MAE 4351 final report Benjamin Wagner

SENIOR DESIGN:MAE 4351 Project

Ref.: MAE 4350-001-2016Date: 03. mai. 2023Page: 1 of 54 PagesStatus: Report #1

Mars Flight Vehicle: The Daedalus - Performance

Signatures:Name: Signature: Dept.: Date:

Author: Benjamin Wagner Benjamin Wagner MAE 5/3/2023

Seen: Dr. Bernd Chudoba MAE

Summary:

As part of the requirements to complete the capstone design project for the University of Texas at Arlington, the Albatross Aeroworks design team has been tasked with designing a bush plane for flight on Mars. This challenging task will require the combination of design elements for high altitude aircraft on Earth, due to the thin atmosphere on Mars, and the ruggedness of aircraft like the Piper Cub to handle landing on the rugged and varied terrain. This particular report focuses on flight performance, in addition landing gear configuration and basic design, solar power options, battery selection and rocket selection are also addressed. A database of mission dates from a NASA database is included to allow for preliminary review of possible launch dates which do not exceed rocket capabilities. A literature review of applicable material is given in order to create a knowledge base for the analysis that will come in future reports. An MDA and IDA are also included to cover data flow. In addition, a methodology for determining performance characteristics and usable solar energy is also addressed in detail. A brief history of Mars exploration and characteristics of concern are detailed to develop a reliable knowledge base from which a successful design can be derived. The final result yielded a takeoff ground roll of 493 m on the sand. A range of 183.7 km was found from a mission profile analysis. Two possible mission launch dates were also determined between 2025 and 2030 which met performance limits of the Delta IV Heavy rocket.

Distribution:

Institution: Dept.: Name:

The University of Texas at Arlington MAE Dr. Bernd Chudoba

The University of Texas at Arlington



Work Disclosure Statement

The work I performed to document the results presented in this report was performed by me, or it is otherwise acknowledged.

Date: 5/3/2023

Signature: Benjamin Wagner




Contents

WORK DISCLOSURE STATEMENT......................................................................................................................... 2

LIST OF FIGURES.................................................................................................................................................. 5

LIST OF TABLES................................................................................................................................................... 6

NOMENCLATURE................................................................................................................................................ 7

I. INTRODUCTION....................................................................................................................................... 8

A. ALBATROSS AEROWORKS........................................................................................................................................8B. PROPOSAL.........................................................................................................................................................10C. FLIGHT PERFORMANCE.........................................................................................................................................11D. SCHEDULE OF EVENTS..........................................................................................................................................12

II. MARS.................................................................................................................................................... 12

A. MARS EXPLORATION...........................................................................................................................................12B. MARS CHARACTERISTICS......................................................................................................................................12C. MARS TERRAIN..................................................................................................................................................13

III. LITERATURE REVIEW.............................................................................................................................. 15

A. PERFORMANCE REFERENCES.................................................................................................................................15B. SOLAR POWER REFERENCES..................................................................................................................................15C. DESIGN DECISION REFERENCES..............................................................................................................................15

IV. ENGINEERING METHODS........................................................................................................................ 16

D. MULTI-DISCIPLINARY ANALYSIS.............................................................................................................................16E. FLIGHT PERFORMANCE ANALYSIS...........................................................................................................................17F. MISSION PROFILE...............................................................................................................................................19

V. FLIGHT PERFORMANCE METHODOLOGY.................................................................................................20

A. PRIMARY FLIGHT PERFORMANCE PARAMETERS.........................................................................................................20B. DETERMINATION OF FLIGHT PERFORMANCE PARAMETERS..........................................................................................20C. MARS SOLAR POWER MODEL...............................................................................................................................22

VI. DESIGN DECISIONS................................................................................................................................. 24

A. LANDING GEAR SELECTION...................................................................................................................................24B. BATTERY SELECTION............................................................................................................................................26C. SOLAR CELL SELECTION........................................................................................................................................27D. ROCKET SELECTION.............................................................................................................................................28E. MISSION DATES..................................................................................................................................................31

VII. RESULTS................................................................................................................................................. 31

A. INPUTS.............................................................................................................................................................32B. PERFORMANCE VALUES........................................................................................................................................33C. SOLAR POWER ASSESSMENT.................................................................................................................................34D. MISSION PROFILE...............................................................................................................................................35E. ROCKET ANALYSIS...............................................................................................................................................37

VIII. ABET REQUIREMENTS............................................................................................................................ 38




A. OUTCOME C: DESIGN SYSTEM, COMPONENT OR PROCESS TO MEET NEEDS..................................................................38B. OUTCOME D: AN ABILITY TO FUNCTION ON MULTIDISCIPLINARY TEAMS.......................................................................39C. OUTCOME E: UNDERSTAND PROFESSIONAL AND ETHICAL RESPONSIBILITY.....................................................................39D. OUTCOME F: AN ABILITY TO COMMUNICATE EFFECTIVELY..........................................................................................40E. OUTCOME H: UNDERSTAND IMPACT OF ENG. SOLUTIONS IN GLOBAL AND SOCIETAL CONTEXT.........................................41F. OUTCOME I: RECOGNIZE THE NEED AND ABILITY TO ENGAGE IN LIFELONG LEARNING......................................................41

IX. CONCLUSION......................................................................................................................................... 42

APPENDIX......................................................................................................................................................... 43

REFERENCES...................................................................................................................................................... 53




List of FiguresFigure I-1: Albatross Aeroworks Organizational Structure……………………………………………………………….……..9Figure I-2: The Piper Cub, an Iconic Bush Plane………………………………………………………………………………..10Figure II-1: Pathfinder Terrain Image……………………………………………………………………………..…………….13Figure II-2: Damage to Curiosity’s Wheel……………………………………………………………………………………….13Figure II-3: Dune Fields on Mars-Equatorial Perspective………………………………………………………………………14Figure II-4: Dune Fields on Mars’ Southern Hemisphere………………………………………………………………………14Figure IV-1: Multi-Disciplinary Analysis……………….…………………………………………………………………..........17 Figure IV-2: Flight Performance Analysis Methodology……………………………………………………………….……….18 Figure III-4: Global Hawk Mission Profile…………………………………………………………………………………...….19Figure III-5: Proposed Mission Profile……………………………………………………………………...……………………19Figure V1-1: Ski Landing Gear…………………………………………………………………………………….……………..25Figure VI-2: Geometric Estimation of Landing Gear Positions………………………………………………………………..26 Figure VI-3: Delta IV Heavy Payload Fairings…………………………………………………………………………………29 Figure VI-4: Standard Fairing vs Asymmetric Fairing………………………………………………………………………….30 Figure VI-4: IRVE Reentry System…………..…………………………………………………………………………………..30 Figure VII-1: Thrust Available by velocity………………………………………………………………………………………32 Figure VII-2: Solar Irradiance at 0.2 Optical Depth on Mars Equator………………………………………………………..34 Figure VII-3: Solar Irradiance at 0.2 Optical Depth at 30 Degree South Latitude……………………………………………35 Figure VII-4: Days per 30 Degrees True Anomaly…………………………………………………………………………...….35 Figure VII-5: Range Illustrated from UTA Campus……………………………………………………………………………36




List of TablesTable I-1: Project Schedule………………………………………………………………………………………...………...…...12Table VI-1: Battery Data……………………………………………………………………..……………………………..…….26Table VI-2: Oxis Energy Ultra-Light Lithium Sulfur Pouch……………………………………..……………………….……26Table VI-3: Rocket Database…………………………………………………………………………………...…………………28Table VI-4: Martian vs Earth Calendar Reference……………………………………………………………….……………30 Table VII-1: Propulsion Data Table 1.…………………………………………………………………………………………...32 Table VII-1: Propulsion Data Table 2…………………………………………………………………………………………….32Table VII-2: Aerodynamics Outputs……………………………………………………………………………………………..33Table VII-3: Sand Takeoff Ground Roll and Time………………………………………………………………………………33Table VII-4: Bedrock Takeoff Ground Roll and Time………………………………………………………………………….33Table VII-5: Half-Mission Profile Data…………………………………………………………………………………………..36 Table VII-6: Mission Profile Based Range……………………………………………………………………………………….36 Table VII-7: Delta IV Maximum Performance……………………………………………………………………………….…37 Table VII-8: Mars Southern Summer Missions with Returns………………………………………………………………….37Table VII-9: Delta-V Qualifying Mars Missions…………………………………………………………………………………37Table A-1: Normalized Net Irradiation for Surface Albedo of 0.1……………………………………………………………..43Table A-2: Normalized Net Irradiation for Surface Albedo of 0.4……………………………………………………….….....44Table A-3: Normalized Net Irradiation for Surface Albedo of 0.25……………………………………………………………45Table A-4: Mars Bulk Parameters………………………………………………………………………………………………..46Table A-5: Mars Orbital Parameters……………………………………………………………………………………………..46Table A-6: Mars Observational Parameters……………………………………………………………………………………..47Table A-7: Martian Atmospheric Characteristics………………………………………………………………………….……47Table A-8: Historical Log of Mars Missions……………………………………………………………………….…………….47Table A-9: One-Way Mars Missions…………………………………………………………………………………………..…50Table A-10: Mars Missions with Return Trajectories…………………………………………………………………………..51Table A-11: Dune Field Distribution by Latitude……………………………………………………………………………….52Table A-12: Dune Field Distribution in Craters…………………………………………………………………………………52




Nomenclaturea -Average AccelerationA -AlbedoAR -Aspect RatioAU -Astronomical UnitC -Specific Fuel ConsumptionCI -Coefficient of Solar IrradianceCL -Coefficient of LiftCD -Coefficient of DragCD0 -Zero-Lift Coefficient of DragD -Drage -Orbital EccentricityE -EnduranceFAA -Federal Aviation AdministrationFAR -Federal Aviation RegulationFBO -Fixed Base of OperationsG -Gravityh -Hour AngleH -AltitudeI -Solar IrradianceIG -Ground Level IrradianceI0 -Average Solar IrradianceIsp -Specific ImpulseIDA -Individual Disciplinary AnalysisIRVE -Inflatable Reentry Vehiclekbat -Gravimetric Energy DensityL -LiftM -Mean AnomalyM0 -Starting MassMf -Final MassMDA -Multi-Disciplinary Analysis

MTV -Mars Transfer VehicleP -Length of Mars DayPAvail -Power AvailablePElec -Electrical Powerr -Current Sun Mars Distancer -Average Sun-Mars DistanceR -RangeRT -Turning RateROC -Rate of ClimbS -Planform AreaSG -Ground RollT -ThrustV -VelocityVSTALL -Stall VelocityVTD- -Touchdown VelocityV/P -Velocity-Power RatioWi -Initial WeightWf -Final Weightz -Solar Zenith Angleδ -Solar Declinationε -Mars Axial TiltηBC -Solar Cell Efficiencyηdischarge -Battery Discharge EfficiencyηSC -Solar Cell Efficiencyθ -LatitudeθT -True Anomalyρ -Air Densityψ -Maximum Solar Inclinationμr -Coeff. of Friction for Ground Operationsω -Turning Rate




I. IntroductionHE purpose of this project is to fulfill requirements of the capstone design project for the MAE 4351 class “Aerospace Vehicle Design II” at the University of Texas at Arlington. The assigned project is the engineering

of a multi-purpose aircraft for flight on Mars. The aircraft must be capable of flying both manned and unmanned. It must be able to land on and takeoff from a significant portion of the terrain on Mars. The aircraft must be safe and offer sufficient reusability to make it a worthwhile addition to a Mars mission. The following is a synopsis of requirements and recommendations given by Dr. Chudoba regarding the design of the aircraft1:

T

prove manned flight is possible optionally piloted platform (manned/unmanned) first deploy autonomous platform followed by manned utilization identify customer-base for autonomous platform demonstrate practical self-take-off/landing identify practical & robust propulsion system demonstrate astronaut & aircraft interface incorporate GA rescue system explore least-effort transit/deployment concept investigate survivability/longevity in Martian environment incorporate aerial platform into promising commercial (e.g. SpaceX) or research (e.g. NASA) plan

This particular report focuses on flight performance, a discipline which quantifies a wide range of aircraft flight characteristics. These will include concepts such as range, endurance, flight ceilings and rate-of-climb among others. Success on this mission is not measured individually by any single benchmark, but from a more practical perspective of the overall value the aircraft will present to a Mars mission.

In addition to performance characteristics, other design aspects will be considered. Solar power, an option for supplying electrical power to run the aircraft will be modeled in order to determine its feasibility. If sufficient, the aircraft may be powered or charged using solar power. If not, an alternate power source must be utilized. Battery and solar cell selection, a necessary partner to solar power, is also addressed in this report. The selection of a high specific density battery which can operate in the cold weather conditions on Mars is necessary for the success of this mission.

Landing gear and suspension systems have also been designated to the performance group. While this is primarily a structural consideration, the selection of a landing gear system and appropriate suspension have a direct effect on takeoff and landing considerations and will be selected with performance consideration as primary.

Rocket selection has also been made a responsibility of the performance group. While this is non-standard and not related to flight performance, performance characteristics of the rockets themselves must be determined. The rockets must be capable of launching the MTV and its payload into a 200 km circular LEO. This is the orbit from which Mars transfer trajectories are calculated by NASA, and serves to simplify this first order approximation. The estimates from this analysis may include a significant margin of error and should only be used as a starting point for any major analysis. The findings will then be referenced to a database of potential Mars launch dates to produce a short list of mission timeframes.

A. Albatross AeroworksAlbatross Aeroworks was formed as one of two teams in the class; an eight person team which was divided into

separate disciplines according to the team member’s skillset and the team needs. The members and their positions are as follows:

Chief Engineer- Tim Mushrush Aerodynamics- Zack Ferguson Structures- Enrique Rodriguez Propulsions- Gabe Lopez Controls/Stability- Guerric Makengo Performance- Ben Wagner CAD- Kelly Wright Cost Analysis- Zac Webster




Figure I-1: Albatross Aeroworks Organizational Structure2

Each member will handle most of the analysis of their system as shown, but for any multi-system design there will be overlap and collaboration in many decisions and calculations. The chief engineer is Timothy Mushrush and will handle synthesis where the overall aircraft will be “meshed” together and steered toward certain design goals. This could be due to cost, weight, mission goal, propulsion availability.

Aircraft are designed to complete specific mission goals and perform in a specific way. Ben Wagner is responsible for the aircraft performance. Aircraft performance analysis treats the plane as a point mass and uses fundamental laws of physics to formulate equations that determine flight characteristics such as range, endurance, climbing rate, takeoff, and landing distance.

For this, all the other systems must work in unison. Gabriel Lopez is responsible for propulsion of the aircraft. He will perform an analysis on electric motor driven propeller propulsion systems versus rocket propulsion, with an emphasis on the Mars Daedalus required needs. Suitable electric motors that were in production during the time of development of the aircraft will be compared. With this an engine performance model based on an optimal propeller configuration to meet the requirements of the aircraft mission will be conceptualized.

Being a solar aircraft, the solar cells are an important part of not only energy generation to generate thrust for propulsion, but need to be considered for structures. Enrique Rodriguez leads the design of structures. This system will require an analysis of weights, inertias, materials, and the structures themselves. Analyzing the weights that span throughout the aircraft to identify the center of gravity, which will create a reference axis system. Inertias of the aircraft, the structural geometry and materials will all combine as well. These are of upmost importance for all other systems.

The CAD/geometry system that is led by Kelly Wright will focus on the dimensions of the aircraft as well as 3-D design. The findings of this group lay the foundation for calculations needed in other groups such as aerodynamics, structures, and controls.

Guerric Makengo leads the Controls/Stability group. This system develops the analysis of all the elements or features such as the empennage that keeps the aircraft stable and reliable in its mission. It verifies that the sum of all the moments must equal zero for an aircraft to be in equilibrium. It determines the trim conditions based on the inputs such as weight, control surface dimensions, and lift & drag coefficient terms from other groups.

The lift & drag are determined by the aerodynamics group which is led by Zack Ferguson. Aerodynamics uses 2-D analysis initially to find lift and moment of the aircraft based off of lifting surfaces. Then a 3-D finite wing analysis is done. The overall drag of the aircraft is also found.

Zac Webster is in charge of cost analysis. This is the quantification and integration of lifecycle, operation, development and manufacturing cost. It also entails analyzing revenue and valuation techniques. This means that working closely with structures & CAD/geometry to keep track of materials and understanding operation costs associated with the performance group will be critical as well.




B. ProposalThe first requirement for the teams was to assemble proposals for presentation to Dr. Chudoba. These

presentations were to consist of ideas for a business case for a Mars aircraft, and to supplement the original requirements. The teams were given one week to generate the proposals which were due on June 15th.

Early proposal ideas included a lighter-than-air design and vertical take-off. The airship idea was quickly determined to be unfeasible for manned operation as the thin Mars air would require a massive rigid airship. The approximation found that for every kilogram being lifted a vacuum of greater than 50 m3 would be required. Thus lifting an astronaut an equipment of approximately 150 kg would require a rigid airship of at least 7500 m 3. This value doesn’t account for any structure of the airship, and assumes a vacuum inside, both obviously impractical. A VTOL aircraft was found to require excessive power, yielding any design to have a very short flight duration and ineffective range. After these results the team determined the more efficient fixed-wing STOL design was necessary.

Albatross Aeroworks developed two primary proposals. The first was a supersonic aircraft to be released during an orbital entry to do reconnaissance of possible locations for a manned mission to Mars. The aircraft would be unmanned and for a single use only.

The second proposal was a combination of UAV and manned aircraft. The UAV was to deliver rovers and automatic testing equipment to designated locations, and the manned craft was for collecting samples and performing repairs on remote equipment. Both aircraft would be very similar in design, differing primarily in sizing considerations. They would both be minimalistic and designed for rugged terrain.

The competing team proposed a sample collection system. The aircraft would be delivered from orbit, enter the Martian atmosphere and land. It would then collect a sample, which would be launched back to orbit via rocket to rendezvous with the spacecraft for return to Earth.

In the end, certain aspects of a bush plane, an aircraft deigned for operations in rugged landscapes, were incorporated into the requirements. The aircraft is required to takeoff from and land on the Martian surface, it must also transport an astronaut and all of his associated gear. The final mission, while very similar to the original has given the team a clear direction to begin the associated work.

Figure I-2: The Piper Cub, an Iconic Bush Plane

The business case of this mission has multiple aspects. It is both of public and scientific interest as well as a proof of concept for private gain. This makes the case for including an aircraft for a Mars mission of high interest to many.

Public interest in space exploration tends to waver depending on the NASA as firsts were accomplished time and again. Public interest in the space program has greatly decreased in recent years as low orbit missions have become pedestrian in comparison Neil Armstrong’s first steps on the moon. However, the departure of the first manned mission to Mars would most certainly provoke massive public interest on a feverish level. In order to further peak that interest, the inclusion of the first man to fly on a foreign planet on one of those pioneering missions will certainly help to maintain that interest. Many proposals for Mars aircraft have been put forward, but none have sought to tackle the challenge of carrying an astronaut. This would bring the man back to the forefront of exploration, eliminate the drone and the rover and allow humans to leave footprints in locations never seen by the naked eye.

Scientifically, the astronaut could do more and collect samples and up close images from a variety of locations beyond the capabilities of a simple rover. More of the planet could be seen in minutes than curiosity traveled in its first 2 years. The astronaut can be directly involved in collecting samples of interest from sights which may span 100 km or more. The value to




science of having varied sample rather than a few rocks found meters apart is invaluable to increasing our understanding, or perhaps even finding proof of life.

As a proof of concept, the Mars Daedalus serves to show flight is possible, even in the thin air and landing in the rough terrain can be done. Once astronauts on Mars become a regularity, rather than a dream, the need to repair equipment may present itself. Nations spend hundreds of millions, if not billions of dollars to put rovers and other scientific equipment on Mars, rather than losing the equipment to the first mechanical failure, it can now be repaired. The astronaut would fly from his current location with a 3d printed replacement part for the machine and perform repairs on sight. While the expense of this may be several million dollars, saving a multi-billion dollar investment is well worth the price.

The Mars Daedalus is a project to do more than simply show that manned flight on Mars is possible. It will generate public interest in science, open up doors for faster and broader exploration, and show us we can breathe new life into projects that would once have been lost.

C. Flight PerformanceDesigning an aircraft for flight on Mars presents a slew of challenges. These challenges affect every group

involved with the aircraft design process. Mars’ thin atmosphere reduces available lift, the remoteness of another planet create challenges in transporting and deploying the aircraft. Every aspect of this aircraft’s design must be tailored to handle the challenges presented.

For the performance group, four design decisions have been assigned. The first of which is landing gear system selection. Due to the rugged terrain the aircraft will require rugged and flexible landing gear. The fine sand mandates a larger tire to reduce the rolling friction coefficient, but the thin air calls for light weight landing gear design. A review of landing gear designs will be performed. The advantages and disadvantages of each will be compared in order to make an appropriate selection.

The propulsion system has yet to be selected, so work for both rocket and electrical power propulsion are required. To this end, battery selection was the next assigned design task. The temperatures on Mars are very cold and wide in range, a battery will need to operate in cold climate conditions. The lightest possible battery suitable for the job is also necessary, this means a high specific energy. A review of battery types and a search for industry information will be conducted

In order to charge the batteries for the aircraft, solar power is the most appropriate. The options are to use this system to charge the aircraft batteries, or to install solar cells on the aircraft skin for direct power. A charging system can be left at the landing site, or be brought along. The selection of different solar cells to meet the needs is necessary.For the direct power supply, the lightest and most efficient system is a must. The same cells would be appropriate for a charging system that travels with the aircraft. For a stationary charging system, weight is less of a concern, so the highest efficiency is desired to reduce charging times and increase the usability of the aircraft. A review of industry information will be performed to find solar cells for this purpose.

The standard flight performance considerations are complicated by Mars conditions. An atmospheric model must be built to determine properties like density and pressure. Takeoff and landing performance are a major focus of the project, the aircraft cannot succeed in its design purpose if unable to safely perform ground operations. A slow takeoff and landing speed are preferred to increase the safety, and sufficient power is needed to takeoff with high rolling coefficients in sand.

In order to redistribute workloads based on team progress, rocket selection has been reassigned to the performance team. This task is not a major design consideration, but a necessary one. The design of an aircraft that ignores the limitations of weight and size will very possibly exceed what can be transported to Mars. A survey of current rockets will be conducted in order to determine possible solutions. In addition, the size of the rockets will be addressed and compared with a selected packing method (selected by synthesis). The most appropriate will be selected at that time. If none of the rockets are suitable or capable the design will be sent back to the synthesis team.

The data from the rocket selection process will include determination of a maximum delta-V, the specific impulse required for the mission, for the MTV. This value will be compared to a database of Mars missions, and cross referenced to missions with return trajectories to determine possible launch dates between 2020 and 2030. Should the Mars Daedalus utilize solar powered propulsion, the dates will also be referenced to determine that any stay on Mars will cover a significant time in which enough solar irradiance is available at ground level. This process will allow Albatross Aeroworks to narrow down to a few select launch opportunities over the given time frame.

To accomplish the required tasks in a simple and transferable method, Microsoft Excel will be used. Excel allows for the establishment of a series of algebraic equations in order to calculate values under a range of




conditions, or to accomplish a series of calculations quickly. The outputs can be organized into tables and spreadsheets can be transferred to other members of the team electronically.

D. Schedule of EventsAs in any industrial endeavor, a schedule is required to ensure timely delivery of materials to avoid costly delays

and customer issues among many things. In this particular case bi-weekly reports are required as well as two presentations, one mid-term and one final. Included is a table of due dates running through the semester:

Table I-1: Project Schedule

Week Deliverable20-Jun Report 14-Jul Report 218-Jul Report 31-Aug Report 4

14-Aug Final Report

In addition to the listed schedule, two presentations will occur during this time. The first of which will be during the 25-July and the final presentation 8-Aug.

II. MarsThis section includes a brief history of Mars missions. In addition, characteristics of Mars pertaining to the

mission are defined. A discussion of Mars terrain is included as it relates to areas in which the aircraft can operate.

A. Mars ExplorationThe exploration of Mars has been ongoing for decades, but still no man has set foot on the planet. The first

attempt to reach Mars was a launch in 1960 that failed to reach Earth orbit, the Russian Korabl 4 and a series of failed mission by Russia and the US followed. However, in 1964 NASA’s Mariner 4 became the first spacecraft to reach Mars. Dozens of missions have been launched since, many resulting in failure. One of the most successful missions to date has been Curiosity, a motor-vehicle sized rover weighing in at about 900 kg. Curiosity has been conducted an exploration of the Gale Crater for nearly 4 years. Table A-7 in the appendix contains a historical list of missions to Mars.

B. Mars CharacteristicsA collection of data from reputable sources is one of the first requirements for any kind of research. While this is

discussed in the next section on how this pertains to the performance discipline and other engineering related tasks assigned to the flight performance team. This section covers data collected on Mars and a brief history of missions to Mars.

The most complete single source for data on Mars would be from NASA. In the appendix several tables are included which contain data obtained from publicly accessible sources posted by NASA. The following data is some of the more relevant13:

Surface Gravity: 3.71 m/ss

Orbital Period: 686.973 days Perihelion: 206.62*106 km Aphelion: 249.23*106 km Length of Day: 24.658 hours Axial Tilt: 25.19 degrees Surface Pressure: 6.36 mb at mean radius Surface Density: .020 kg/m3

Average Temp: 210 K




It also valuable to consider the primary make-up of the Martian atmosphere. Below is a list of the primary atmospheric constituents13:

Carbon Dioxide (CO2) 95.32% Nitrogen (N2) 2.7% Argon (Ar) 1.6% Oxygen (O2) 0.13% Carbon Monoxide (CO) 0.08%

The data listed in the tables, some of which is included above, will be the primary source for data on Mars. These values are utilized in producing solar models as well as atmospheric data essential to performance calculations.

C. Mars TerrainThe Martian terrain is also a serious concern.

While some areas are rolling hills and sand dunes, others have jagged rocks and even dense heavy boulders. Obviously, these areas are unsuited for landing an aircraft. It is necessary to assess the recurrence of usable and unusable terrain to understanding the feasibility of this project. Extensive reading of articles and research from reliable sources was conducted to collect data on the Martian terrain and better understand the challenges.

The accompanying image was taken by Pathfinder, a Mars rover launched by NASA. While the background reveals a large plane with what appear to be two small mountains in the background, the foreground is rocky and unforgiving. Terrain of this nature is probably unsuitable for standard takeoff vehicle, and has caused damage to aluminum wheels on Curiosity.

As described previously, the Martian terrain can be unforgiving. Shown here is damage to one of Curiosity’s wheels caused by razor sharp rocks on the Mars surface. In order for a vehicle to operate on this terrain without extensive damage and need for replacement wheels it would have to operate at very slow speeds, or utilize extremely heavy-duty tires. Even with heavy tires the rough terrain would cause major stability issues for any fast moving aircraft during takeoff and landing. Additionally, heavy duty tires would mandate much higher weights. Operating in this rocky terrain will require the design of tires and suspension to guarantee safe operation. These issues will be addressed below in the landing gear design section and stability testing will be the responsibility of the controls and stability group.

Alternately, Mars also has large sandy areas with rolling hills and flat bedrock. The flat bedrock areas would be relatively safe for the aircraft to land, but determining their prevalence on the Mars surface is difficult due to the small percentage of Mars which has been surveyed in this level of detail. However, a significant percentage of dune fields have been mapped.

Data from the Astrogeology Science Center, a division of the USGS (United States Geological Survey), includes maps of dune fields and classification of types. This system is called the Mars Global Digital Dune Database (MGD3). This information can be utilized to determine safe landing areas. Landing in relatively flat areas among the sand dunes gives the advantage of a softer and smoother area to put down. Adversely, the coefficient of rolling friction is increased in these areas and power required to takeoff is increased. This value will be assumed to be 0.4 considering the fineness of the sand on


Figure II-1: Pathfinder Terrain Image17

Figure II-2: Damage to Curiosity’s Wheel17



Mars18. From the USGS findings, many of the sand dunes are located in craters. The bowl shape of a crater gives a shield from wind and allows sand to collect inside. While the wind still shapes the dunes, the sand tends to remain.

MGD3 maps have been completed between the latitudes of 65° N and 65° S and some mapping has been performed beyond these limits, but is not complete. Maps are included below marking the position of dune fields as well as tables in the appendix detailing the total area and distribution by longitude. The yellow areas on the map represent dune fields inside of craters, while the red are dune fields not in craters.

Figure II-3: Dune Fields on Mars-Equatorial Perspective20

It is apparent from figure II-3 that the southern hemisphere has the vast majority of the dune fields on Mars. The utilization of solar power for this mission would make the southern summer a necessary time for flight activities to ensure sufficient solar power. While the southern summer is shorter than the northern it also is closer to the sun, giving higher maximum solar irradiance.




Figure II-4: Dune Fields on Mars’ Southern Hemisphere20

III. Literature ReviewA review of sources for this project was performed during the preparation of the proposal. This section presents

the primary sources for information used for developing methodology and deriving valuable data. It is separated and organized into the different design tasks assigned to the performance group.

A. Performance ReferencesDeveloping a knowledge base is crucial to creating a valid methodology for design of an aircraft. In order to

model flight performance characteristics, several sources ae referenced: Introduction to Aircraft Flight Mechanics by Thomas Yechout, a text book from the AIAA education

series, contains basic flight performance analysis equations and explanations of their applicability Fundamentals of Aircraft and Airship design by Leland Nicolai, also from the AIAA education series, this

text book expands on performance concepts providing additional information that may prove essential Reverse Engineering the Solar Impulse, by Benjamin Wagner, capstone design paper from spring semester,

many of the current design requirements are addressed in the previous project

These sources offer standardized performance equations. The AIAA publications contain information with simple algebraic analysis of the performance characteristics of interest on most aircraft. The Solar Impulse report contains information pertaining calculation of endurance and range for electrical and solar powered aircraft. The methodology development requires significant information from these sources.

B. Solar Power ReferencesThe modeling of solar power on Mars will aid in the determination of its feasibility as a power source to directly

power flight, or as a recharging source to permit battery powered, electrical flight. In order to model solar power and electrical data, the following sources will be used:




NASA Technical Memorandum 102299: Solar Radiation on Mars, a NASA technical document dealing with solar irradiance on Mars

Design of Solar Powered Airplanes for Continuous Flight by Andre Noth, an academic paper addressing the design of solar powered aircraft

Fast and Accurate Estimation of Solar Irradiance on Martian Slopes, a technical article by Asymeric Spiga and Francois Forget

Atmospheric Effects on the Utility of Solar Power on Mars, by Robert Haberle: a research article on Mars solar irradiance and its diffusion in the Mars atmosphere

Building a solar energy model for Mars requires information on the methodology to employ it for powering aircraft, this can be found in the paper by Andre Noth. The NASA technical papers give data for Mars and the necessary information to build a model based on reasonable assumptions and measured phenomena.

C. Design Decision ReferencesMultiple design decisions have been made the responsibility of the performance group for this project. In order

to aid in the process of selecting materials for batteries and solar cells, as well as landing gear and suspension configuration, the following sources will be utilized:

Handbook of Batteries Vol. 3, by David Linden, a comprehensive manual on batteries Aviation Maintenance Technician Handbook-Airframe Vol.2, reference for existing landing gear and

suspension design Aircraft Landing Gear Design: Principles and Practices, by Norman Currey, part of the AIAA educational

series

Industry data will be researched and utilized to aid in the design process. Many design decisions made on aircraft are directly related to finding a manufacturer with a current product that will meet the needs of the project. The selection of solar cells and batteries will be performed by exploring existing options. The AMT handbook is used as a reference to build a multiplicity of options for landing gear design. While basic design analysis of the landing gear will be addressed reviewing the AIAA publication.

IV. Engineering MethodsThe process of reverse engineering an aircraft is complex. It requires application of several engineering

disciplines and often many iterations of the overall process to complete. Each individual discipline is responsible for crucial steps of the analysis, and must work together efficiently to complete the task, as a problematic error may require re-worked data from several sources in order to correct.

D. Multi-Disciplinary AnalysisDesigning an aerial vehicle for flight on Mars is a complicated if not daunting task. This undertaking requires

organized methodology and teamwork. In order to do so a multi-disciplinary process must be formulated to understand and help regulate the flow of information between design groups.

The design process is demonstrated in the MDA by illustrating the flow of information from one group to another during the engineering process. The process begins by establishing a knowledge base, which includes information available about ultra-light aircraft and Mars. This knowledge base is then considered with the mission requirements as well as pertinent regulations in order to understand the characteristics and requirements for a successful design. The information is then given to the teams who complete their individual design processes and deliver the data back to the synthesis team for validation. The synthesis team then checks all data to ensure cohesion. After verifying, the data is compared it to a solution space which was generated earlier in the engineering process. This final step determines whether or not the reverse engineering process generated a successful design.

The MDA for this project highlights the usage of an alternate sizing methodology. While the Loftin sizing method can be used for aircraft flying subsonic on Earth, it is not usable for aircraft on Mars. However, an alternate sizing methodology for Mars allows for an aircraft designed for Earth to be resized for the conditions on Mars. This process operates as a first step in the design process.




Figure III-1: Multidisciplinary Analysis

E. Flight Performance AnalysisAs explained in the previous section the knowledge base and all associated data is shared with each group in the

design process of the Mars aerial platform. Each team utilizes this information to facilitate their separate processes. However, simply supplying the information and leaving each group to their own devices would not result in successful design. Rather, each group must create their own individual disciplinary analysis in order to facilitate efficient and timely processing of data for the completion of the project on its necessarily rigorous schedule.

Included below is the individual disciplinary analysis for the performance group. The same information flow process is applied in a general sense to the MDA as well as the IDA for each group. In the performance IDA, input data, which includes calculated values from several other groups, is used to begin the analysis process. In the second level of cells, most of the major values required to calculate performance data are referenced. Following that is the consideration for major variables, such as air density and velocity, and also mission requirements. The knowledge base is applied to begin the quantification process, as sources to back up analysis methods are a necessity. The major




variables are then applied to their areas of concern and deliverable data is then generated. That data is checked for accuracy against accepted values, and if requirements are met, the data is delivered to the synthesis group.

When data fails to fall into accepted parameters all processes are reviewed. From verifying data, to testing programs for flaws, the information must be checked for accuracy and validity. If no apparent mistakes are located, then the processes and assumptions behind them must be examined. Once these steps are completed and data from other groups has been reviewed, another iteration is performed with the new information. This process repeats until reaching one of two conclusions, either the aircraft converges to the solution space established by the synthesis team, or the design is found to be inherently flawed.

For this project, the IDA takes on a slightly different form. While still resembling any IDA for performance, the responsibilities have been broken into two area, engineering decision and performance analysis. The engineering decisions are based on the knowledge base, which includes industry data, while the needs and inputs of other groups also affect decision making. The decisions themselves become major factors into performance analysis, as battery selection will affect endurance, and landing gear selection will affect takeoff and landing performance.

Figure IV-2: Flight Performance Analysis Methodology




F. Mission ProfileA mission profile is a generalized case in which the aircraft will likely be used. Mission profiles will vary from

aircraft to aircraft depending on their design and specific purposes. A standard passenger aircraft will have a simple mission profile, including a takeoff, cruise and landing. However, the altitude and distance traveled and time will vary greatly between aircraft and their roles. Other aircraft will utilize more complex mission profiles, such as the Global Hawk, a surveillance drone used by the US Air Force. The Global Hawk mission profile would include a loitering portion, a segment in which the Global Hawk circles a position of interest conducting surveillance before returning to land at a designated base. In the case of the Solar Impulse, a solar aircraft with indefinite endurance, two separate cruise altitudes are utilized, a high altitude daytime cruise to collect more solar energy, and a low altitude nighttime cruise to improve efficiency.

Figure III-4: Global Hawk Mission Profile15

Establishing a mission profile for the aircraft requires a review of any standard scenario for which the design may be used. Currently there are two basic designs being considered, rocket propulsion and electrical propulsion. An alternative electrical propulsion system with a rocket assisted takeoff has also been proposed. The general mission profile is consistent across all designs; a standard takeoff, cruise and landing profile. However, since any mission will require a return to a fixed base of operations, FBO, the mission profile must include a second identical segment.

Figure III-5: Proposed Mission Profile




V. Flight Performance MethodologyFlight performance is a quantitative discipline in aircraft design. It may have stringent requirements placed on it

by customer needs or mission requirements, and it is a composite of other disciplines that gives a representative picture of the aircrafts capabilities. This section discusses those parameters and how they are calculated.

A. Primary Flight Performance ParametersThe primary measures of flight performance are as follows:

Range and endurance: A measure of the maximum distance and time an aircraft can travel respectively, often measured idealistically using the Breguet range and endurance equations.

Cruise Altitude and speed: The standard altitude and speed an aircraft will fly at with maximize efficiency.

Takeoff and landing field length: A measure of the necessary distance for an aircraft to land and takeoff.

Rate of Climb: A measure of the speed at which the aircraft gains altitude. Can be derived from excess power.

Stall speed: Aircraft speed at which flow separates from the upper surface of the wing causing a loss in lift.

Glide Ratio: A unitless measure of the unpowered descent of an aircraft, represented by forward velocity divided by downward velocity. This is also directly proportional to CL/CD.

Maneuverability: An analysis of the aircraft’s ability to pitch, roll and yaw. Typically quantified in to radians per second and radius of the maneuver.

B. Determination of Flight Performance ParametersThe two primary variables necessary for flight performance analysis are lift and drag. These are found simply by

considering atmospheric conditions and flight speed. Inputs for Cd0 and CL must be considered for the aircraft as well. The equations are as follows3:

D= (0.5 ) ρV 2 Sref [CD0+K CL

2 ] (1)

L=(0.5) ρV 2 Sref C L (2)

Since the project has not settled on a propulsion system, the range must be calculated for both electrical and rocket propulsion systems. The range and endurance for a rocket propulsion system can be approximated by the Breguet equations. The Breguet range equation and endurance equations are referenced from Fundamentals of Aircraft and Airship Design3:

R= LD

VC

ln [ W i

W f ] (3)

E= LD

1C

ln [ W i

W f ] (4)

In an electrical system, the time a charge lasts is directly proportional to the mass of the battery. At minimum, the mass must be sufficient to last until the solar panels have resumed charging the batteries. According to Design of Solar Powered Airplanes for Continuous Flight the minimum mass of the battery can be found using the following method4:

mbat=T night

ηdischarge kbatPelec (5)




The term ηdischarge represents the efficiency of the battery during discharge and kbat the gravimetric energy density.Assuming we know the power requirements of the aircraft, and the performance characteristics of the battery

itself we are now free to determine an appropriate battery weight based on endurance requirements, or solve for the endurance of the aircraft based on a known battery mass. We also have the ability to vary power usage in order to reduce necessary battery weight or increase endurance. In order to maximize endurance for a given battery mass and type we must minimize the power requirements. Solving purely for time yields the following equation:

T=mbat ηdischarge kbat

Pelec(6)

In order to produce a range value, the endurance equation can simply be multiplied by velocity. However, this only yields a range for maximum endurance flight. The optimal value for range will be found by maximizing V/P. The equation is as follows:

R= VPelec

mbat ηdischarge kbat (7)

This has been reorganized to isolate the V/P term for reference. Cruise speed and altitude are found as part of the process for determining maximum range. When V/P is found at a maximum the associated altitude and velocity are the cruise range and speed for the aircraft. Again, this will be discussed in more detail in the engineering methods section.

For cruise-climb, a slightly different method must be employed; the maximization of the ratio of C L2/D gives the

optimal rate. Stall speed is the minimum speed at which an aircraft can continue to generate lift. Below this speed separation

occurs on the airfoil and the aircraft will begin to stall. Stall speed is an important reference for other values and is calculated as5:

V STALL=√ 2 WρS CLmax

(8)

Rate of climb is another important reference for the mission profile. The Mars Daedalus needs to operate at an acceptable altitude to fly above most obstacles it will encounter. ROC is utilized to determine both service ceiling and absolute ceiling. Service ceiling is considered the altitude at which ROC is 100 ft/min (Assuming a small AOA (less than 15 degrees) ROC can be found as5:

ROC=V (T−DW ) (9)

The power required at a given velocity can be determined from the following equation5:

Preq=12

ρ CD S V 2 (10)

Takeoff ground roll will be calculated using the average acceleration method. The thrust produced at V=0 and at takeoff velocity will be determined. In addition, the rolling drag for the different terrain conditions will also be considered. Taking these values will allow for a start and end acceleration to be compared and determine whether a numerical approach or a simple estimation for average acceleration will be sufficient. For all cases, the calculation will also assume no wind, therefore ground velocity equals airspeed velocity. Takeoff into the wind would be




utilized to reduce the necessary ground speed, so the most important case to consider would be with no wind. The acceleration for takeoff is found using this equation11:

SG=V 2

2a(11)

Alternately, the takeoff distance can be found by integrating the acceleration at different points during lift off should the acceleration be found too far from linear to make a good estimate. The values will be found at different points and distances will be calculated from one point to the next.

For the aircraft landing roll, we must consider the different design possibilities. Both rocket and propeller propulsion can utilize reverse thrust, so the landing equation includes reverse thrust. If the design of the aircraft does not include a reverse thrust system, the thrust term can be set to 0. The landing roll can be calculated using the following equation5:

SGLanding=W

gρS (CD−μr CL)ln ¿¿ (12)

Glide ratio is a unitless ratio that represents the rate at which an aircraft descends. Glide ratio can be found simply by the ratio L/D. This condition must be considered to approximate the best approach for landing, both reducing the forward and downward velocities to increase safety. A powered approach with a high AOA may be preferable to a low powered approach for this configuration. In order to maximize the glide ratio the following condition must be met5:

CD0=K CL2 (13)

Maneuverability is a measure of an aircraft’s ability to change direction. It describes turn ratio, roll capability and other aspects of the aircraft’s characteristics outside of straight flight. For an aircraft whose structure is light and likely to fail under intense maneuvers, only turn rate and radius will be reviewed.

Turn radius is the measure of the radius of the curvature an aircraft takes during a smooth, unchanging turn. Turn rate is a measure of the speed at which an aircraft moves around a turning maneuver. Turn radius and turn rate can be approximated using the following equations, respectively5:

RT=V 2

g √n2−1(14)

The final turn radius and rate of turn values will be based on the maximum load structure that can be applied to the wings, this value will be calculated by the structures team. A necessary safety factor will be added to ensure the safety of crew and vehicle during emergency maneuvers.-

C. Mars Solar Power ModelA solar system is necessary to either power or charge batteries for an electrical propulsion system on the aircraft.

Determining the solar irradiance and power supplied from solar cells is required to determine the feasibility of this option. A review of previously applied methods (the reverse engineering of the solar impulse) and research on solar irradiance on Mars was conducted to produce viable methodology.

Through this research it was determined that the methodology used for reverse engineering the solar impulse was insufficient to describe solar irradiance on Mars. This is due to several factors.

First, Mars’s orbital eccentricity is much greater than Earth’s (.0935 vs .0167). This results in significant differences between solar irradiance at perihelion versus aphelion, while on Earth solar irradiance can be assumed constant. Second is atmospheric make-up.




On Mars, the effects of air molecules on light diffusion can be ignored for any light waves of greater than 200 nm10. This is due to the atmosphere being made of over 95% carbon dioxide and extremely thin relative to Earth. On Earth diffusion from air molecules and water vapor is significant.

Dust is the primary driver of diffusion on Mars, on Earth, dust storms are uncommon and overall diffusion due to cloud cover and other effects can be approximated with simple models. While on Mars, massive dust storms can cover the entire planet and can complicate solar irradiance models.

The maximum amount of sunlight available at the top of the Mars atmosphere can be given by10:

I=cos ( z ) I0( rr )

2

(15)

Where I is the solar irradiance, z is the zenith angle, I0 is the average irradiance (590 W/m2), r is the average Sun-Mars distance (1.52AU) and r is the current Sun-Mars distance. The cosine of the solar zenith angle is10:

cos ( z )=sinθsinδ+cosθcosδcos(h) (16)

The hour angle, h, is measured as the time from local noon and calculated as10:

h=2πt /P (17)

Where P is the length of a Martian day (88775 sec). To determine sinδ the following is applied10:

sinδ=sinεsin Ls (18)

Epsilon is the axial tilt of Mars (25.2 degrees) and LS is the aereocentric angle, an angular measure of the position of the Sun during a Mars orbit. LS of 0 corresponds to the vernal equinox of the northern hemisphere.

The Sun-Mars distance, r, can be found from the following equation10:

( rr )=1+ecos(Ls−LS

P)1−e2 (19)

LSP is the aereocentric angle at perihelion (250°).

Following this methodology produces the maximum Solar Irradiance at the top of the Mars atmosphere. The value, h, hour angle, can be set to zero to find the daily maximum.

The next step in this process is finding the solar irradiance at ground level. Determining the value directly involves complex calculation and integration for each case. Fortunately a simplification is available for this method. Tables producing a coefficient based on assumed conditions can be applied quickly to produce a solar irradiance value. These conditions are the albedo and the optical depth.

Albedo is the solar irradiance attributed to reflected light from a surface. Optical depth is a measure of the extinction coefficient, or the amount of light diffusion due to passing through a medium. In this case the primary medium is dust in the Martian atmosphere. According to Haberle’s paper, an albedo value of 0.25 is typical on Mars, and the optical depth likely ranges from 0.1 in clear conditions to 6.0 in severe dust storms 10. In Haberle’s paper two tables for irradiance coefficients are made available. The first for an albedo of 0.1, the second for 0.4. These tables allow the user to obtain a coefficient to use based on solar zenith angle and optical depth. A third table was generated using linear interpolation for 0.25, an approximation recommended for albedo values between 0.1 and 0.410. These coefficient values, CI, can be applied to find the solar irradiance at ground level:

I G=IC I

1−A(20)




Where A is the value for albedo, CI is taken from the tables A-1, A-2 and A-3. Using an hour angle, h=0, provides the daily maximum for the given conditions.

From this data we can integrate between times of the day to determine power supplied from the solar cells.

∑∫i

f

(IG η¿¿SC sinh−Pelecreq)dt=charge ¿ (21)

Where ηsc is the efficiency of the solar cells. For in flight usage the power term is subtracted due to the demands of the propulsion system and other flight systems. This term can be assumed to be 0 for ground charging.

There is a significant possibility that the Mars Daedalus will be limited in its usage on Mars. This will be the result of reduced solar irradiance due to a significant orbital eccentricity. At its maximum 717 W/m2 reach the surface of Mars, at its minimum it is only 493 W/m2 10. There is a need to understand not only that it is possible to use the aircraft, but to also define a time frame for its usage. It would be a significant waste of resources to launch a mission that would be unable to fulfill its purpose. In order to answer this question the aereocentric angle must be referenced to calendar dates.

The aereocentric angle references the vernal equinox as 0°. The position of perihelion is at the aereocentric angle of 250°. If we assumed the Mars orbit was circular, we could simply create a ratio of days to degrees (687/360) and have a solution. However, Mars eccentricity is significant enough to take into account. In order to reference these positions, the aereocentric angle is referenced directly to an angle from Perihelion. All aereocentric angles under 250 degrees have 110 degrees added, all angle of 250 degrees or more have 250 degrees subtracted from them.

Once this is done we have the orbital term true anomaly, θT. The true anomaly must then be referenced to the mean anomaly, M, a geometric representation of the position of body which corresponds to its position in an elliptical orbit. The mean anomaly allows for time calculations to be performed between any two points on an elliptical orbit. In this particular case, the positions are referenced to calendar days.

M=2 tan−1(√ 1−e1+e

tanθT

2 )− e√1−e2 sinθ1+ecosθ

(22)

The time from perihelion to the mean anomaly is found as:

M=2πT

t (23)

Solving for t and putting T in terms of days produces the Mean anomaly position in days from perihelion. This allows us to take any angle Ls and determine a date on the Martian calendar.

To test the solar irradiance calculator inputs with known results must be entered. As the only two known values are the maximum irradiance of 717 W/m2 and the minimum of 493 W/m2, these values will be checked for. Logically these would fall at perihelion, Ls=250° and aphelion, Ls=70°. The results are as follows:

Ls=250° and latitude 23.6° S: 717.8 W/m2

Ls=70° and latitude 23.6° N: 493.5 W/m2

This values for solar irradiance at perihelion and aphelion validate the use of the irradiance calculator for this project.

VI. Design DecisionsAs was discussed previously three design decisions were made the responsibility of the performance group.

These include the selection of an appropriate landing gear configuration, batteries for the aircraft should an electrical propulsion system be selected, and solar cells to either power or charge the electrical system. This section of the report deals with those design decisions.

A. Landing Gear Selection




To facilitate landing gear selection, a review of standard landing gear systems was performed. This process included reviewing several sources for landing gear types, configurations and design considerations. The following are the primary landing gear arrangements and types for small aircraft:

Tail wheel landing gear: two wheels positioned in front of the center of gravity and a third positioned on the tail, typically requires higher takeoff speeds, and provides more stability on the ground than other types

Tricycle landing gear: two wheels positioned shortly aft of the center of gravity and one in front, lower takeoff speeds and less stability on the ground

Retractable: landing gear that stow inside the fuselage during flight, reduces parasite drag but weighs more and requires space inside the fuselage

Shock absorbing systems are necessary for an aircraft landing in rough terrain. The shock of landing is reduced with a proper shock absorber, and the load in the airframe and especially in the wings is alleviated. The following are a list of shock absorbing systems used for landing gear for small aircraft:

Rigid: the impact of landing is reduced solely through the use of flexible tires and most of the load is transferred to the airframe

Leaf-spring: the use of spring steel absorbs some of the impact by flexing and transfers the shock to the airframe at a reduced rate

Shock struts: pneumatic or hydraulic struts which absorb the impact and dampen it by pressing fluid through an orifice

Bungee Cord system: rigid structure is attached to airframe by a hinge, the landing gear movement is resisted by a bungee cord or other spring system

Research on the Mars terrain has revealed several distinct repeating features. One of which is dune fields, hundreds of which exceed one square kilometer and can be in the hundreds if not thousands of square kilometers 20. Additionally, areas of bedrock exist which a landing could occur on, however, estimates on smooth bedrock size and regularity have not been located. Difficult if not impossible for this design to operate in, is rocky fields which can extend several kilometers in diameter. As shown in figure II-2, these rocks due extensive damage to wheels.

The sandy terrain on Mars still presents issues, even large areas which are more suitable for aircraft operation are covered in small dunes measuring several centimeters high. To allow for the best possible result the most rugged standard landing gear will be utilized, a bungee cord or spring system will be used to greater flexibility in the landing gear without presenting the concern of failure which would accompany a hydraulic are pneumatic strut. The landing gear will not retract, but tires can be covered by a fairing to improve aerodynamic qualities, reducing parasite drag.

After surveying additional information on aircraft, alternate options were found to improve landing performance. The use of skis on the sand dunes would reduce the friction coefficient from 0.4 to 0.2. Some

types of skis have holes which allow the tires to continue their contact, but prevent the tire from sinking into the sand, others allow the tire to sit on top of the ski. Either system can be designed with a quick release catches that allow for easy removal should there be an intent to land on rock bed. Additionally, the use of skis is not a new practice. Skis have been installed on aircraft for landing gear for decades, from the P-51 to the C-130 and even the Chinook helicopter, this method is reliable and often used for landing on challenging surfaces.

While sand presents power issues a model for the performance of tires on a rock bed produce a rolling resistance coefficient of approximately 0.05. This performance far exceeds that of landing on sand, however the lack of a smooth surface free from potential damage from rocks is difficult to guarantee. An assumption that small stones should not pose a major issue for this aircraft will is made. An attempt to model a basic spring mass damper system was determined insufficient to model the system, as the contribution of tire deformation is significant.

Originally a tail aft landing gear was selected for better ground stability. However, to meet takeoff needs, a change to tricycle gear was made. This change will allow for rotation during the takeoff run and reduce the


Figure V1-1: Ski Landing Gear22



necessary takeoff velocity. Due to the short length of the aircraft, about 6 meters, the inclusion of tricycle landing gear should not compromise the design.

Figure VI-2: Geometric Estimation of Landing Gear Positions

B. Battery SelectionIf an electrical propulsion system is selected, a battery is necessary. There are two primary considerations for

battery selection due to the nature of this mission, the operating temperature and specific gravity. In addition, the battery must be rechargeable, sometimes called a secondary battery.

The weather on Mars is very cold compared to Earth. The diurnal temperature range recorded by NASA at the Viking 1 landing site was found to be 184 to 242 K6. Curiosity sensors found temperature ranges from -103 to 37 F (198 to 276 K)7. Any battery utilized for operating on Mars would have to be capable of operating at a very low temperature with little insulation.

To accommodate the need for a low speed, aircraft weight needs to be minimized. To achieve sufficient power to run the aircraft and minimize weight, a high specific gravity battery is preferred. The total energy needs to be sufficient to power the aircraft, while also supplying sufficient peak power. Both of these conditions are being considered as part of the design process.

Included below is data from research on batteries. Research of industry information was also conducted to supplement data.




Table VI-1: Battery Data8

Table VI-2: Oxis Energy Ultra-Light Lithium Sulfur Pouch9

The table above is industry data for a battery to start manufacture in 2019. Current batteries achieve similar performance with a specific energy of 325 Wh/kg. Oxis data gives operational temperatures as low as 243 K9. The most suitable option for batteries is the Oxis Energy Ultra-Light Lithium Sulfur Pouch, this will be used in the future for all analysis involving batteries and will be used as an energy source for an electrical propulsions system.

C. Solar Cell SelectionThe selection of a solar cell will be dependent on one of three scenarios. The first of which is as a power supply

to run the aircraft’s electrical propulsion. The second is a charging system for batteries to accompany the aircraft. The third is a stationary charging system.

A system that accompanies the aircraft must be lightweight and efficient. It must also supply enough power to supplement the aircraft, preferably provide all the aircrafts power requirements. This system could be left active to charge batteries after landing. Drawbacks include additional weight and thus more power required for flight.

A system solely for charging the aircraft batteries must do so quickly. The charge must occur in less time than astronaut life support systems would last. Complications with this system include accounting for minimum charging time of batteries to avoid overheating, thus requiring minimum time on site for astronauts, and risk of failure on site stranding astronauts. This system becomes less feasible due to these additional risks.

A stationary charging system could be kept at the landing site with astronauts, this system would be repairable and weight considerations would be negligible compared to an onboard system and could serve to charge and power additional equipment. Conversely, this system would limit range to a single battery discharge.

After a review of industry data on ultra-lightweight cells, the final selection of the SolAero IMM4J CIC was made. The specifications of concern are as follows11:




49 mg/cm2 (.49 kg/m2) Minimum 33.0% efficiency

The selection of this cell was based on a review of over a dozen similar purpose ultra-light cells. The SolAero product outperformed them all in both weight and efficiency, making them the best fit for any scenario.

D. Rocket SelectionAs previously addressed, there are two requirements for rocket selection. The ability to deliver the payload mass,

and the appropriate area in which the payload can be packed. This process will not be completed until nearly all design and engineering decisions have been made, but will serve as a check to ensure aspects of the design are not becoming unfeasible throughout the process.

In order to save weight, the launch of the Daedalus will be addressed as a separate launch, this process is a common proposal for manned space missions, to save weight on an individual launch the payload is divided into multiple parts so that life support systems and food supplies for astronauts can be more robust, and manned trips can be planned for faster travel time, which requires much more fuel. Equipment is launched before the departure of astronauts on a slower journey to allow for fuel savings. Launching an entire manned mission and associated equipment becomes impractical, if not impossible, as the weight of the rocket and fuel is exponentially related to the weight of the payload.

Determining the requirements of a rocket to launch a payload is a complicated process, and not a primary focus of the project. A rough estimate is the only consideration being made, and the selection of an existing rocket that meets these estimates will be performed. The first step is to collect data on current rockets being used for space launch. Most rockets have been excluded from this list as payloads were grossly insufficient to launch an MTV plus fuel into orbit. The following are American rockets which have not been retired:

Table VI-3: Rocket DatabaseRocket Database First Stage Second StageRockets Sp Impulse (s) Struc Mass (kg) Prop Mass (kg) Sp Impulse (s) Struc Mass (kg) Fuel Mass (kg) Diameter (m)Falcon 9 v1.1 282 23100 395700 340 3900 92670 5.2Atlas V 311.3 21054 284089 450.5 2316 20830 3.05Delta IV Heavy 360 78000 601200 462 3480 27220 5.1Antares 297 18700 242000 304 1010 12887 2.34

All the rockets listed include multiple stages, a rocket can be analyzed stage by stage backward through the launch process. The final stage, which is the MTV and payload will be used to escape orbit and transfer to Mars. The equation for estimating rocket size ignores aerodynamic forces. It calculates based on weights, specific impulse and change in velocity. The equation is as follows12:

M 0=M f∗e∆V

I sp g0 (24)

Where M0 is the starting mass, Mf is the final mass and Isp is the specific impulse of the rocket. The final mass includes the structure mass and payload of the rocket, and the starting mass also includes fuel. This process can be repeated for the stages, where the payload of the preceding stage equals the total mass of the following stage. All of the values are pulled from the database, and delta-V is adjusted in order to minimize the rocket’s weight.

To further accommodate the limitations, the diameter of the rocket is considered and payload fairing sizes are approximated to determine reasonable dimensions for the transport area. This data will be given to the synthesis team in order to determine an appropriate packing method. Should a viable packing method not be possible for a particular rocket, it will be eliminated.

In order to accelerate the selection process, the Centaur, a later stage rocket, will be used to model the MTV for any rockets with insufficient data. After a review of available data the Centaur was found to have the most complete information, a long history of use and has successfully launched large payloads to Mars. The relevant data for the Centaur is included below21:




3.05 m diameter (stored inside fairing) 12.68 m height 2247 kg mass 20830 kg propellant mass maximum 450.5 s specific impulse

In the case of the Delta IV heavy, sufficient information was available on its second stage to replace the Centaur for modeling purposes. This stage is known as the Delta Cryogenic rocket. Listed below is a set of data used for modeling21:

5 meter diameter 3490 kg structure mass 27220 kg fuel mass maximum 462 specific impulse

An alternative to these rockets is available, the Falcon 9 Heavy, which SpaceX lists as capable of delivering a payload of 13,600 kg to Mars. Certainly this would accomplish the mission, but sufficient data on the rocket is not yet public so the analysis cannot be performed.

A survey of fairings has allowed for a more detailed analysis of packing for the aircraft. The Delta IV Heavy utilizes several different fairings. While smaller fairings are inadequate to carry the MTV and payload, a larger fairing manufactured by Boeing can be modified to fit the MTV and give sufficient space. In addition, a company called ATA has patented an asymmetric fairing which has a max width of approximately 15m and length of 29.2 m. This fairing would allow sufficient space to pack the aircraft and MTV without excess manipulation and disassembly of the design for packing. However, the asymmetric fairing is 1600 kg more massive than a standard fairing, this constitutes more mass than the estimated payload.

Figure VI-3: Delta IV Heavy Payload Fairings24




It is also necessary to consider a reentry system for a mission to Mars. While no analysis will be performed, it is necessary to include the mass of a reentry system for launch calculations. A survey of reentry systems was conducted in hopes of finding a light weight system.

For most missions involving putting equipment on the surface of Mars, parachutes have been utilized. Viking, Pathfinder and Curiosity have all utilized parachutes for reentry. Curiosity made use of a system called the “sky crane26,” a rocket powered crane which slows to a hover and lowered Curiosity on to the Martian surface. Other smaller rovers have used an airbag system. A large system of inflated bubbles which provide drag and prevent damage from impacting the planet’s surface. The basic advantages to these systems are as follows:

Both proven systems Parachute-airbag system is

simple and light weight Sky crane system

guarantees soft landing

Adversely these systems have drawbacks for the Mars Daedalus design which make them less likely to succeed. It is light weight but bulky. It is also more delicate and any excessive impact would likely cause damage.

The disadvantages of the systems are given below: Airbag system introduces large forces on structures Sky crane system is very heavy, over 2400 kg26

An alternate reentry system, IRVE (inflatable Re-entry Vehicle), is currently in a testing phase. NASA has developed this system as a method to increase the diameter of the reentry vehicle without adding excess weight. The additional diameter gives a larger area, resulting in additional dynamic breaking and permits the payload to be slowed without complex additional systems. The test plan shows that the impact velocity should be less than 77 km/hour on Earth. With additional area, that velocity can be slowed more. Moreover, the shielding itself will act as an airbag, thus a parachute and airbag system are included in a single assembly and the payload does not reaccelerate. While the system presents similar advantages as the parachute-airbag system, it doesn’t have a history, and may not be reliable.

LDSD (low density supersonic decelerator) is another system under development for use in Mars landing, due to the size, 8m across, and weight of 3120 kg27, it is not practical for this purpose. The LDSD is a heat shield, with airbags and parachute packed behind it, and it utilizes a rocket for control.

For the purposes of weight savings, the IRVE will be selected as the reentry system. At 1400 lbm (635 kg), this system carries significant advantages in reducing the necessary space during transport, and mass to be delivered. While additional testing and development is necessary, it will be applied for this system to model a lightweight reentry system.


Figure VI-4: Standard Fairing vs Asymmetric Fairing25

Figure VI-4: IRVE Reentry System30



E. Mission DatesThere are three requirements to consider for the purpose of narrowing down a list of possible mission dates: A mission does not exceed the delta-V capabilities of the MTV with payload Mission dates precede manned dates by a reasonable time frame Mission dates overlap dates with sufficient solar irradiance if solar power is selected for propulsion

A database of potential Mars mission dates is available through NASA online. Several sets of constraints can be applied. Using this database, two sets of information were established. One of which is one-way trips to Mars, the other is Mars trajectories with returns. Both of these will be used for comparison to determine viable mission dates. A database of these missions is given in the appendix.

A reference between Mars calendar dates and Earth calendar dates is necessary to reference the days on Mars in which there will be sufficient solar power for flight. A table was located at the website for the Planetary Society that did exactly this. An excerpt of that table is included below:

Table VI-4: Martian vs Earth Calendar Reference21

Spring equinox Summer solstice Autumnal equinox Winter solstice (Ls = 0°) (Ls = 90°) (Ls = 180°) (Ls = 270°)

33 Jun 18 2015 Jan 03 2016 Jul 04 2016 Nov 28 2016

34 May 05 2017 Nov 20 2017 May 22 2018 Oct 16 2018

35 Mar 23 2019 Oct 08 2019 Apr 08 2020 Sep 02 2020

36 Feb 07 2021 Aug 25 2021 Feb 24 2022 Jul 21 202237 Dec 26 2022 Jul 12 2023 Jan 12 2024 Jun 07 2024

38 Nov 12 2024 May 29 2025 Nov 29 2025 Apr 25 202639 Sep 30 2026 Apr 16 2027 Oct 17 2027 Mar 12 2028

40 Aug 17 2028 Mar 03 2029 Sep 03 2029 Jan 28 2030

Mars year

VII. ResultsAs a part of this process, results showing that the necessary data has been delivered is mandatory. This section

deals with data and will deliver results that coincide with the most recent iterations. From the original iteration the primary performance focuses were range and takeoff and landing ground roll. Those values are given below:

Takeoff ground roll on sand: 1087.8 m Takeoff ground roll on bedrock: 972 m Landing on sand: 881.6 m Landing on rockbed: 1007.8 m Ideal Range: 124.6 km Mission Profile Range: 115.9 km

These results were accompanied with a set of recommendations to improve performance for future iterations. Through the process several adjustments were made. The following actions were taken based on those recommendations:

Recommended at least 50% increase in wing area Wing area increased from 11.5 to 22.4 m2

Recommended increase in engine thrust for takeoff Engine thrust increased across takeoff range Split wingspan in half for packaging, included Add reverse thrust for landing, not included Include minimum 50 kg battery, maximum 100 kg 70 kg battery included




Include solar cells for additional power 10.98 kg of solar cells included

A. InputsFrom several different groups input data was obtained. This data was received from several different groups and

processed for results. The data received and groups delivering each are as follows:

Synthesis: Wing area: 22.4 m2

Structures: Weight: 1476.2 N

Propulsions:Table VII-1: Propulsion Data Table 128

0m TO 750m 1500m 0m landPreq 12.06227485 29.63979139 29.3954793 12.092265drag N 271.0623562 240.9739137 235.1638344 274.8242drag lbs 60.93723012 54.17308047 52.86692293 61.782927

Table VII-1: Propulsion Data Table 228

1500m Cruise 750m Cruise Takeoff Landingpitch (deg)

rpm0.985 0.985 0.985 0.985

1 1 1 10.95 0.95 0.95 0.95

0.727 0.7324 0.6767 0.67650.92 0.92 0.92 0.92

a1= 1.47 1.46 1.58 1.58200 200 200 200

0 0 0 0a2= 217.39 217.39 217.39 217.39

29395.5 23639.8 12062.30 12092.3043.43 34.71 19.27 19.32

Figure VII-1: Thrust Available by velocity28




Aerodynamics:Table VII-2: Aerodynamics Outputs29

Takeoff Low Alt. Cruise High Alt. Cruise LandingAltitude m 0 750 1500 0

Velocity m/s 42.75 118.00 118.00 42.50Roe 0.02110 0.020350 0.019350 0.02110

Mach 0.1719 0.4751 0.4759 0.1709CL 3.4346 0.4691 0.4929 3.5218CD 0.8374 0.0770 0.0783 0.8748

From the excel sheets for both performance and aerodynamics, edits can be made which allow for values to be found at different velocities. This editing method was used to pull values across different takeoff velocities to estimate takeoff distances for multiple cases.

B. Performance ValuesIt was determined that takeoff ground roll could not be assumed constant. Acceleration moved from 4.3 to 1.3

m/s2. In order to estimate the takeoff ground roll the Heun method was utilized. The Heun method is a step integration where acceleration is assumed near constant for a range, and the midpoint of the range is used as reference. This continues through the full set of velocities until takeoff velocity is reached. From this the distance traveled over that range is calculated from equation (11) and all values are summed for a final value. Time is found by taking the reference velocity and dividing the distance traveled by that value. Total time is found again by summing the values. Takeoff velocity was estimated to be 42.75 m/s by the aerodynamics. The following table gives all the necessary inputs and calculated values for takeoff on both sand and on bedrock, coefficients of rolling friction were assumed 0.2 and 0.05 respectively:

Table VII-3: Sand Takeoff Ground Roll and Timevelocity 5.0 15.0 25.0 35.0 42.5mass 397.9 397.9 397.9 397.9 397.9lift 20.0 182.0 505.0 992.0 1466.0friction co 0.2 0.2 0.2 0.2 0.2gravity 3.7 3.7 3.7 3.7 3.7weight 1476.2 1476.2 1476.2 1476.2 1476.2weight-lift 1456.2 1294.2 971.2 484.2 10.2friction 291.2 258.8 194.2 96.8 2.0drag 4.3 39.0 108.7 213.7 316.1thrust 1994.7 1503.2 1145.3 920.0 850.0net force 1699.2 1205.4 842.4 609.5 531.9accel 4.3 3.0 2.1 1.5 1.3seg dist 11.7 49.5 118.1 228.5 85.1total dist 11.7 61.2 179.3 407.8 492.9seg time 2.3 4.1 7.2 11.7 11.6total time 2.3 6.4 13.6 25.2 36.8

Table VII-4: Bedrock Takeoff Ground Roll and Time




velocity 5.0 15.0 25.0 35.0 42.5mass 397.9 397.9 397.9 397.9 397.9lift 20.0 182.0 505.0 992.0 1466.0friction co 0.1 0.1 0.1 0.1 0.1gravity 3.7 3.7 3.7 3.7 3.7weight 1476.2 1476.2 1476.2 1476.2 1476.2weight-lift 1456.2 1294.2 971.2 484.2 10.2friction 72.8 64.7 48.6 24.2 0.5drag 4.3 39.0 108.7 213.7 316.1thrust 1994.7 1503.2 1145.3 920.0 850.0net force 1917.6 1399.5 988.0 682.1 533.4accel 4.8 3.5 2.5 1.7 1.3seg dist 10.4 42.6 100.7 204.2 84.9total dist 10.4 53.0 153.7 357.9 442.8seg time 2.1 3.5 6.1 10.2 10.4total time 2.1 5.6 11.8 22.0 32.4

For decreasing landing roll, a parachute was added. The estimation of the parachute’s effect was to multiply the drag coefficient by 7.6. The inclusion of reverse thrust was declined in favor of this method. From equation (12):

Landing on sand: 341.8 m Landing on rockbed: 498.1 m

By equation (6) the ideal endurance and range were found to be as follows: 0.69 hrs 293.1 km

These values are limited in that there is no solar power included. Later in this report, a basic mission profile is developed to refine the data.

To determine the time spent in climb and descent phases, it is necessary to calculate rate of climb. While the aircraft is capable of producing much more thrust than is used, a value was selected to represent better efficiency over this time frame. The selection was made numerically and the data was included on the mission profile.

Climbing phase: Power required: 29.4 kw Power available: 200 kw Power setting: 44 kw Rate of climb: 9.72 m/s Time to climb to 3000 ft: 308 sec

Descent phase: Power required: 34.72 kw Power available: 200 kw Power setting: 15 kw Rate of descent: 9.92 m/s Time to descend: 303 sec

C. Solar Power AssessmentAs described in the solar power analysis, a model was built to calculate solar power at different times of the

Martian year, at different latitudes and times of day. As discussed earlier, albedo was assumed to be 0.25. From this model the following data was determined:




Figure VII-2: Solar Irradiance at 0.2 Optical Depth on Mars Equator

Figure VII-3: Solar Irradiance at 0.2 Optical Depth at 30 Degree South Latitude

Figure VII-4: Days per 30 Degrees True Anomaly

For the first iteration the solar panels were assumed to cover the full wing area of 22.4 m2. Based on perihelion at 30 degree south latitude the values for solar power were as follows:

Midday solar irradiance: 701.9 W/m2

Solar power max: 2617.5 W




Currently, maneuverability data cannot be established, as no maximum force load estimates have been received from structures. As this is not crucial to mission success, it may be omitted.

D. Mission ProfileTo approximate a mission profile for the Mars Daedalus the power consumption and time of each stage had to be

approximated. Takeoff and landing values, as well as climb and descent data, has been addressed above. Power requirements were provided from the propulsions group. Only two values can be adjusted during this step to achieve desired results, takeoff time and time at cruise. As inputs for a first step, a battery mass of 100 kg was included. Recommendations to improve performance must be suggested for future iterations. The following is data from this process:

Battery mass: 70 kg Specific density: 450 Wh/kg Total energy: 31.5 kJ

In order to simplify the model only a half of the two flight mission plan was calculated. To prevent stranding the astronaut, the duration was limited to half the energy available from the batteries. The table is as follows:

Table VII-5: Half-Mission Profile Dataphase time (hours) time (s) start theta (rad) final theta solar energy (Wh) Ereq (Wh) Net Energy Sum (Wh)

takeoff 0.0102 37 1.57 1.57 43.88 197.23 -153.36 -153.36climb 0.0857 308 1.57 1.59 367.19 4414.41 -4047.22 -4200.58cruise 0.2628 946 1.59 1.66 1124.40 11412.44 -10288.04 -14488.62descent 0.0840 303 1.66 1.68 358.35 1474.70 -1116.35 -15604.97landing 0.0092 33 1.68 1.69 39.04 177.10 -138.06 -15743.02

The sum energy, measured in Wh, was capped at 15,750 for the first half. While some charging will occur while the pilot is on the ground at a rate of 2486.6 W, the short duration that the astronaut will be detached is not sufficient to make major considerations. Currently solar power supplies approximately 9% of energy needed for this flight profile. Should solar power be excluded for powering the aircraft, the time at cruise would be reduced by approximately 160 seconds.

The total endurance, a more applicable methodology, can be found by adding the time for each of the flight phases. This gives us 1397 seconds, 23.3 minutes. This value leaves approximately less than 0.5% of power left over for each half-mission stage.

Range can be determined by applying the time spent in each flight condition and multiplying by its velocity. From this we obtain the following data:

Table VII-6: Mission Profile Based Rangephase Time (sec) Velocity (m/s) dist (km)climb 308 118 36.4cruise 946 118 111.6

descent 303 118 35.7183.7

It is important to note, this calculation is for the half-mission profile. This allows the astronaut to fly to a location, perform a task and return to the home base. To better visualize this value, a radius of this range was placed on a map centered on the UTA campus. The associated image reveals that the distance traveled by the aircraft would allow it to travel to Temple, and Tyler, Texas, as well as Ardmore, Oklahoma.




Figure VII-5: Range Illustrated from UTA Campus31

E. Rocket AnalysisUtilizing the analysis given in equation (24), it was found that the current best performing rocket, which also

meets the packing requirements is the Delta IV Heavy. The data found for this rocket is given below:

Table VII-7: Delta IV Maximum PerformanceStage 2 Stage 1

Delta V 8.14 5.69 km/sSpecific Impulse 462 311.3 secgravity 0.00981 0.00981 km/s^2structure mass 3480 78000 kgpayload mass 1900 32418.13 kgfinal mass 5380 110418.1 kgtotal mass 32418.13 711584.6 kgfuel mass 27038.13 601166.5 kg

An estimated value of 1900 kg for the payload mass was found by summing the different necessary components for the flight. These included the following:

Aircraft mass without payload: 217 kg Inflatable reentry system: 650 kg Supersonic reentry parachute: 300 kg Aircraft Enclosure 83 kg

The analysis concludes that the Delta IV Heavy can serve to transport the payload into LEO at 200 km. This is estimated to be a delta-V of 7.78 km/s. The first stage accomplishes 5.69 km/s of the impulse, after which the second stage detaches and reaches orbit. This leaves 6.05 km/s of delta-V available for the transfer to Mars.

From this point an analysis of missions with returns can be performed selecting missions with stays primarily during the summer for the Mars southern hemisphere. This will eliminate as many possibilities and prevent repeating the analysis a second time:

Table VII-8: Mars Southern Summer Missions with Returns




The second step of this analysis is to select the prospective launch dates which fall within the delta-V requirements. Simply ordering the database by the delta-V and cutting missions above the mark accomplishes this. These missions can then be ordered against the return missions to determine which can deliver the Mars Daedalus within a window of approximately 12 months of the manned mission’s arrival. The qualifying missions are as follows:

Table VII-9: Delta-V Qualifying Mars Missions

Earth_Departure Dest_ArrivalDuration

(days)C3 (km2/s2)

Injection DV(km/s)

Post-injection DV (km/s)

Total DV (km/s)

Route

Nov-24-2026 Jul-22-2027 240 14.3 3.86 1.06 4.92 EMDec-10-2026 Jul-06-2027 208 26.1 4.36 1.45 5.81 EM

Future iterations may include more, or exclude some of these missions. It is important to note that this is a first order analysis and needs much more research before any real conclusions can be made.

A request for data was made by cost analysis. Based on mission profiles, what would be the most payload mass that could be carried to orbit and delivered to Mars on a selected launch date? By selecting from the previous data, the mission with the lowest delta-V was selected to maximize payload mass. From this the following data was calculated:

Table VII-10: Maximum Payload EstimateStage 2 Stage 1

Delta V 7.02 5.68 km/sSpecific Impulse 462 311.3 secgravity 0.00981 0.00981 km/s^2structure mass 3480 78000 kgpayload mass 3500 32850.23 kgfinal mass 6980 110850.2 kgtotal mass 32850.23 712033.9 kgfuel mass 25870.23 601183.6 kg

This leaves 1600 kg of potential payload for additional equipment. This can include rovers, sensors, food, or many other required supplies for a mission to Mars.

VIII. ABET RequirementsThis section is devoted to addressing the ABET requirements for this project and class. Each requirement will be

addressed individually in the following sections of this report. The purpose of this is to show that learning objectives have been met and that the student has progressed in their education and is ready for graduation and joining the job market as a successful engineer.

A. Outcome C: Design System, Component or Process to Meet Needs




From the ABET requirements submitted to the team:

Plan to accomplish:1) Specify technical and managerial requirements for assigned team design project.2) Request comprehensive literature search using professional-quality resources.3) Task the teams & individual students to produce a team/individual semester task & time plan aimed at producing a timely and performing system or component.4) Specify mile-stone deliverables and their expected quality.5) Question students during weekly contact team meetings about their design approach meeting the project requirements.

Plan to demonstrate:1) Require and grade be-weekly individual design reports throughout the semester. See samples of bi-weekly design reports in course exhibit.2) Require and grade by every student a report chapter explicitly outlining the product development strategy and overall development plan.3) Require and grade a detailed report discussing the design of a system, component or process to meet overall mission objectives.

Key Assignment for Outcome C: The mid-term individual report and the final semester report, both documents demonstrate the student’s ability to design a system, a component and or process to mission within realistic constraints such as economic, environmental, social, political, ethical, health and safety, manufacturability and sustainability. The very nature of this class is to synthesize individual components into the systems context. The students must use multi-disciplinary and disciplinary methodologies, analysis, geometry visualization, and possibly testing in the design of their components and systems. See samples of individual (a) mid-term reports and (b) final semester reports in course exhibit.

Each individual discipline addresses this requirement through their respective IDA and engineering methodology. The IDA serves as a tool to understand the multi-disciplinary aspects of the analysis of separate disciplines. Essentially, the design of an aircraft can be broken into discipline, but must be approached holistically with each discipline serving not only its purpose, but the purpose of other disciplines. The methodology development of each discipline is a product of the disciplinary analysis and helps to fulfill the requirements as well.

B. Outcome D: An Ability to Function on Multidisciplinary TeamsThe ABET requirements are as follows:

Plan to accomplish:1) Divide class into two competing design teams, each consisting of disciplinary sub-team structure.2) Assist teams in choosing a corporate identity and task to formulate a product development strategy and business case.3) Introduce the chief engineers to define a project multi-disciplinary methodology (MDA) and responsibilities, and all disciplinary engineers to formulate corresponding individual disciplinary methodologies (IDA) & responsibilities.4) Task project teams to produce (a) team/individual weekly update presentation(s), (b) bi-weekly individual project report.

Plan to demonstrate:1) Collect from each team the written organization plan, assigned duties and time schedule for completing tasks. See samples of bi-weekly design reports in course exhibit.2) Require each team member to evaluate other members of the team as to his or her contributions to the design project. Use these evaluations in determining final grades.3) Observe and grade each team member’s contributions during weekly meetings and bi-weekly progress reports. See samples of bi-weekly design reports in course exhibit.




Key Assignment for Outcome D: It is the very nature of this course to require students to work in teams on a project. The student team must respond to a request for proposal (RFP) with their ‘corporation’ which they have formed to work on this semester project. A chief engineer is selected who organizes managerial and engineering activities of the disciplinary groups. For each design discipline (e.g., aerodynamics, structures, cost), disciplinary sub-teams and team leaders are formed, dividing the disciplinary workload while depending upon each other to accomplish the group goals. The groups must work through problems of under-achievers, design trade-offs, and work towards a published schedule & budget. The teams always use a multi-disciplinary approach to the project and utilize theory and experience gained from a variety of courses. All oral and written reports demonstrate the inner workings of the team. These reports are required to pass the course. See samples of individual (a) mid-term reports and (b) final semester reports in course exhibit.

Throughout the process, the goals of outcome D were met. Additional work beyond the standard disciplinary analysis was required, which was distributed among the teams. Much of this work was redistributed to other teams which were in better position to handle the needs due to more successful time management or better knowledge of a particular task. As the performance lead certain tasks were expected, but a great deal of additional work was undertaken. General landing gear selection, modeling of solar irradiance on Mars, rocket sizing and rocket payload estimation as well as selection of potential mission launch dates were all tasks generally outside of performance which were accomplished by this group. It is easily observable that the performance group worked hard to meet the dynamic needs of the teams design project.

C. Outcome E: Understand Professional and Ethical ResponsibilityFrom the ABET requirements:

Plan to accomplish:

1) Classroom discussions about professional and ethical responsibility of the aerospace vehicle designer and technology forecaster.2) Dedicate one lecture to flight vehicle safety, certification, and incident & accident investigation.3) Reading assignment of selected chapter in Aircraft Safety – Accident Investigations, Analyses & Applications by S.S. Krause, McGraw-Hill, 20034) Organize a speaker from the FAA to introduce the subject flight safety and certification.

Plan to demonstrate:1) Require an individual report chapter addressing design for safety & reliability. The overall design methodology needs to contain a concrete approach to design for safety & certification next to design for mission objectives. This bi-weekly report assignment counts as a major part of the course grade. See samples of bi-weekly project reports in course exhibit.2) The students are required to research and document a project-relevant flight vehicle accident design case study. From the case study, it is required to take certain lessons learned into account with the safety methodology devised. Particular emphasis is directed to the discussion of engineering responsibility and the often fatal consequences in wrong-doing. See samples of bi-weekly design reports in course exhibit.3) Document in the individual report how Outcome F does relate to the individual student project’s responsibility.

Key Assignment for Outcome F: This particular bi-weekly report assignment explicitly requires describing the systematic approach implemented towards professional & ethical responsibility overall resulting in enhanced flight safety.

As part of the analysis process it is the ethical responsibility to not only perform all tasks diligently and honestly, but to also submit all assumptions made in the analysis process. This is a task taken seriously, that any assumptions made regarding data were addressed in the results section, and all input data was drawn from respective groups and cited accordingly. This process allows assumptions to be checked and results to be more reliable, thus improving the safety of the aircraft.




D. Outcome F: An Ability to Communicate EffectivelyAs per ABET requirements:

Plan to accomplish:1) Specify requirements for written reports and oral MS PPT presentations: (a) organization & presentation, (b) content & originality, (c) practical application and feasibility, (d) addressing the average audience consisting of the decision-maker, specialist and layman. Outline use of text, tables, figures, and graphs.2) Individual students receive written bi-weekly report feedback.3) Three student-faculty contact opportunities per week are utilized to develop the skills of efficient oral communication and giving MS PPT update presentations.

Plan to demonstrate:1) Grade bi-weekly individual written reports. See samples of bi-weekly project reports in course exhibit.2) Grade mid-term and final reports. See samples of reports in course exhibit.3) Grade mid-term and final team & individual presentation performance: (a) presentation material, and (b) oral presentation skills. See samples of recorded mid-term and final presentations in course exhibit.

Key Assignment for Outcome G: The mid-term report & presentation and the final report & presentation serve to document the student’s ability to communicate effectively.

The establishment of effective means of communication is essential for any team to complete a project. In order to enable effective communications a method for direct communication and another for sharing of data and research was established. The application Groupme was utilized to keep the team in quick contact with each other. For data sharing the online cloud service Google Drive was used to allow for sharing of data. These programs accelerated the process and gave an avenue for effective communication.

E. Outcome H: Understand Impact of Eng. Solutions in Global and Societal ContextThe ABET requirements for this outcome are as follows:

Plan to accomplish:1) Dedicate one report chapter to the history of aerospace. Emphasize on the catalyst effect engineering solutions have on human development in the past, presence and future.2) Select a pertinent case study discussing the implications of emerging aerospace technologies on the (a) environment, (b) global and domestic economy, and (c) global and domestic societies and politics.

Plan to demonstrate:1) Grade bi-weekly individual written reports. See samples of bi-weekly project reports in course exhibit.2) Grade mid-term and final reports. See samples of reports in course exhibit.3) Grade mid-term and final team & individual presentation performance: (a) presentation material, and (b) oral presentation skills. See samples of recorded mid-term and final presentations in course exhibit.

Key Assignment for Outcome H: The report chapters and presentations are discussing the impact of engineering solutions in the global and societal context. The reports/presentations are graded on the organization and content of the cause-effect relations engineering solutions will have on the emerging aerospace technology in the global and societal context. See samples of individual written reports covering Outcome H.

This requirement was fulfilled by the history of Mars exploration. A brief history discussing the successes and failures of missions to Mars. The purpose of this is to gain from the knowledge of those who came before. A quote which best captures what is being accomplished by this is one commonly credited to Isaac Newton, “If I see further it is because I stand on the shoulders of giants.”

F. Outcome I: Recognize the Need and Ability to Engage in Lifelong Learning




The ABET requirements are as follows:

Plan to accomplish:1) Require an extensive literature search to be performed during the first weeks of class. The search is documented in the bi-weekly project reports covering project relevant aspect from the (a) past, the (b) presence, and the (c) projected future.2) Require the literature search chapter to discuss the significance of project-relevant past, present and future design knowledge, and in particular how the design knowledge has been evolving (increasing or decreasing) over time.3) Require each individual student to build a disciplinary data-base (DB, like overview tables & figures) and knowledge-base (KB), like lessons learned, design guideline, trend lines, etc.) overall aimed at retaining, organizing, and making available relevant design data, information and knowledge.4) Require each student to utilize and enrich the DB & KB with new information generated throughout the project. Emphasize the need to engage in lifelong learning in order to efficiently shape future engineering products.

Plan to demonstrate:1) Require and grade the particular bi-weekly report as a major part of the course grade aimed at delivering the primary DB & KB for the project. See samples of bi-weekly progress reports in course exhibit.2) Require an individual student progress presentation as a major part of the course grade aimed at introducing the student’s disciplinary DB & KB for the project. See samples of DB & KB presentations in course exhibit.

Key Assignment for Outcome I: According to the overall project time plan, the students are required to document their individual working DB & KB in two ways: (a) bi-weekly report, and (b) DB & KB update presentation.

Not only does a project of this magnitude require an extensive research, the research continue through most of the project. Many smaller items may be overlooked when a project of this scale begins, and additional research is necessary to fill the gaps. This particular project highlighted the need to move out of the comfort zone of aeronautical engineering and into other aspects, rocket sizing and modeling of solar irradiance are two examples of this. Many future projects in the professional world require that engineers work outside their focused skill set and find solutions to problems that will require attaining new knowledge.

IX. ConclusionFrom the above analysis, it can be observed that not only is flight on Mars possible, but can be a major asset to a

Martian mission. After several iterations, a solution was produced that offered a reasonable takeoff ground roll, and long range. These characteristics will allow the Daedalus to be an integral part of the exploration of Mars. The Daedalus represents a huge leap in the potential for exploring Mars.




Appendix

Table A-1: Normalized Net Irradiation for Surface Albedo of 0.110

albedo 0.1 zenith angle0 10 20 30 40 50 60 70 80 85

0.1 0.883 0.883 0.881 0.879 0.875 0.868 0.855 0.83 0.757 0.640.2 0.866 0.865 0.862 0.857 0.85 0.836 0.813 0.768 0.651 0.5080.3 0.848 0.847 0.842 0.835 0.826 0.8016 0.773 0.712 0.571 0.4330.4 0.83 0.829 0.823 0.814 0.801 0.776 0.736 0.663 0.51 0.3850.5 0.813 0.811 0.804 0.793 0.778 0.748 0.701 0.619 0.462 0.3510.6 0.796 0.793 0.785 0.772 0.755 0.721 0.668 0.579 0.422 0.3250.7 0.778 0.775 0.766 0.752 0.732 0.695 0.638 0.546 0.393 0.3060.8 0.761 0.758 0.748 0.732 0.711 0.67 0.609 0.514 0.366 0.2880.9 0.744 0.741 0.73 0.713 0.69 0.646 0.583 0.486 0.345 0.274

1 0.728 0.724 0.712 0.694 0.669 0.623 0.557 0.459 0.324 0.2611.1 0.712 0.708 0.696 0.676 0.649 0.602 0.535 0.438 0.309 0.251.2 0.695 0.691 0.679 0.658 0.63 0.581 0.513 0.416 0.295 0.241.3 0.679 0.675 0.662 0.64 0.611 0.56 0.491 0.395 0.28 0.231.4 0.664 0.66 0.646 0.624 0.594 0.542 0.473 0.379 0.27 0.2221.5 0.649 0.644 0.63 0.607 0.576 0.524 0.455 0.364 0.26 0.2151.6 0.634 0.629 0.615 0.591 0.559 0.506 0.437 0.348 0.249 0.2071.7 0.619 0.614 0.599 0.575 0.542 0.489 0.42 0.333 0.24 0.21.8 0.605 0.6 0.585 0.561 0.527 0.475 0.406 0.321 0.232 0.1941.9 0.591 0.586 0.571 0.546 0.513 0.46 0.392 0.31 0.225 0.188

2 0.578 0.572 0.557 0.532 0.498 0.445 0.377 0.298 0.217 0.1822.25 0.544 0.539 0.523 0.498 0.463 0.411 0.346 0.272 0.2 0.168

2.5 0.514 0.509 0.493 0.467 0.432 0.381 0.319 0.252 0.187 0.1572.75 0.483 0.478 0.462 0.436 0.401 0.352 0.293 0.231 0.173 0.146

3 0.457 0.452 0.436 0.411 0.376 0.329 0.273 0.216 0.162 0.1373.25 0.431 0.426 0.411 0.386 0.352 0.307 0.254 0.202 0.152 0.129

3.5 0.405 0.4 0.385 0.361 0.328 0.285 0.236 0.187 0.142 0.124 0.363 0.358 0.344 0.321 0.29 0.251 0.208 166 0.127 0.1075 0.289 0.285 0.273 0.253 0.227 0.196 0.162 0.131 0.101 0.0866 0.229 0.226 0.216 0.178 0.178 0.154 0.128 0.104 0.081 0.069




Table A-2: Normalized Net Irradiation for Surface Albedo of 0.410

albedo 0.4 zenith angle0 10 20 30 40 50 60 70 80 85

0.1 0.594 0.593 0.592 0.591 0.588 0.583 0.575 0.558 0.509 0.430.2 0.586 0.585 0.583 0.58 0.576 0.566 0.551 0.52 0.441 0.3440.3 0.578 0.577 0.574 0.569 0.562 0.549 0.527 0.485 0.389 0.2950.4 0.569 0.567 0.564 0.557 0.549 0.532 0.504 0.454 0.35 0.2640.5 0.56 0.558 0.553 0.546 0.535 0.515 0.483 0.426 0.318 0.2420.6 0.55 0.548 0.543 0.534 0.522 0.498 0.462 0.401 0.292 0.2250.7 0.54 0.538 0.532 0.522 0.509 0.483 0.443 0.379 0.273 0.2130.8 0.531 0.528 0.522 0.511 0.496 0.467 0.425 0.358 0.256 0.2020.9 0.521 0.518 0.511 0.499 0.483 0.452 0.408 0.341 0.242 0.193

1 0.511 0.509 0.501 0.488 0.47 0.438 0.392 0.323 0.228 0.1841.1 0.501 0.499 0.49 0.476 0.458 0.424 0.377 0.309 0.219 0.1771.2 0.492 0.489 0.48 0.465 0.446 0.411 0.363 0.295 0.209 0.1711.3 0.482 0.479 0.469 0.454 0.433 0.398 0.349 0.281 0.2 0.1641.4 0.472 0.469 0.459 0.444 0.422 0.386 0.337 0.271 0.193 0.1591.5 0.463 0.46 0.45 0.433 0.411 0.374 0.325 0.26 0.186 0.1541.6 0.453 0.45 0.44 0.423 0.4 0.363 0.313 0.25 0.18 0.1491.7 0.444 0.44 0.43 0.413 0.389 0.352 0.302 0.24 0.173 0.1441.8 0.435 0.431 0.421 0.403 0.379 0.342 0.292 0.232 0.168 0.141.9 0.426 422 0.411 0.394 0.37 0.332 0.283 0.224 0.163 0.136

2 0.417 0.413 0.402 0.384 0.36 0.322 0.273 0.217 0.158 0.1332.25 0.395 0.391 0.38 0.362 0.336 0.299 0.252 0.199 0.147 0.124

2.5 0.375 0.371 0.36 0.341 0.316 0.279 0.234 0.185 0.138 0.1162.75 0.354 0.35 0.339 0.32 0.295 0.259 0.216 0.171 0.128 0.108

3 0.336 0.332 0.321 0.303 0.277 0.243 0.202 0.161 0.121 0.1023.25 0.319 0.315 0.304 0.285 0.261 0.228 0.189 0.151 0.114 0.096

3.5 0.301 0.297 0.286 0.268 0.244 0.212 0.176 0.141 0.107 0.0914 0.271 0.267 0.257 0.24 0.217 0.189 0.157 0.125 0.096 0.0825 0.219 0.216 0.207 0.192 0.173 0.149 0.124 0.101 0.078 0.0666 0.176 0.173 0.166 0.154 0.137 0.119 0.099 0.081 0.063 0.054




Table A-3: Normalized Net Irradiation for Surface Albedo of 0.25albedo 0.25 zenith angle

0 10 20 30 40 50 60 70 80 850.1 0.7385 0.738 0.7365 0.735 0.7315 0.7255 0.715 0.694 0.633 0.5350.2 0.726 0.725 0.7225 0.7185 0.713 0.701 0.682 0.644 0.546 0.4260.3 0.713 0.712 0.708 0.702 0.694 0.6753 0.65 0.5985 0.48 0.3640.4 0.6995 0.698 0.6935 0.6855 0.675 0.654 0.62 0.5585 0.43 0.32450.5 0.6865 0.6845 0.6785 0.6695 0.6565 0.6315 0.592 0.5225 0.39 0.29650.6 0.673 0.6705 0.664 0.653 0.6385 0.6095 0.565 0.49 0.357 0.2750.7 0.659 0.6565 0.649 0.637 0.6205 0.589 0.5405 0.4625 0.333 0.25950.8 0.646 0.643 0.635 0.6215 0.6035 0.5685 0.517 0.436 0.311 0.2450.9 0.6325 0.6295 0.6205 0.606 0.5865 0.549 0.4955 0.4135 0.2935 0.2335

1 0.6195 0.6165 0.6065 0.591 0.5695 0.5305 0.4745 0.391 0.276 0.22251.1 0.6065 0.6035 0.593 0.576 0.5535 0.513 0.456 0.3735 0.264 0.21351.2 0.5935 0.59 0.5795 0.5615 0.538 0.496 0.438 0.3555 0.252 0.20551.3 0.5805 0.577 0.5655 0.547 0.522 0.479 0.42 0.338 0.24 0.1971.4 0.568 0.5645 0.5525 0.534 0.508 0.464 0.405 0.325 0.2315 0.19051.5 0.556 0.552 0.54 0.52 0.4935 0.449 0.39 0.312 0.223 0.18451.6 0.5435 0.5395 0.5275 0.507 0.4795 0.4345 0.375 0.299 0.2145 0.1781.7 0.5315 0.527 0.5145 0.494 0.4655 0.4205 0.361 0.2865 0.2065 0.1721.8 0.52 0.5155 0.503 0.482 0.453 0.4085 0.349 0.2765 0.2 0.1671.9 0.5085 211.293 0.491 0.47 0.4415 0.396 0.3375 0.267 0.194 0.162

2 0.4975 0.4925 0.4795 0.458 0.429 0.3835 0.325 0.2575 0.1875 0.15752.25 0.4695 0.465 0.4515 0.43 0.3995 0.355 0.299 0.2355 0.1735 0.146

2.5 0.4445 0.44 0.4265 0.404 0.374 0.33 0.2765 0.2185 0.1625 0.13652.75 0.4185 0.414 0.4005 0.378 0.348 0.3055 0.2545 0.201 0.1505 0.127

3 0.3965 0.392 0.3785 0.357 0.3265 0.286 0.2375 0.1885 0.1415 0.11953.25 0.375 0.3705 0.3575 0.3355 0.3065 0.2675 0.2215 0.1765 0.133 0.1125

3.5 0.353 0.3485 0.3355 0.3145 0.286 0.2485 0.206 0.164 0.1245 0.10554 0.317 0.3125 0.3005 0.2805 0.2535 0.22 0.1825 83.0625 0.1115 0.09455 0.254 0.2505 0.24 0.2225 0.2 0.1725 0.143 0.116 0.0895 0.0766 0.2025 0.1995 0.191 0.166 0.1575 0.1365 0.1135 0.0925 0.072 0.0615




Table A-4: Mars Bulk Parameters13

Table A-5: Mars Orbital Parameters13




Table A-6: Mars Observational Parameters13

Table A-7: Martian Atmospheric Characteristics13

Table A-8: Historical Log of Mars Missions14

# Launch Date Name Country Result Reason

1 1960 Korabl 4 USSR (flyby) Failure Didn't reach Earth orbit

2 1960 Korabl 5 USSR (flyby) Failure Didn't reach Earth orbit

3 1962 Korabl 11 USSR (flyby) Failure Earth orbit only; spacecraft broke apart

4 1962 Mars 1 USSR (flyby) Failure Radio Failed

5 1962 Korabl 13 USSR (flyby) Failure Earth orbit only; spacecraft broke apart

6 1964 Mariner 3 US (flyby) Failure Shroud failed to jettison

7 1964 Mariner 4 US (flyby) Success Returned 21 images

8 1964 Zond 2 USSR (flyby) Failure Radio failed

9 1969 Mars 1969A USSR Failure Launch vehicle failure

10 1969 Mars 1969B USSR Failure Launch vehicle failure






13 1971 Mariner 8 US Failure Launch failure

14 1971 Kosmos 419 USSR Failure Achieved Earth orbit only

15 1971 Mars 2 Orbiter/Lander

USSR Failure Orbiter arrived, but no useful data and Lander destroyed


USSR Success Orbiter obtained approximately 8 months of data and lander landed safely, but only 20 seconds of data

17 1971 Mariner 9 US Success Returned 7,329 images

18 1973 Mars 4 USSR Failure Flew past Mars

19 1973 Mars 5 USSR Success Returned 60 images; only lasted 9 days


USSR Success/Failure Occultation experiment produced data and Lander failure on descent

21 1973 Mars 7 Lander USSR Failure Missed planet; now in solar orbit.

22 1975 Viking 1 Orbiter/Lander

US Success Located landing site for Lander and first successful landing on Mars

23 1975 Viking 2 Orbiter/Lander

US Success Returned 16,000 images and extensive atmospheric data and soil experiments

24 1988 Phobos 1 Orbiter USSR Failure Lost en route to Mars

25 1988 Phobos 2 Orbiter/Lander

USSR Failure Lost near Phobos

26 1992 Mars Observer US Failure Lost prior to Mars arrival

27 1996 Mars Global Surveyor

US Success More images than all Mars Missions

28 1996 Mars 96 Russia Failure Launch vehicle failure

29 1996 Mars Pathfinder US Success Technology experiment lasting 5 times longer than warranty

30 1998 Nozomi Japan Failure No orbit insertion; fuel problems

31 1998 Mars Climate Orbiter US Failure Lost on arrival

32 1999 Mars Polar Lander US Failure Lost on arrival

33 1999 Deep Space 2 Probes (2)

US Failure Lost on arrival (carried on Mars Polar Lander)

34 2001 Mars Odyssey US Success High resolution images of Mars

35 2003 Mars Express Orbiter/Beagle 2

ESA Success/Failure Orbiter imaging Mars in detail and lander lost on arrival




Lander

36 2003 Mars Exploration Rover - Spirit

US Success Operating lifetime of more than 15 times original warranty

37 2003 Mars Exploration Rover - Opportunity

US Success Operating lifetime of more than 15 times original warranty

38 2005 Mars Reconnaissance Orbiter

US Success Returned more than 26 terabits of data (more than all other Mars missions combined)

39 2007 Phoenix Mars Lander US Success Returned more than 25 gigabits of data

40 2011 Mars Science Laboratory

US Success Exploring Mars' habitability

41 2011 Phobos-Grunt/Yinghuo-1

Russia/China Failure Stranded in Earth orbit

42 2013 Mars Atmosphere and Volatile Evolution

US Success Studying the Martian atmosphere

43 2013 Mars Orbiter Mission (MOM)

India Success Develop interplanetary technologies and explore Mars' surface features, mineralogy and atmosphere




Table A-9: One-Way Mars Missions19

Earth_Departure Dest_Arrival Duration (days) C3 (km2/s2)Injection DV

(km/s)Post-injection DV

(km/s)Total DV (km/s)

Route

Jun-14-2025 Feb-12-2027 608 29.1 4.48 5.37 9.84 EMJun-30-2025 Mar-16-2027 624 30.4 4.53 5.11 9.64 EMJul-16-2025 Apr-17-2027 640 32.2 4.6 4.91 9.52 EMJul-16-2025 May-03-2027 656 31.3 4.57 4.8 9.37 EMAug-01-2025 Jun-04-2027 672 32.8 4.63 4.68 9.3 EMAug-01-2025 Jun-20-2027 688 32.3 4.6 4.62 9.23 EMFeb-25-2026 Aug-07-2027 528 56.2 5.54 2.41 7.94 EVMMar-13-2026 Dec-29-2027 656 11.9 3.76 1.4 5.15 EMMar-29-2026 Nov-11-2027 592 43.1 5.04 2.7 7.73 EMMar-29-2026 Dec-13-2027 624 18.7 4.04 2.07 6.11 EMMar-29-2026 Dec-29-2027 640 11.5 3.73 1.93 5.66 EMMar-29-2026 Nov-27-2027 608 29.2 4.48 2.33 6.8 EMJun-17-2026 May-05-2028 688 9.9 3.66 1.15 4.81 EMJun-17-2026 Apr-19-2028 672 11.2 3.72 1.35 5.07 EMJul-03-2026 Jun-22-2028 720 10.3 3.69 0.82 4.51 EMJul-03-2026 Jun-06-2028 704 10.8 3.7 0.92 4.63 EMOct-23-2026 Aug-07-2027 288 8.9 3.62 0.81 4.44 EMNov-08-2026 Sep-08-2027 304 9 3.63 0.62 4.25 EMNov-24-2026 Aug-07-2027 256 13.8 3.84 0.85 4.68 EMNov-24-2026 Jul-06-2027 224 14.7 3.88 1.42 5.3 EMNov-24-2026 Jul-22-2027 240 14.3 3.86 1.06 4.92 EMDec-10-2026 Jul-06-2027 208 26.1 4.36 1.45 5.81 EMDec-10-2026 Jun-20-2027 192 26.5 4.37 1.98 6.35 EMDec-10-2026 Jun-04-2027 176 27 4.39 2.73 7.12 EMDec-26-2026 May-19-2027 144 52.6 5.4 3.86 9.26 EMDec-26-2026 Jun-04-2027 160 50.7 5.33 2.79 8.12 EMJun-20-2027 Dec-31-2028 560 26.3 4.36 4.71 9.07 EMJul-06-2027 Feb-01-2029 576 27.7 4.42 4.51 8.92 EMJul-06-2027 Feb-17-2029 592 27.6 4.42 4.41 8.83 EMJul-22-2027 Apr-06-2029 624 28.9 4.47 4.22 8.68 EMJul-22-2027 Mar-21-2029 608 28.7 4.46 4.27 8.73 EMAug-07-2027 May-08-2029 640 29.4 4.49 4.15 8.64 EMDec-29-2027 Mar-21-2029 448 50 5.3 2.29 7.59 EVMJan-14-2028 Mar-21-2029 432 30.5 4.53 3.21 7.74 EVMJan-14-2028 Mar-05-2029 416 30.5 4.53 3.51 8.05 EVMJan-30-2028 Jan-03-2030 704 20.9 4.14 1.49 5.63 EMMar-02-2028 Jul-11-2029 496 10.5 3.69 3.58 7.28 EVMMar-18-2028 Mar-08-2030 720 18.9 4.05 1.24 5.29 EMMay-21-2028 Dec-18-2029 576 34.5 4.69 1.94 6.63 EMMay-21-2028 Jan-03-2030 592 28.2 4.44 1.52 5.96 EMJun-06-2028 Feb-04-2030 608 24.8 4.3 1.27 5.57 EMJun-22-2028 Mar-08-2030 624 26.1 4.35 1.15 5.5 EMNov-29-2028 Aug-28-2029 272 10.9 3.71 1.01 4.72 EMNov-29-2028 Sep-29-2029 304 8.2 3.59 0.85 4.45 EMNov-29-2028 Sep-13-2029 288 8.8 3.62 0.85 4.47 EMDec-31-2028 Aug-28-2029 240 15 3.89 1.29 5.17 EMDec-31-2028 Aug-12-2029 224 13.2 3.81 1.39 5.2 EMDec-31-2028 Jul-27-2029 208 12.9 3.8 1.78 5.58 EMJan-16-2029 Jul-27-2029 192 23.9 4.26 1.79 6.06 EMJan-16-2029 Jul-11-2029 176 24.3 4.28 2.45 6.73 EMFeb-01-2029 Jun-25-2029 144 48 5.22 3.5 8.73 EMFeb-01-2029 Jul-11-2029 160 46 5.15 2.52 7.67 EMJun-25-2029 Nov-19-2030 512 23.4 4.24 3.97 8.21 EMJul-11-2029 Dec-21-2030 528 24.7 4.3 3.82 8.12 EMJul-11-2029 Jan-06-2031 544 24.6 4.29 3.76 8.05 EMJul-27-2029 Feb-23-2031 576 25.8 4.34 3.64 7.98 EMJul-27-2029 Feb-07-2031 560 25.8 4.34 3.67 8.01 EM




Table A-10: Mars Missions with Return Trajectories19

Earth_Departure Dest_Arrival Dest_Departure Earth_ReturnStay time

(days)Duration

(days)C3

(km2/s2)Injection DV

(km/s)Post-injection DV

(km/s)Total DV (km/s)

Nov-08-2026 Sep-08-2027 Sep-10-2028 Aug-12-2029 368 1008 9 3.63 1.21 4.84Nov-08-2026 Sep-08-2027 Sep-10-2028 May-24-2029 368 928 9 3.63 1.28 4.91Nov-08-2026 Sep-08-2027 Aug-25-2028 Jul-27-2029 352 992 9 3.63 1.22 4.85Nov-08-2026 Sep-08-2027 Sep-10-2028 Jun-09-2029 368 944 9 3.63 1.27 4.9Nov-08-2026 Sep-08-2027 Sep-10-2028 May-08-2029 368 912 9 3.63 1.51 5.14Nov-10-2026 Jul-08-2027 Aug-07-2027 Feb-28-2028 30 475 11.3 3.73 6.01 9.74Nov-10-2026 Jul-13-2027 Aug-12-2027 Mar-04-2028 30 480 11.2 3.72 5.99 9.71Nov-15-2026 Jul-08-2027 Aug-07-2027 Feb-28-2028 30 470 11.8 3.75 6.02 9.77Nov-15-2026 Jul-03-2027 Aug-02-2027 Feb-23-2028 30 465 12 3.76 6.08 9.84Nov-20-2026 Jul-03-2027 Aug-02-2027 Feb-23-2028 30 460 13.3 3.82 6.1 9.91Nov-24-2026 Oct-10-2027 Sep-10-2028 May-08-2029 336 896 13.7 3.83 1.72 5.56Dec-10-2026 Aug-23-2027 Sep-10-2028 May-08-2029 384 880 23.9 4.26 1.72 5.99Dec-10-2026 Aug-23-2027 Sep-26-2028 Apr-22-2029 400 864 23.9 4.26 2.16 6.43Dec-10-2026 Aug-23-2027 Sep-26-2028 Apr-06-2029 400 848 23.9 4.26 2.93 7.19Dec-26-2026 Sep-08-2027 Aug-25-2028 Mar-21-2029 352 816 39.7 4.9 4.28 9.18Dec-26-2026 Sep-08-2027 Sep-26-2028 Apr-06-2029 384 832 39.7 4.9 3.13 8.03Feb-12-2027 Apr-03-2028 Sep-26-2028 Apr-22-2029 176 800 16 3.93 6.04 9.97Nov-14-2028 Aug-11-2029 Sep-10-2029 Apr-08-2030 30 510 9.8 3.66 5.48 9.14Nov-14-2028 Aug-16-2029 Sep-15-2029 Apr-13-2030 30 515 9.5 3.65 5.49 9.14Nov-14-2028 Jul-27-2029 Aug-26-2029 Mar-24-2030 30 495 11 3.72 5.58 9.3Nov-14-2028 Aug-01-2029 Aug-31-2029 Mar-29-2030 30 500 10.5 3.69 5.53 9.22Nov-14-2028 Aug-06-2029 Sep-05-2029 Apr-03-2030 30 505 10.1 3.68 5.49 9.17Nov-14-2028 Jul-22-2029 Aug-21-2029 Mar-19-2030 30 490 11.7 3.74 5.66 9.4Nov-29-2028 Sep-29-2029 Nov-03-2030 Sep-19-2031 400 1024 8.2 3.59 1.37 4.97Nov-29-2028 Sep-29-2029 Nov-03-2030 Jun-15-2031 400 928 8.2 3.59 1.57 5.16Nov-29-2028 Sep-29-2029 Nov-03-2030 Jul-01-2031 400 944 8.2 3.59 1.42 5.01Nov-29-2028 Sep-29-2029 Oct-02-2030 Sep-03-2031 368 1008 8.2 3.59 1.41 5.01Dec-15-2028 Oct-15-2029 Nov-03-2030 Jun-15-2031 384 912 10.1 3.68 1.75 5.43Dec-15-2028 Oct-15-2029 Nov-03-2030 May-30-2031 384 896 10.1 3.68 2.11 5.78Dec-19-2028 Aug-01-2029 Aug-31-2029 Apr-03-2030 30 470 9.2 3.64 5.78 9.41Dec-19-2028 Aug-01-2029 Aug-31-2029 Mar-29-2030 30 465 9.2 3.64 5.8 9.43Dec-24-2028 Aug-01-2029 Aug-31-2029 Mar-29-2030 30 460 10.4 3.69 5.78 9.47Dec-24-2028 Jul-27-2029 Aug-26-2029 Mar-24-2030 30 455 10.3 3.68 5.84 9.52Dec-29-2028 Jul-27-2029 Aug-26-2029 Mar-24-2030 30 450 12 3.76 5.84 9.6Dec-29-2028 Jul-22-2029 Aug-21-2029 Mar-19-2030 30 445 12.1 3.76 5.93 9.69Jan-03-2029 Jul-17-2029 Aug-16-2029 Mar-14-2030 30 435 14.4 3.86 6.06 9.92Jan-03-2029 Jul-22-2029 Aug-21-2029 Mar-19-2030 30 440 14.3 3.86 5.93 9.79Jan-16-2029 Aug-28-2029 Nov-03-2030 Jun-15-2031 432 880 23 4.22 1.9 6.13Jan-16-2029 Aug-28-2029 Oct-18-2030 May-14-2031 416 848 23 4.22 2.91 7.13Jan-16-2029 Aug-28-2029 Nov-03-2030 May-30-2031 432 864 23 4.22 2.26 6.48Feb-01-2029 Sep-13-2029 Oct-18-2030 May-14-2031 400 832 37.4 4.81 3.01 7.82Feb-01-2029 Sep-13-2029 Sep-16-2030 Apr-28-2031 368 816 37.4 4.81 3.97 8.79Apr-22-2029 Jun-12-2030 Nov-03-2030 Jul-01-2031 144 800 14.9 3.88 5.21 9.1May-08-2029 Jul-14-2030 Nov-03-2030 Jul-01-2031 112 784 17 3.97 5.16 9.13May-24-2029 Aug-15-2030 Dec-05-2030 May-30-2031 112 736 19.7 4.09 5.68 9.77May-24-2029 Aug-15-2030 Dec-05-2030 Jul-01-2031 112 768 19.7 4.09 5.19 9.27May-24-2029 Aug-15-2030 Dec-05-2030 Jun-15-2031 112 752 19.7 4.09 5.33 9.41




Table A-11: Dune Field Distribution by Latitude20

Table A-12: Dune Field Distribution in Craters20




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29Ferguson, Zackery, Mars Flight Vehicle: The Daedalus-Aerodynamics, Senior Design Capstone Report, Mechanical and Aerospace Engineering Department, University of Texas at Arlington, Arlington, TX 2016

30IRVE and IXV-NASA’s and ESA’s New Reentry Vehicles, Exo Cruiser. Web 8/1/2016, http://dodlithr.blogspot.com/2013/10/irve-and-ixv-nasas-and-esas-new-re.html

31Google Maps-Radius. Web 8/15/2016. http://obeattie.github.io/gmaps-radius/?lat=32.633706&lng=-96.847194&z=8&u=km&r=183.7


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