A Feasibility Study of a Mars Scout Vehicle to Study Methane

Transcription

1 GIT SMART: A Mission to Measure Atmospheric Methane on Mars Matt Daskilewicz, Zhi Deng, Kathy Goben, Ramraj Harikanth, Erik Kabo, SoYoung Kim, Kybeom Kwon, Kathleen Stokes, Daili Zhang Faculty Advisor: Prof. Dimitri N. Mavris April 24, 2006 A Feasibility Study of a Mars Scout Vehicle to Study Methane May 21-25, 2006 i

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3 A Feasibility Study of a Mars Scout Vehicle to Study Methane Advisor: Dr. Dimitri Mavris Program Manager: Erik Kabo Chief Engineer: Kathy Goben Team Members: Matt Daskilewicz, Zhi Deng, Ramraj Harikanth, So Young Kim, Kybeom Kwon, Kathleen Stokes, and Daili Zhang Abstract The recent discovery of methane, a known biomarker, in the Martian atmosphere has generated sufficient scientific interest to warrant in situ studies of methane on Mars. This study lends itself to NASA s new, costeffective class of missions called Mars Scout missions, and would be addressing the Mars Exploration Program Analysis Group s first goal, determine whether life ever arose on Mars. This paper addresses the feasibility of a vehicle performing a Mars Scout mission to characterize Martian methane by describing the design of a complete Scout mission architecture. Only a UAV platform met all operational requirements, and out of all possible UAV concepts, the unpropelled balloon provided the most feasible, cost-effective option for this particular mission. Our feasibility study concludes that an uncontrolled balloon, also called an aerobot, is the best design for the next step in Martian methane research. The design resulted in four aerobots, flying at altitudes of 6 km, 7 km, 8 km, and 9 km, that are delivered to Mars in a single aeroshell. With present day technologies, no other concept can achieve the science goals at a competitive cost and acceptable level of risk, and the Mars Scout aerobot concept offers the lowest risk, most cost-effective solution for the methane characterization mission. Introduction In 2004, President George W. Bush announced the Moon, Mars, and beyond vision for space exploration. In order to plan for human exploration of Mars, we must first obtain a better understanding of the planet by conducting robotic exploration missions. 1 NASA has formed the Mars Exploration Program Analysis Group (MEPAG) to prioritize the scientific and technological objectives that must be accomplished, with the first goal being to determine whether life ever arose on Mars. 2 The recent discovery of methane in the Martian atmosphere directly relates to this goal because methane is a known biomarker. This discovery has generated sufficient scientific interest to warrant studies of methane on Mars, the first of which should verify the discovery and attempt to identify the methane source. This study lends itself to NASA s cost-effective class of missions called Mars Scout missions. 3 This paper addresses the feasibility of a vehicle performing a Mars Scout mission to characterize Martian methane by describing the design of a complete Scout mission architecture that resulted from a series of downselection studies. Throughout the design process, the design space was kept as open as possible, with downselection based on the evaluation of how well each concept fulfilled the science goals and Scout mission constraints. Whenever possible, physics-based models were created and used to evaluate designs quantitatively. Scientific Goals Observations made by the Planetary Fourier Spectrometer on the ESA Mars Express Orbiter in 2004 detected trace concentrations of methane in the Martian atmosphere. 4 Given methane s relatively short chemical lifetime on Mars, it is hypothesized that a methane source currently exists on Mars to supply this concentration. Furthermore, methane is a known biomarker, with biogenic methane comprising nearly all methane sources on Earth. 5 Thus there is considerable interest in studying Martian methane to support the search for life on Mars, one of the four science goals for Mars identified by MEPAG. 2 If sustainable methane sources can be identified on Mars, there may also be potential to collect Martian methane as a fuel source for future exploration missions, 6 giving the mission a practical, as well as purely scientific, validation. In situ studies of the Martian atmospheric methane will allow for simultaneous studies of the structure and dynamics of the Martian atmosphere, which will improve the understanding of spacecraft entry, deployment and operations on Mars and reduce risk in future exploration missions. This knowledge will become critically important as NASA plans manned missions to Mars, and so an in situ atmospheric study is warranted. Capabilities Required A study of the methane on Mars will characterize methane concentrations by seasonal and diurnal cycles, and ambient conditions. Concentrations have been detected on the level of 10 ppb, significantly lower than the 2

4 average Earth value of 1700 ppb. 4 Thus a meaningful characterization will require very precise measurements, on the order of 1 ppb, and this became a design requirement for the science instrumentation. The most promising technique for achieving this level of precision is IR spectroscopy, which measures trace gas concentrations by their characteristic absorptions of certain infrared wavelengths. In order to further characterize the methane source as biogenic (produced as a byproduct of life) or abiogenic, the isotopologue ratios 13 CH 4 / 12 CH 4 and 12 CH 3 D/ 12 CH 4 must be measured. As observed on Earth, biogenic methane will exhibit lower isotopologue ratios than the naturally occurring ratios of these isotopes. This is attributed to the phenomenon of isotope fractionation, whereby organisms preferentially take in lighter isotopes (in this case 12 C and H). 7 It is believed that Martian organisms would also exhibit fractionation, thus by comparing observed values for these ratios with the natural Martian ratios of carbon and hydrogen isotopes, we can determine if the methane is biogenic. Because these ratios are very small, typically between 10-5 and 0.01, current methods of trace gas detection are not sensitive enough to make these measurements remotely; measurements must be made within the atmosphere. Isotopologue measurements are typically made using a mass spectrometer; however recent advances in optics technology will allow this function to be performed by an infrared (IR) spectrometer at a significantly reduced mass and power. In addition to measuring methane concentrations, the mission will take high-frequency measurements of temperature, pressure and wind velocity. These measurements will allow for a long-term, wide ranging study of Martian atmospheric dynamics at a resolution not previously possible. Platform Downselection The science instruments must be integrated into a platform that meets all operational requirements while minimizing mission risk. For example, measurement of the isotopologue ratios requires that the spectrometers be immersed in the atmosphere under consideration. Additionally, scientists currently disagree on the possible locations of methane within the Martian atmosphere, and so the platform must have as large a search area as possible. 8 This requires that the platform s range is maximized, and that the platform can operate in multiple altitudes. Furthermore, long endurance allows the platform to prolong its search, allowing for a long-term study of Martian methane. The required flight characteristics of the science instruments drive the platform decision; the optimal platform will be capable of operating within the atmosphere, over large ranges and multiple altitudes, and for a long endurance. The four platform categories identified as being available for a Mars mission are lander, rover, unmanned aerial vehicle (UAV), and orbiter; these platforms capabilities with respect to the required characteristics are summarized in Table 1. Table 1: Platform Selection Platform Long Endurance Large Range Multiple Altitudes Lander Rover UAV Orbiter Atmospheric Immersion The UAV was selected because it is the only platform that achieves all required flight characteristics. A major turning point in the design process, this first downselection decision restricts the design space to vehicles capable of atmospheric flight. All subsequent design decisions stem from the platform downselection. Concept Generation With the selection of the UAV platform, the design space broadens to include all of the possible UAV concepts. Brainstorming identified the chief functions that define flight-capable vehicles and possible profiles, as well as alternatives for each function. These functions and their alternatives were placed into a matrix of alternatives, provided in Table 2. Selecting one alternative from each function creates a vehicle-profile concept, most of which are infeasible. Four general types of vehicle-profile concepts emerge from this matrix of alternatives. A fixed-wing vehicle utilizes forward velocity and a propulsion system and typically has full control over its direction and altitude. Similarly, a rotorcraft vehicle may behave in the same way but obtain lift via rotational velocity. Such a vehicle may be simplified to vertical take-offs and landings and no control over its direction, being blown by the wind to a new location. An airship with a propulsion system can occasionally control its direction while being blown by the wind. The simplest concept is an unpropelled, uncontrolled balloon without a propulsion system. 3

5 Vehicle Flight Profile Table 2. UAV Concept Matrix of Alternatives, with Sample Concept Highlighted Functions Alternatives Source of lift Forward velocity Rotational velocity Lighter than air Source of thrust Propulsion system Wind driven Propulsion plus wind Range < 10 km < 100 km < 1000 km > 1000 km Endurance Hours Days Months Years Artificial intelligence None Low Medium High Degree of control No control Altitude Direction Altitude & direction Landings Continuous flight Flight with landings All concepts were subjected to a qualitative analysis in order to further pare down the concepts to the optimal options. For instance, a fixed-wing vehicle that lands and takes off introduces much more complexity and a higher level of risk into the design without offering any advantages over a fixed-wing vehicle in continuous flight. Similarly, airships and balloons do not gain any advantages by landing. The analysis resulted in the identification of an optimal profile for each vehicle type, and the most advantageous concepts are summarized in Table 3; these concepts are studied in more detail to determine the optimal vehicle-profile configuration. Concept Fixed Wing Rotorcraft Rotational velocity Airship Table 3. Optimal Concepts Vehicle Functions Flight Profile Attributes Artificial Degree of Source of Lift Source of Thrust Range Endurance Intelligence Control Altitude and Forward velocity Propulsion system < 1000 km Months Medium to high direction Lighter than air Landings Propulsion and < 100 km Months Low to medium Altitude Yes wind, or propulsion Propulsion and < 1000 km Months Low Direction No wind Balloon Lighter than air Wind < 1000 km Months Low None No No Vehicle Concept Selection Having established the need for an airborne platform and generated a number of configuration and operational concepts, a selection process was needed to select the right concept for this mission. The first step of this process was to define a set of selection criteria by which the concepts could be ranked. Next a conceptual sizing of each concept was carried out for a range of mission profiles and sizing variables. Each concept was then ranked against the downselection criteria, using preliminary sizing analyses and historical trends where appropriate. This led to the selection of a best concept, which was carried on to the next phase of design. Concept Selection Criteria Definition The selection criteria included both quantitative and qualitative figures of merit (FOMs) based on the mission constraints and science goals. Scout missions are subject to launch vehicle and cost constraints. The mission will use the largest launch vehicle provided for scout missions, the Delta II 7925, and a Viking heritage aeroshell. This launch, entry, and deployment system constrains the concept to a maximum dimension of 2.65 m and a maximum mass of 568 kg. If proposed in the 2006 Mars Scout competition, the mission must be capable of being launched by December 31, 2011, and Principal Investigator costs are capped at $475 million, in real year dollars. 9 Additionally, a set of qualitative FOMs was used to rank the concepts in terms of criteria that could only be estimated at this stage in the design process. These included scientific capability, probability of success, required complexity, feasibility, and trajectory control. Concepts were ranked in these FOMs based on a combination of current research, historical trends and engineering experience. Vehicle Concept Analysis Faced with choosing between four very different vehicle concepts, the set of sizing variables was chosen to be configuration-independent whenever possible. In order to make the most objective comparison, the sizing variables values were chosen to represent the state-of-the-art baseline for each concept, while the mission- 4

6 dependent variable ranges were kept constant across all concepts. In this way, each concept may be analyzed as the best in its class, while the capability provided by each is held nearly constant. The final list of sizing variables and their ranges is shown in Table 4. For each concept, the total vehicle mass was divided into four components: fixed equipment mass, power subsystem mass, lifting surface mass, and propulsion subsystem mass. It is anticipated that the fixed equipment mass will be driven primarily by the science instrumentation, and as such may be considered independent of the vehicle type at this level of detail. The other three masses are scaled according to the propulsive power and lift force required to maintain flight, defined as a constant velocity cruise for the fixed-wing and airship configurations, and a hover for the rotorcraft and balloon. Table 4. Sizing Variables for Vehicle Concept Downselection Fixed-wing Rotorcraft Airship Balloon Fixed Equipment Mass (kg) 5 to 50 5 to 50 5 to 50 5 to 50 Power Subsystem Density (kg/w) to to to 0.25 n/a Lifting Surface Density (kg/m2) Propulsion Subsystem Density (kg/w) n/a Altitude (km) 1 to 10 1 to 10 1 to 10 1 to 10 Velocity (m/s) 0 to to to 200 n/a Lifting Surface Aspect Ratio 5 to 30 5 to 50 n/a n/a Because the mission is intended to be a long-term study of the Martian atmosphere, only renewable power supplies were considered. As such, the power system density is varied to represent different candidate technologies, with the lower limit corresponding to solar cells and the upper limit corresponding to radioisotope thermoelectric generators (RTGs). (Though not strictly renewable, RTGs would provide power for the entire mission without consuming mass.) The propulsion system is assumed to comprise small electric motors with propellers, fans or rotors to convert the power to thrust or lift force. Finally, the lifting surface includes the wing or rotor for fixed-wing and rotorcraft respectively, and the balloon material for the airship and balloon. The density is slightly higher for the fixed-wing and rotorcraft as these designs would require additional structural support in the wing or rotor; however the values for all four concepts represent light membrane structures, which are assumed to be necessary given Mars s extremely thin atmosphere. Sizing Study Results Using the above variables, each concept was sized according to the power required to maintain flight as determined by energy-based constraint analyses. A Monte Carlo simulation was used to change the input variables in such a way as to fill the design space, resulting in thousands of possible designs for each concept. This method allowed for rapid analysis of the entire design space, which would not be possible using a point-design approach. It was found that the power system density is the most constraining variable, demonstrating an upper limit of feasibility beyond which the vehicle s weight grows faster than its lifting capability. The other variables are less of an impact on feasibility, but still strongly influence the sizing and how the designs were ranked against the FOMs. The feasible results of the study are shown in the graphs in Figure 1. The scatterplot shows each design represented by its launch mass and 5.3 m constraint constraining dimension (wingspan for fixed-wing, 2.65 m constraint rotor diameter for rotorcraft, and tank length for airship and balloon). The results were filtered by the maximum dimensions that could fit in the aeroshell (5.3 m, allowing for one fold in wings or rotors), as well as the maximum allowable payload mass, 568 Figure 1. Results of Concept Selection Study 568 kg constraint 5

7 kg. As can be seen, the mass limit was not as constraining as the dimension limit. Figure 1 shows that the airship and balloon can easily be packaged in the aeroshell, while the fixed wing and rotorcraft concepts cannot fit without excessive folding. Qualitative Comparison of Concepts The above results suggest that while several concepts can accomplish the mission, a balloon is the most robust to the launch vehicle constraints. This is not surprising, as the balloon and airship primarily use flexible structures, which can easily be folded to fit a given dimension, while the other concepts require rigid members. A complete analysis of the concepts, however, must consider not only the resulting vehicle size, but also aspects of the vehicle related to capability, probability of success, feasibility, and control. These are difficult to predict to any degree of certainty considering that there is almost no historical data from past planetary air vehicles. To overcome this, the team chose to use a qualitative ranking Table 5. One of the Pugh Matrices for Concept Downselection. Fixed-wing Rotorcraft Airship Balloon Vehicle Complexity Probability of Success Control Feasibility Scientific Capability technique in the form of a Pugh matrix. 10 In this process the concepts were ranked by relative comparisons with each other, rather than by assigning each concept an absolute value in each metric. Multiple Pugh matrices were created, each one using a different concept as the baseline. One example is shown in Table 5. Here it can be seen that relative to the airship, a balloon offers a more simple design, but poorer control. A fixed-wing aircraft conversely is more complex, but offers greater control. However, a fixed-wing vehicle would have to be very marginal to fly in Mars s thin atmosphere, and as such has a lower probability of success and feasibility. After studying several such matrices, it became clear that a balloon would offer competitive science capability while being significantly simpler than the other concepts. In short, the added control offered by the more complex concepts is not believed to add significant scientific capability to the mission to justify their increased risk and cost. It must be reiterated that the mission is intended to be a long-term study of the Martian atmosphere, and Martian methane in particular. At this point in time, little is known about where to find methane, but computer simulations suggest it should be fairly evenly distributed throughout the atmosphere. 11 As such, the balloon s lack of a propulsion system is not seen as a detriment; allowing the balloon to simply drift with the wind will provide sufficient coverage of Mars over the mission lifetime. For all of the above reasons, the balloon concept was chosen as the best vehicle type to complete this mission. The selection of a balloon as the design concept is a major milestone that fully defines the design space to be considered in the remainder of the project. The rest of this report will focus on the detailed design of the balloon concept, often referred to by NASA as an aerobot when referring to planetary missions. This term will be used herein to distinguish the entire vehicle from the balloon envelope. Aerobot Design With the selection of an unpropelled aerobot as the best type of vehicle for this mission, sizing and synthesis of the appropriate subsystems had to be performed. As envisioned by the team, the aerobot will comprise two main components: a gondola and a balloon. The gondola carries all of the science instruments, computing and communications equipment, power sources, and other necessary subsystems. The balloon in turn provides a means of lifting the gondola and, through its interaction with the Martian winds, a means to transport the gondola. These relationships are shown in more detail in the functional decomposition in Figure 2. Here each function is color coded according to which subsystem it is dependent on, with grey functions being performed by an off-vehicle asset. Using this architecture, the design of the gondola is only loosely dependent on the balloon design. The gondola is designed with the science instruments as its payload, and each support subsystem in turn is sized to provide functionality for the science instruments. The fully designed gondola becomes the payload for the balloon, which is then sized to lift and transport the gondola. These interactions are depicted in the design structure matrix shown in Figure 3. Because the balloon is mostly independent of the intricacies of the gondola design, the design of the gondola and balloon will be described separately below. Baseline 6

8 Figure 2. Functional Decomposition for a Mission to Characterize Martian Methane Figure 3. Design Structure Matrix of Information Flow between Sizing Spreadsheets 7

9 Gondola Subsystems In order to correctly size the balloon, its payload, or gondola, needs to be sized first. In general, a successful vehicle will be capable of operating the science instruments, collecting their data, and storing this data until it can be relayed to a Mars orbiter. These operations require power, thermal protection, and a structure in which to mount the required hardware. Thus, the necessary subsystems are color-coded in the functional decomposition and identified as the science payload, power, command and data handling, communications, thermal, and structure subsystems. Tools were created to size each subsystem and integrated as seen in Figure 3; the resulting mass and power trends allowed appropriate hardware to be chosen for each subsystem. Science Payload Several methane detection devices and other instruments of interest, such as an atmospheric package and cameras, were identified and studied in detail. The science payload sizing tool calculated the mass and power required for hundreds of instrument combinations. It was decided to use two Quantum Cascade Laser (QCL) spectrometers and the atmospheric package, because this combination provided the minimum mass and power required and still fulfill all mission requirements. Each QCL spectrometer can be designed to perform different functions, and having the second spectrometer onboard provides redundancy and reduces risk. Each spectrometer requires approximately 3 watts of power to operate and weighs 0.25 kilograms. Both spectrometers will take a reading approximately every 12 seconds, and each reading will require 8192 bits of data, however by processing this data with an onboard computer each reading can be compressed to only 320 bits. 7 The atmospheric package will consist of a thermometer, pressure transducer, anemometer, and three accelerometers. The total package will require 0.3 watts of power and weigh 0.03 kilograms. Readings will be taken every second, and each reading will require 72 bits of storage (six instruments at 12 bits/instrument). Command and Data Handling The command and data handling (C&DH) subsystem collects and saves the science data until it is relayed to Earth. In addition, the C&DH subsystem records housekeeping data to be eventually sent to Earth and executes any commands sent from Earth to the vehicle. Such a system requires a central processing unit, data acquisition board, and flash memory, and the transit to Mars necessitates that such equipment be radiation hardened. A survey of potential radiation hardened computers resulted in the identification of the RAD6000, which has flown successfully on Mars Pathfinder and the Mars Exploration Rovers (MER). 12 The C&DH subsystem will implement the RAD6000 processor, manufactured by BAE Systems, because this processor s past success in Martian missions substantially lowers the mission risk. The RAD6000 is lightweight and has plenty of processing power with low power consumption. Weighing 0.9 kg, the particular RAD6000 configuration for the methane characterization mission requires less than 7.5 W maximum power and 3.1 W during periods of low power. 13 A back-up RAD6000 will provide redundancy. Both processors will be plugged into a backplane along with 256 MB of flash memory (which will provide several days worth of storage), and the data acquisition board. The entire subsystem requires an estimated 7.5 W and weighs 3 kg. Communication Subsystem The communications subsystem relays the data collected by the C&DH subsystem to Earth. The data can either be sent directly to the Deep Space Network (DSN) on Earth, or first relayed to a Mars orbiter, which will then transmit the data to the DSN when convenient. The direct approach allows for time sensitive data and commands to be transferred but requires a larger, more powerful antenna. Because the scientific instruments do not collect time sensitive data and the aerobot does not require commands from Earth in order to perform its daily operations, the vehicle can utilize a smaller, less massive communication system that will transmit to orbiters, such as Mars Odyssey, Mars Reconnaissance Orbiter, and other future orbiters. The orbiters will then send the data to the DSN. Compatibility with the ELECTRA UHF transceiver onboard the Mars Reconnaissance Orbiter and future orbiters will guarantee that the aerobot can transmit data once on Mars. CMC Cincinnati Electronics, manufactures ELECTRA-compatible UHF transceivers such as the C/TT-505, which was chosen for this mission. 14 This transceiver weighs 1.85 kilograms and requires 45.5 watts for transmitting and 6 watts for receiving. 15 The transmission and receiving frequencies will be in the 400 MHz range. When paired with a monopole antenna, the entire system weighs approximately 2.4 kilograms. A similar system utilizing a CMC Cincinnati UHF transceiver and monopole antenna operated successfully onboard MER. 16 The aerobot will be capable of transmitting for two to eight minutes per orbiter pass, at a rate up to 256 kbps, 16 and may experience up to three orbiter passes per day. To conserve power, the aerobot will only transmit once per day, requiring an average power of 0.72 watts. A single two-minute pass at 128 kbps will be sufficient for 8

10 transmitting multiple days worth of mission data. If the aerobot is unable to transmit for several passes, the C&DH subsystem s storage capacity is capable of saving the newer data without having to erase the existing data. Thermal Subsystem The thermal subsystem must be in place to protect the temperature-sensitive instruments and electronics from the cold nighttime Martian temperatures (as low as 170 K). In order to conserve power, passive methods of heating were chosen for the gondola subsystems. Reference 17 provides an equation for estimating thermal subsystem mass, resulting in a required mass of 0.23 kilograms. Power Subsystem The power subsystem generates, stores, and distributes the necessary power to the subsystems listed above. Several power generation and storage methods exist, and the decision of which method(s) to utilize was driven by mass. Due to the long endurance required by this mission, the power source should be renewable. Solar arrays and solar thermal dynamic cycles are the only options without an on-board fuel source, and solar arrays, being much lighter in mass than solar thermal dynamic cycles, are the optimal power source for this mission. Due to high efficiencies, the baseline solar array will consist of triple junction gallium arsenide (GaAs) cells; such cells have reached laboratory efficiencies of 35.2% and have been successfully used on past Martian missions. 18 This technology is expected to improve in the future, so it is assumed that the baseline solar array will consist of 30% efficient cells with a specific performance of 100 W/kg. Although most systems will be turned off at night to conserve power, secondary batteries will provide power to the atmospheric package and C&DH subsystem. The battery pack will consist of Lithium Ion cells, which exhibit energy densities as high as 160 W hrs/kg and do not suffer from memory. 19 The Li-Ion battery pack for this mission will provide 28 V and 3.5 A hrs. The power subsystem s control and distribution power requirements and masses were estimated using approximations from Reference 17. The power subsystem needs to generate an average of 77.4 W per diurnal cycle, and it weighs a total of 3.43 kg. The required solar array area, if mounted onto a cylindrical body, is 2.72 m 2. The power budget is provided in the gondola design summary. Structure Subsystem The structure subsystem provides a foundation in which to mount and protect the above subsystems. This supporting structure consists of the gondola, which hangs underneath the balloon. For purposes of this feasibility study, the gondola is assumed to be a simple cylinder. The structure subsystem mass is calculated to be 1.7 kg. The final dimensions of each gondola will be a radius of 0.2 m and a height of 1.1 m, resulting in a volume of m 3. The gondola s surface area is 1.38 m 2 (not including the top and bottom of the gondola), requiring that 1.34 m 2 of the solar array be deployed underneath the gondola, effectively lengthening the cylindrical structure after the entry and atmospheric deployment phases have taken place. The gondola dimensions were selected to fit inside the aeroshell, which will be discussed in the launch, entry, and deployment section of the report. The complete mass budget is provided in the gondola design summary. Gondola Design Summary The gondola s mass and power budgets are presented in Table 6. The science payload, C&DH, and communication subsystems combined volumes equal 6400 cubic centimeters. The total volume of the gondola is cubic centimeters, leaving an adequate cubic centimeters for the thermal, power, and structure subsystem. Table 6. Gondola Mass and Power Budgets Subsystem Mass (kg) Power (W) Science Instruments Command &Data Handling Communications (transmitting) Thermal Power Structure TOTAL Balloon Design While knowledgeable of the basic principles of balloon flight, the team was inexperienced with the challenges facing modern balloon designers and with what additional challenges operating on Mars might present. A literature search was employed to determine the design-driving trades and limitations that should be considered. Some results were expected, such as materials selection being driven by weight and gas retention concerns. Others were unexpected, such as the structural benefits of pumpkin-shaped balloons, or the need to regulate balloon temperature to prevent CO 2 deposition in Mars s atmosphere. Additionally, the team sought the advice of Dr. Henry 9

11 Cathey, chair of American Institute for Aeronautics and Astronautics Balloon Systems Technical Committee. Design variable descriptions are given below, along with the corresponding options or ranges and how the final values were chosen. Final specifications were arrived at by simultaneous trade studies across all design variables. Balloon Type Two distinct balloon architectures were identified, zero-pressure and superpressure. A zero-pressure balloon is open to the atmosphere at the bottom and relies on a thermal difference between the lifting gas and ambient conditions to reduce the internal density and create lift. Designs for zero-pressure balloons on Mars have been proposed that would use solar heating to keep the balloon aloft by day and descend by night. 20 Superpressure balloons are closed to the atmosphere and maintain a constant volume, with internal pressure variable but always greater than ambient pressure. In this way, altitude is held constant for the duration of the mission (unless ballast is carried for altitude control.) There are also combination balloons that use two envelopes, such as a superpressure core surrounded by a zero-pressure balloon. 21 This option was seen to be overly complex and not suited for Mars s extremely thin atmosphere, and thus was not considered for this study. The superpressure design was chosen in order to accommodate a long endurance mission and to remain at altitudes above 6 km to avoid most terrain. The varying altitude of a zero-pressure design was not seen to offer any advantages for the desired mission and would add unnecessary risk to the mission with each landing/liftoff cycle. Balloon Structure Until recently, superpressure balloons used for atmospheric research on earth were typically spherical, unreinforced film structures. As early as 1919 it was known that natural shape balloons, with an ovular profile in the shape of Euler s elastica, offer superior stress distribution compared to spherical balloons; however material and manufacturing limitations prevented their use. 22 Specifically, singular points exist at the poles of the balloon where meridional tension theoretically becomes infinite. 23 Only with the recent invention of ultra high-strength materials has this constraint been overcome by incorporating reinforcing meridional load tapes into the balloon structure. These divide the balloon into axisymmeteric gores, giving it a pumpkin-like appearance. The final milestone in reaching a modern balloon design was to add additional material between the load tapes, so that rather than stretch and bulge out when inflated, a bulge would be manufactured into each gore to reduce stress. This type of structure was demonstrated in flight as early as 1999 and will be used for this mission. Number of Balloons Based on estimates of balloon mass and pre-deployment volume, it was anticipated that a single entry vehicle could carry multiple balloons. This option offers significant risk mitigation by building redundancy into the system and allows for a wider coverage area on Mars. Each balloon is designed to float at a different altitude to provide a vertical distribution of the science instruments and to greatly reduce any chance of collision between the balloons. It was decided to plan for a series of balloons at 1 km altitude increments, the lowest being at 6 km to avoid most terrain. Limited by the launch vehicle mass and volume constraints, it was found that four balloons could be carried within the launch vehicle and aeroshell, at altitudes of 6 km, 7 km, 8 km, and 9 km. Load Tape Material Selection The most important factor in selecting a load tape material is tensile strength, which must be very lightweight as with all components on the balloon. In addition, the material must meet constraints for operation in Mars s atmosphere. The material should exhibit strong insensitivity to extended exposure to UV radiation, must not become brittle in temperatures as low as 170 K, and must not suffer from creep. Finally, the material must maintain its properties when braided into a rope and subjected to stress concentrations such as knots or splices. Candidate materials are shown in Table 7. Table 7. Load Tape Material Candidates Material Density (g/cc) Tenacity (N/tex) Ult. Strength (Mpa) Density (g/m) Ult. Load (kn) Kevlar Spectra Zylon Vectran As can be seen, Zylon has the best strength to weight ratio of the available materials. Much research has focused on using Zylon for this application, and it has been found that this material meets the additional constraints 10

12 on Mars. 24,25 While there has been some concern about Zylon s resistance to UV radiation, this can be overcome by treating the fibers or sewing them into a sleeve in the balloon envelope. It is suggested that for a 100-day mission, the fiber load should not exceed 60% of its ultimate strength to avoid creep and maintain an appropriate safety margin; this suggestion became a design requirement for load tape sizing. 26 Envelope Material Selection Envelope material selection presents the most significant challenge in balloon design. This material must be lightweight, as the balloon must lift itself before any payload can be considered. The material must also be impermeable to the lifting gas, such that at the end of the long-endurance mission the balloon still contains sufficient gas to stay aloft. Radiative properties must be considered to regulate internal temperature to prevent CO 2 deposition on the balloon, and to ensure that the material remains at a temperature above its brittleness point. The superpressure design requires a positive gauge pressure inside the balloon, and the envelope must support any stress not carried by the load tapes. Furthermore the material must be resistant to tearing, especially when considering the folding that will be required before deployment, and must exhibit excellent tear propagation resistance. Finally, the material must maintain its properties under the Martian UV radiation. At present, the most feasible option is to create a coextrusion of several materials, each serving a subset of the requirements, which is common in balloon construction. After exploring many candidate materials, a 3 ply coextrusion consisting of a 0.25 mil Vectra film bound to a light aramid (Kevlar) scrim with a thin layer of adhesive in between provides the best option. Vectra is a liquid crystal polymer (the same type of material used in the Vectran fiber in Table 7) that exhibits extraordinary gas retention. The aramid scrim provides resistance to tear propagation. Aramid was chosen over Zylon for this application because the need for UV protection on such a large area would offset Zylon s weight advantage. An additional layer of light adhesive film will be placed between the plies to promote adhesion between the dissimilar materials. The final material weight is estimated at 21 g/m2. This is significantly lighter than comparable materials that Figure 4. Film Density Trade Study have been proposed for Earth ultra-long-duration balloons (ULDBs); however the extremely thin atmosphere on Mars necessitates such a marginal material. Figure 4 demonstrates the change in total launch mass of the balloons as film density changes. The total vehicle mass is very sensitive to film mass, and so developing new material technologies to reduce the film mass while maintaining or improving resistance to tears, pin-holing, and radiation is an important topic for future research. Material Leakage Trade Study As mentioned, Vectra was chosen for its superior gas retention. Data provided by the manufacturer suggests leakage rates so low as to be almost immeasurable for a 100-day mission. In reality, the leakage from manufacturing imperfections or pinholes formed as a result of packaging will far out weigh the raw material leakage. A trade study was conducted to determine the sensitivity of mass loss to leakage. The results are shown in Figure 5. Figure 5. Leakage Trade Study Gore Structure Two variables related to gore structure were considered: the total number of gores, and the additional material built into each gore. Manufacturing a bulge into the gores reduces stresses in the balloon at the expense of additional weight. The quantity of material added was specified as the path length along the gore between two adjacent load tapes (s) divided by the linear distance between the two load tapes (c), as shown in Figure 6. A range of s/c from 1 to π/2 was examined, such that the upper limit corresponds to a semicircular gore. The s/c ratio was treated as longitudinally constant, and so the total material weight was approximated as the material weight of a nongored balloon multiplied by s/c. 11

13 The number of gores most strongly affects the stresses in the inflated balloon. By adding additional gores, the local radius of curvature of each gore is reduced, thus reducing the barrel (longitudinal) stress. However, balloons with too many gores have been observed to inflate improperly. To avoid inflation problems, an upper limit on the number of gores that may be used was imposed on the design, based on a relation from Calladine. 22 Figure 6. Gore Shape Structurally, the balloon benefits from using the maximum allowed number of gores; however, each gore adds additional seams that increase the likelihood of imperfections that will cause lifting gas leakage. The required total load tape length is also directly proportional to the number of gores, though this does not significantly affect the total balloon mass. The final design was chosen to have eighty gores, with s/c = 1.1, which gives a strength margin of 1.9 while only adding 10% material mass. Gondola Tether Design Zylon was chosen as the tether material for the same reasons it was chosen for the load tapes. The strength requirement was set at 5 kn, which is approximately equal to a 45 g force on a 11.4 kg gondola. This offers a significant factor of safety to allow for sufficient strength to withstand the stresses of deployment, as well as tether degradation due to long-term UV exposure. For this breaking strength, the tether mass was calculated as 2.08 g/m. The length of the tether connecting the gondola to the balloon has implications for dynamic oscillations and keeping the gondola-mounted solar cells out of the shadow of the balloon, with secondary influences on weight and deployment. Unfortunately these trades are not well defined, and while a case could be made for a longer or shorter tether, neither is very conclusive. Coupling this with the fact that any reasonable length tether will compose just a small fraction of the total balloon weight, and will otherwise have little quantifiable effects on the balloon system, a decision was made to simply scale the tether length with balloon diameter. Scaling was chosen such that the tether length equals twice the balloon diameter. This gives a reasonable estimate of tether weight in the absence of a detailed analysis, and is sufficient for this study s goal of demonstrating the feasibility of an aerobot concept. Lastly, it is noted that the tether will not be a single filament, but will comprise several ropes connecting axisymmetrically around the circumference of the balloon, so as to distribute the gondola load. As such, the total length of tether material required will exceed this separation length. The final tether mass of each balloon is estimated at 1.2 kg. Lifting Gas Selection Helium and hydrogen were considered as candidate lifting gasses. These two gasses differ in atomic mass by only 2 amu, and so in Mars s atmosphere (average molecular weight of 44 amu) the buoyancy benefit of hydrogen is not significant enough to be the primary decision driver. Rather, the primary trades governing the choice of gasses are the storage system mass and complexity, leakage rates, and safety. Hydrogen offers significantly lower leakage than helium, and as such requires a less complex storage system for the long trip from Earth to Mars. However there is a danger of explosion before exiting Earth s atmosphere. The exact mass savings of using hydrogen instead of helium is unknown, and because such systems are built to order, even estimating the mass is difficult. One source estimates the gas handling system mass for a particular Mars balloon to be 14 kg for hydrogen compared to 75 kg for helium. 27 Using these values, the mass of a helium containment system would be roughly equal to the mass of a single aerobot. For this reason, hydrogen was chosen for the current design. This value of 14 kg for the inflation system was applied to each aerobot, and does not include the mass of the hydrogen storage tanks described below. Storage Tank Design Hydrogen may be stored as a cryogenic liquid, a compressed gas, or trapped in a metal hydride compound. For this project it is assumed that the hydrogen is stored as a compressed gas, which represents the current best option considering the long storage time requirement. Storage tanks for the hydrogen were designed to be of the minimum mass and volume possible. Carbon fiber reinforced plastic tanks were chosen for their superior strength to weight ratio. The final design calls for each balloon to be inflated by six tanks, each pressurized to 30 MPa. The combined masses of the six tanks for each aerobot are later provided in Table 8. Gondola Mass Growth A trade study was conducted to determine the robustness of the design to gondola mass growth and reduction. These changes could have many causes, for example payload growth due to carrying additional instruments, or solar cell growth due to technology not maturing as quickly as anticipated. Conversely, an 12

14 unanticipated breakthrough could lead to a reduced gondola mass. The results are shown in Figure 7 and indicate a gondola mass margin of 25%. Aerobot Design Summary and Baseline Comparison A mass breakdown of the aerbots is shown in Table 8. A summary of the final balloon design is shown in Table 9, with specifications of past Mars balloon designs shown for comparison. The Mars 96 was a Russian design, while the MAP and MABS were NASA designs. These baselines provide a benchmark for Martian balloon missions; comparing the GIT SMART aerobot s specifications to the baseline Mars balloons indicates that the GIT SMART design is realistic. A representation of the aerobot is shown in Figure 8. Figure 7: Trade Study of Flight Mass vs. Gondola Mass Table 8. Aerobot Mass Summary Altitude Balloon Gas Tanks Inflation Gondola And Tether TOTAL 6k k k k TOTAL Table 9. Comparison of Final Aerobot Design with Baselines Planet Mars Mars Mars Mars System Mars '96 MAP MABS GIT SMART Shape Cylinder Sphere Sphere Natural Diameter (m) 13.2x x17/33x19/35x20/38x21 Volume (m 3 ) /10500/12700/15400 Payload (kg) Balloon Mass (kg) /61.6/69.9/79.7 Altitude (km) 2 to /7/8/9 Material Mylar Nylon-6 My/Kev/PE Vectra/Kev Area Density (g/m 2 ) Yield/Ult Str (N/m) 787/ /2790 NA/2426 NA/3000 Max Diff. Press (Pa) Stress Level (N/m) Technology Readiness Levels In choosing components for the aerobot design, preference was given to those with high Technology Readiness Levels (TRLs) whenever possible. This helps to mitigate risk and reduce cost, as these components have already been proven to some extent. The final design does, however, contain some revolutionary enabling technologies, which are understandably at a lower TRL. The superpressure balloon concept itself has been identified by NASA as an Figure 8. 3-D Aerobot Representation Table 10. Aerobot Component TRLs Subsystem Design Component TRL Balloon Superpressure Balloon Concept 5 30 Science Payload QCL Spectrometer Components: Cavity Ringdown Spectrometer Optical Cryocooler IR Detector Quantum Cascade Laser Atmospheric Package CDH C/TT-505 Transceiver 9 Power Li-Ion Batteries Multijunction GaAs Cells 8 4 Computer RAD6000 CPU

15 important technology for use on Mars, and work is being done to advance this to a TRL of 6 in anticipation of future aerobot Scout proposals. The science instruments have also been designed to use new technologies, which will greatly improve the quality of the methane measurements. A summary of the aerobot TRLs is shown in Table Mission Infrastructure In order for the aerobot to fly on Mars and accomplish its goals, this mission must fit into the U.S. space program s infrastructure. Analyzing the system of systems in which the aerobot must operate will conclude the feasibility study by addressing the aerobots launch, entry, and deployment system as well as cost and risk. Launch, Entry, and Deployment The launch system selected for this mission is a Delta II 7925 with a 2.9 m diameter fairing. 30 This launch vehicle can carry 1022 kg to LEO, the largest payload capacity of the launch vehicles provided for Mars Scout missions. The Delta II also has an excellent history of delivering payloads to Mars, making it the lowest risk option. Once in LEO, a transfer stage must be used to send the payload on a trajectory to Mars. This stage weighs approximately one quarter of the mass launched to LEO, or 205 kg. Since the mission specifies that the fall 2011 launch window must be used, a Type II Hohmann transfer departing Earth between October 28, 2011 and November 28, 2011 was selected, resulting in a required C 3 of 9.92 km 2 /s 2 and arrival at Mars during fall This launch window uses the minimum energy to get to Mars while also allowing room for launch delays. Upon arrival at Mars, the transfer stage will slow the payload for entry and then be jettisoned, as the payload directly enters the atmosphere on a ballistic trajectory. A Viking heritage aeroshell like those used for Pathfinder and MER will protect the aerobots during entry. The masses of the heat shield and backshell were estimated using regressions of mass and ballistic entry data from Pathfinder, MER, and the upcoming Phoenix missions, resulting in 67 kg for the heat shield and 82 kg for the backshell. Once entry is completed and the aeroshell is falling through the atmosphere, a disk-gap-band parachute deployment system similar to the systems used for Pathfinder and MER will slow the aeroshell and aerially deploy the aerobots. A parachute mass of 17 kg was estimated in the same way as the heat shield and backshell masses. The balloons will be inflated by the tanks of gas during descent, and the balloon buoyancy will further slow the descent until the balloons are completely inflated, at which point the tanks will detach from the balloon. Table 11 summarizes the weights for the launch, entry, and deployment systems. The final launch mass of 821 kg is well within the 1022 kg constraint of the Delta II 7925, leaving a mass margin of 201 kg. This is a sufficient mass margin to prove the feasibility of the concept and demonstrates the robustness of the design to requirements creep. Cost Analysis The cost analysis was performed using NASA JSC s Spacecraft/Vehicle Level Cost Model (SVLCM) and Advanced Missions Cost Model (AMCM), resulting in estimates for Design, Development, Test, and Evaluation (DDT&E) cost, Life Cycle Cost (LCC), and Production costs. 31 Launch costs were estimated to be $55 million FY06. 30,32 The Principal Investigator (PI) cost was estimated by subtracting the launch cost from the life cycle cost. All models and costs were adjusted to FY06 dollars using the NASA New Start Index Inflation Calculator. Table 12 shows a summary of the mission costs. The total cost to the Principal Investigator was estimated at $429 million FY06. This mission meets the Mars Scout cost constraint of $439 million FY06 (equivalent to $475 million Real Year). Table 11: Launch, Entry, and Deployment System Weights Component Weight (kg) Transfer Vehicle 205 Heat Shield 67 Backshell 82 Parachute 17 Mass Delivered to Mars 450 Total Mass 821 Table 12: Cost Summary Cost Type Cost ($mil FY06) Model DDT&E 163 SVLCM Production 56 SVLCM LCC 484 AMCM Launch 55 Ref. 10, 31 PI 429 PI = LCC - Launch Probability of Success and Risk Assessment The mission s most important FOMs, risk and probability of success, determine the general feasibility of the mission. Every decision made during the design process took these two metrics into account, and the object of each downselection was to select the option that provided the lowest risk and highest probability of success. The total mission probability of success can be calculated from the individual mission segment s probabilities of success, which are summarized in Table