars Student Design Competition Section 10 which is an adjustment to the CER to compensate for the projects unique features which are not accounted for in the historical data. It also takes into account the fact that this mission is using a higher percentage of COTS components when compared to those missions.

By assuming a range of ↵ =0.4to↵ =0.5, and a total spacecraft mass of 30 metric tonnes, a development cost ranging between $2031M and $2539M was obtained, whereas the Flight Unit production cost went from $242M to $303M. Thefore, this model predicted a total mission cost between $2273M and $2842M.

By varying the NAFCOM approximation parameter values for each individual subsystem, amoreaccuratecostoftheoverallmissioncanbeobtained.Thefollowingtableshowsthe results obtained when applying this method, with a confidence estimation level of 90%.

Using historical data points obtained from previous analogous space missions, the ”first pound cost” approach from NAFCOM can be used. The same model was used for most sub- systems but the fact that many COTS components were used was also taken into account. The ensuing risk assessment analysis was performed to provide a cost estimate for di↵erent confidence levels. The results can be seen in the table 16. These results provide validation of the QuickCost model, as the total mission cost is esti- mated as $2.49B. As mentioned above, this cost can be reduced by lowering the development investment in the development of new technologies. The radiation shielding and waste man- agement subsystems are the ones that require the highest amount of development, so reducing the requirement on the level of protection would significantly lower the cost

Model Spacecraft Min Cost $B Nominal Cost $B Max Cost $B QuickCost ComSat 0.228 0.303 0.490 USCM8 ComSat 0.213 0.308 0.402 QuickCost Main 1.87 2.19 3.03 NAFCOM (Spacecraft Level) Main 2.27 3.06 3.84 NAFCOM (Subsystem Level) Main 2.05 2.28 2.51 Table 16: QuickCost Minimum estimated at 50% confidence, Nominal at 75%,Maximumat99%

41 Inspiration Mars Student Design Competition Section 11

Figure 23: NAFCOM Subsystem Analysis

11 Mission Phases and Scheduling

This section provides a general overview of the various mission phases and an outline of a potential schedule.

11.1 Development Schedule

The concept presented has been designed so that most of the elements are COTS (Commer- cial O↵The Shelf) so as to reduce development costs. However, some will need time to be tested and developed, as it is discussed below.

42 Inspiration Mars Student Design Competition Section 11

Waste management: While the reverse osmosis membrane material will be ready on time, the design and verification of the whole pump system will need to be done. As this is a versatile water treatment system, additional funds may be available to accelerate develop- ment.

Bubble radiation shield: This innovative system will require significant testing on Earth and potential verification aboard the ISS. Prototyping can begin very quickly with low cost materials for rapid iteration.

Laser communications: The concept of laser communication from space has been demon- strated on board the LADEE spacecraft. Moreover, the OPALS laser communications pay- load is scheduled for launch to the ISS in March 2014. [23]

Modification of the Dragon and Sundancer: It is anticipated that the Dragon and the Bigelow Expendable Activity Module (BEAM) will be flown with crew before 2018, but additional development to take these spacecraft beyond LEO must be considered. Even if both Bigelow Aerospace and SpaceX wish to make these alterations, the timescale for such development is unknown.

The QuickCost 5.0 model 10 also provides a schedule estimate based on parameters from previous missions, giving a development and production time of 54 months. While this is on pare with the available time, it does not take into account either the accelerating factor of a deadline or the high proportion of COTS hardware that this mission will have.

11.2 Launch and Docking Schedule

11.2.1 Launch and mass distribution

As previously stated 7, the total payload to be distributed amongst the four Falcon Heavy launchers will be of 188.8 tonnes. It has been calculated that the use of Merlin 1D engines for the injection into the desired trajectory would require a combined fuel and tank mass of 157 tonnes. The total dry mass of 31.8 tonnes is to be distributed between the Dragon capsule, which is specified as having a dry mass of 4.2 tonnes [47] and carries a cargo of 6 tonnes, and the Sundancer, with a mass of 8.6 tonnes [1], carrying 4.4 tonnes of cargo.

The two initial Falcon Heavy launches will be dedicated entirely to propellant and tanks, with a total of 47.2 tonnes each. These will be followed by the Sundancer and its cargo,

43 Inspiration Mars Student Design Competition Section 11 as well as a total of 25.6 tonnes of additional propellant. The final Falcon Heavy will carry the Dragon capsule, crew and additional cargo, as well as the remaining 37 tonnes of fuel required.

11.2.2 Launch sequence

Within the next few years, SpaceX hopes to increase its rate of production and launch to ten Falcon Heavy launchers and ten Falcon 9 launchers per year [43]. This corresponds to a minimum delay of about five weeks between each launch. However, a contingency should be applied to this, and therefore two months (about eight weeks) will be taken as the ap- proximate delay between each launch. An additional constraint on the schedule comes from the launch site. By assuming a time delay of two months this would cease being an issue, as demonstrated by SpaceXs two Falcon 9 launches from Cape Canaveral in less than a month (13 December 2013 and 6 January 2014).

This two-month delay will also allow for time to perform the successive in-orbit assem- blies as well as the ensuring verifications. The Merlin 1D engine is fuelled by kerosene and liquid oxygen so propellant boil-o↵is not as much of an issue as it would be for a LOx/LH2 engine. Therefore, the propellant payloads can be safely stored in orbit for longer periods of time. For this reason, the two Falcon Heavy launchers carrying propellant will be launched first, followed by the combination of the Sundancer module with the added cargo and the remaining propellant.

In order to limit the time the astronauts spend in space, which in turn will reduce the dose of radiation inflicted, the food consumed and the physiological e↵ects of habitation in space, the manned Dragon will be launched in the very last instance.

11.2.3 Docking and in-orbit assembly

All rendezvous manoeuvres and docking will be performed autonomously, and will be mon- itored from the ground. Upon completion of the second docking manoeuvre, once the Sun- dancer is attached to the two propulsion modules, the module will be automatically ex- panded. A remote verification of the functionalities of the system will confirm the correct deployment of the module.

Following the docking of the Dragon capsule to the Sundancer module, the crew will en- ter the expandable module and assemble the bubbles. This operation should only be a few

44 Inspiration Mars Student Design Competition Section 11 hours. The assembly of the waste management system will ensue. As this operation may take up to several days, it does not need to be complete before departure from Earth orbit. Following a final verification, the green light will be given for the injection into Mars orbit, with the propulsion burn being accomplished on the 5th of January 2018. The full launch, docking and verification sequence is shown in Fig. 25.

Figure 24: Timeline of launches (orange), dockings (blue) and equipment verifications (light green) before injection into Mars trajectory.

11.3 Interplanetary Events

Following the docking manoeuvres and subsequent verifications, the authorisation will be given to proceed, and the injection into Mars trajectory will be performed on the 5th of January 2018. During this manoeuvre, each Falcon Heavy upper stage will consume the fuel of its additional fuel tank. Then, each stage will be jettisoned, before the ignition of the following stage. Details of the relevant delta v achieved and of the remaining mass can be found in the Launch and Injection into Mars trajectory section, 7. The entire burn sequence will last less than 30 minutes, and will be followed by a final verification.

On the 6th of January,oncethespacecrafttrajectoryhasbeenconfirmedandverified, the ComSat will be deployed through a spring-loaded mechanism. This will ensure that the satellite and the main craft travel as two separate entities, although the di↵erence in velocity between the two will be insignificant and will hence not a↵ect their respective trajectories. 6 Arelativevelocityof25mmperhour,6.94 10 m/s, would prove optimal as it would be ⇥ negligible in orbital calculations, whilst still providing a sucient separation between the two spacecraft within a few days.

On the 9th of January,assumingarelativevelocityof25mmperhour,theseparation

45 Inspiration Mars Student Design Competition Section 11 between both spacecraft will reach 2m. At this point, communication between the two spacecraft will begin, and the ComSat will be used as a relay satellite for all primary com- munications between the mission and Earth.

The fly-by of Mars will occur on the 20th of August 2018. By this time, assuming a relative velocity of 25mm per hour, the distance between both spacecraft will be a safe 135.6m. Upon reaching the periapsis of the fly-by trajectory, the ComSat will perform a burn in order to reduce its velocity and adopt an elliptical orbit about Mars, with a semi- major axis of 28470 km and an eccentricity of 0.994. Subsequent orbital manoeuvres will finally place the ComSat into a circular orbit about mars, at an altitude of 500km (3890km radius). The combined Dragon and Sundancer spacecraft will remain on its unaltered course, returning towards Earth.

The Earth approach will occur on the 21st of May 2019.TheSundancermoduleandthe Dragons trunk will be separated and jettisoned, as the crew will take place in the Dragon capsule alongside food supplies for five days. A first pass through the Earths atmosphere, as described in the Reentry section, will provide the aerocapture: the capsule will be in an elliptical orbit about the Earth. An orbit around this ellipse will take three days. Follow- ing this, on the 24th of May 2019,thefinalre-entryintotheEarthsatmospherewilloccur.

Therefore, the interplanetary travel will last between the injection into Mars trajectory on the 5th of January 2018, and the final Earth re-entry on the 24th of May 2019, or 504 days. When including the scheduled launch date of the astronauts on the 31st of December 2017, the entire manned section of the mission will last 509 days.The sequence of events during the interplanetary trajectory can be seen in Fig. 26.

Figure 25: Interplanetary trajectory and events schedule

46 Inspiration Mars Student Design Competition Section 12

11.4 Risk mitigation and 2021 launch window

Many of the mission components, including the launch segment, the Dragon capsule and the Sundancer module, are to be developed and produced by private companies. This not only means that their development expenses are accounted for by these companies, but also that their development time may prove shorter than that expected for an agency such as NASA. Additionally, the Falcon Heavy will perform its first flight late in 2014, the unmanned Dragon capsule is already in operation and progress is being made towards a manned version, whereas most of the technologies used by the Sundancer will be demonstrated by BEAM starting in 2015 [3].

Similarly, the satellite communication systems to be used on the ComSat are being widely researched, through such missions as LADEE or OPALS, and therefore should then be a proven technology by 2018. Table 1 below gives a risk mitigation matrix, in which delays in each components development are met with their likely e↵ect on the mission. As this table shows, the radiation shield, the Sundancer module, the Dragon capsule and the Falcon Heavy launcher are critical to this design, as their unavailability would lead to abandon the 2018 launch opportunity in favour of that in 2021. However, it is expected that each of these elements will be ready on time.

Conversely, this mission design will not require large modifications if the 2021 launch win- dow were to be adopted. Indeed, despite the longer flight time (597 days from launch of the crewed Falcon Heavy until touchdown, considering the same launch and docking sequence as for the 2018 mission). Food supplies for the additional 88 days will have to be accounted for, but despite the radiation levels being higher due to the solar minimum experienced in 2018 no longer being a benefit, the radiation shield designed for the original scheduled mission will protect the crew suciently to ensure their health and well-being.

Mission component Marginal delay (not ready by 2018) Severe delay (not ready by 2021) Waste management Replace with Isonomer membrane bags, de- Replace with Isonomer membrane bags, de- spite loss in functionality spite loss in functionality Laser communication Replace with high gain antenna, despite loss Replace with high gain antenna, despite loss of data rate of data rate Bubble radiation shield Postpone mission to 2021 window Attempt to replace with polyethylene layers Sundancer module Postpone mission to 2021 window Attempt to use Cygnus- or ATV-type of module Manned Dragon capsule Postpone mission to 2021 window Attempt to use Orion capsule, despite cost considerations Falcon Heavy launcher Postpone mission to 2021 window Attempt to use SLS launcher, despite cost considerations Table 17: Risk and consequence matrix for the mission components which require development. Note that the four last elements are critical to the mission, as a delay will imply opting in favour of the 2021 launch window instead.

47 Inspiration Mars Student Design Competition Section 12

12 Conclusion

This report presented innovative solutions to the problems of interplanetary travel and elab- orated a feasible architecture for Inspiration Mars 2018 fly-by mission. Using COTS hard- ware and incorporating novel technologies and approaches enabled a low mission cost with outstanding system performance. This design employs a Dragon capsule and a Bigelow Sun- dancer module, both of which are close to completion, as well as four launches with the Falcon Heavy, which will perform its first flight before the end of 2014.

The spacecraft’s radiation shield will employ water to provide protection, in addition to the radiation protection provided by the structure of the Sundancer habitat. The waste man- agement system will not only allow for automated disposal of waste and water recycling, but also provide additional radiation protection to the spacecraft. The ComSat will implement laser communications in order to provide unprecedented data return from Mars, providing adefiningmomentforspaceexplorationaswellasexpandinghumanity’sinterplanetaryin- frastructure. This satellite will remain in Martian orbit and constitute the cornerstone to the planet’s subsequent exploration and potential colonisation.

Four Falcon Heavy launchers will be required, for a total payload into 800km altitude orbit of 188.8 metric tonnes. The injection into Mars trajectory will be performed by reusing the upper stages and using some additional fuel carried into orbit, and this will allow for a final dry mass of 31.8 tonnes to be transported to Mars. The Falcon Heavy launch rate is sustainable but ensures critical elements will have a minimum loiter time in LEO before trans-Martian injection. The cost of the mission was calculated with QuickCost and verified with several other models. This gave a total cost estimate of $2.49B with a 75% confidence level, $2.10B with a 50% confidence level and $2.85B with a (90%)confidence level.

48 Inspiration Mars Student Design Competition Section 12

References [18] E. Dadachova and A. Casadevall, “Melanin nanoshells for protection against radiation and [1] B. Aerospace, “Sundancer.” electronic pulses,” Nov. 19 2013. US Patent [2] B. Aerospace, “Next Generation Orbital Space 8,586,090. Stations.” [19] P. Eckart, Spaceflight life support and biospher- [3] B. Aerospace, “”expanding humanity’s future in ics. Torrance,CA; Dordrecht: Microcosm Press; space”.” Website. Kluwer Academic Publishers, 1996. by Peter Eckart.; Revised and expanded from: Life sup- [4] B. Aerospace, “BA330.” port and biospherics, Herbert Utz Publishers, [5] “The earth’s magnetosphere.” Munich, Germany, 1994.; ASE; Life support and biospherics. [6] A. Svitak, “”bigelow tests life support system”,” April 2011. [20] A. I. of Aeronautics, Astronautics, and A. N. S. Institute, Guide for estimating and budgeting [7] “”atv integrated cargo carrier”,” 2013. weight and power contingencies for spacecraft [8] “”orion: America’s new spacecraft for human systems. Washington, DC: American Institute exploration”.” Website. of Aeronautics and Astronautics, 1992. spon- sor, American Institute of Aeronautics and As- [9] S. G. et al., “”radiation protection for interplan- tronautics ; approved, American National Stan- etary manned missions”,” tech. rep., IFN, Italy; dards Institute.; Approved, April 16, 1992. ESA-ESTEC, 2005. [21] NASA, “Anthropometry and biomechanics,” [10] D. Rapp, “Radiation e↵ects and shielding re- Tech. Rep. NASA-STD-3000 275e, NASA, 2008. quirements in human missions to the moon and mars,” Mars, vol. 2, pp. 47–71, September 2006. [22] A. B. et al., “Deep-space optical terminals (dot) systems engineering,” tech. rep., NASA JPL, [11] T. Phillips, “Van allen probes discover a new 2010. radiation belt,” Feb. 28 2013. Science@NASA. [12] J. V. Bailey, “Radiation protection and instru- [23] D. Washington, “Laser communication relay mentation,” Feb. 28 2013. Science@NASA. demonstration, the next step in optical commu- nications,” tech. rep., NASA, 2010. [13] T. ERSMARK, “”geant4 monte carlo simula- tions of the geant4 monte carlo simulations of [24] Mars One, Road Map. the geant4 monte carlo simulations of the inter- [25] “Mars one: Communications system.” national space station radiation environment”,” 2006. [26] D. e. a. Boroson, “The lunar laser communi- cations demonstration (llcd),” in Space Mission [14] M. M. Cohen, “Carbon Radiation Shielding for Challenges for Space Mission Challenges for In- the Habot Mobile Lunar Base,” in 34th Inter- formation Technology,IEEE,ThirdIEEEInter- national Conference on Environmental Systems national Conference, 2009. (ICES), Technical Paper, (Colorado Springs), SAE, July 2004. [27] S. T. et al., “The mars laser communication demonstration,” tech. rep., NASA JPL, 2014. [15] L. W. T. J. E. N. J. W. W. L. C. Simon- sen, “Estimates of Galactic Cosmic Ray Shield- [28] C. Edwards, “The electra proximity link payload ing Requirements During Solar Minimum,” tech. for mars relay telecommunications and naviga- rep., National Aeronautics and Space Adminis- tion,” tech. rep., NASA JPL, 2014. tration, February 1990. [29] H. Hemmati, “Interplanetary laser communica- [16] G. D. Q. J. E. Nealy, L. C. Simpson, “Sei analy- tions,” tech. rep., The Optical Society, 2007. sis of deep space vehicle shielding,” in Shielding [30] F. K. K. F. I. K. J. W. Burnside, D. V. Murphy, Strategies for Human Space Exploration (A. K. “A hybrid stabilization approach for deep-space F. A. C. J. W. Wilson, J. Miller, ed.), vol. NASA optical communications terminals,” Proceedings Conference Publication 3360, NASA, 1997. of the IEEE, vol. 95, pp. 2070–2081, October [17] P. Aerospace., “Ionomer membrane water pro- 2007. cessor,” Nov. 19 2013. IWP Development pro- [31] “Lunar reconnaissance orbiter camera.” gram.

49 Inspiration Mars Student Design Competition Section 12

[32] H. Hemmati, “Deep-space optical terminals [49] B. G. Drake, S. J. Ho↵man, and D. W. Beaty, (dot),” tech. rep., NASA JPL, 2014. “Human exploration of mars, design reference architecture 5.0,” in Aerospace Conference, 2010 [33] J. P. J. Wertz, D. Everett, Space Mission En- IEEE, pp. 1–24, IEEE, 2010. gineering: The New SMAD. Microcosm Press, 2011. [50] L. B. J. S. Center, “Apollo experience report- thermal protection subsystem,” 1974. [34] Honeywell, Honeywell Constellation Series Re- action Wheels. [51] S. Nowlin and L. Thimons, “Surviving the heat: The application of phenolic impregnated carbon [35] SSVB, Earth Horizon Sensor. ablators,” [36] M. B. G. Ruyters, “Plant biology in space: Re- [52] S. S. W. Denise Chow, “Q a with spacex ceo cent accomplishments and recommendations for elon musk: Master of private space dragons,” future research.” 2014. [37] V. N. Svchec, “Space experiments on the devel- [53] J. A. Samareh, R. W. Maddock, and R. G. Win- opment of biological systems for the human life,” ski, “An integrated tool for system analysis of Aviakosmicheskaya i Ekologicheskaya Meditsina, sample return vehicles,” in Aerospace Confer- vol. 47, no. 1, pp. 43 – 46, 2013. ence, 2012 IEEE, pp. 1–8, IEEE, 2012. [38] “fitbit wearable electronics.” [54] R. Mattingly and L. May, “Mars sample return [39] “Google glasses wearable electronics.” as a campaign,” in Aerospace Conference, 2011 IEEE, pp. 1–13, IEEE, 2011. [40] D. e. a. Dinges, “Optical computer recognition of facial expressions associated with stress in- [55] “Dragonlab, fast track to flight.” duced by performance demands,” in Aviation [56] D. of Energy, “Hydrogen Calculator.” Space and Environmental Medicine, Conference paper, SAE, June 2005. [57] Q. Technologies, “Example COTS Hydrogen tank.” [41] S. S. J. Taylor, D. Lee, “Mars reconnaissance orbiter telecommunications,” tech. rep., NASA [58] cpapusa.com, “Example COTS LOX tank.” JPL, 2006. [59] www.pump zone.com, “Centrifugal pump e- [42] D. e. a. Dinges, “Statistical entry, descent, ciency.” and landing performance reconstruction of the [60] W. Alpha, “Mars Solar Irradiance in August Mars Science Laboratory,” in AIAA Atmo- 2018.” spheric Flight Mechanics Conference, Confer- ence paper, AIAA, January 2014. [61] J. H. K. Smith, “Manned mars mission cost esi- mate,” tech. rep., Marshal Space Flight Center, [43] February 1987. [44] SpaceX, Falcon 9 Launch Vehicle Payload User’s [62] Guide, 2009. [45] D. A. Tito, G. Anderson, J. P. Carrico, J. Clark, B. Finger, G. A. Lantz, M. E. Loucks, T. Mac- Callum, J. Poynter, T. H. Squire, et al., “Fea- sibility analysis for a manned mars free-return mission in 2018,” in Aerospace Conference, 2013 IEEE, pp. 1–18, IEEE, 2013. [46] “Rl10 engine.” [47] “The dragon capsule.” [48] R. Crowder and J. Moote, “Apollo entry aero- dynamics.,” Journal of Spacecraft and Rockets, vol. 6, no. 3, pp. 302–307, 1969.

50