salas nunez, luis, mars aerial platform, chief engineer, final report (final)

SENIOR DESIGN: MAE 4351 Project

Ref.: MAE 4351-001-2016 Date: 25. Aug. 2016 Page: 1 of 127 Pages Status: In Progress

The University of Texas at Arlington

MARS AERIAL PLATFORM

CHIEF ENGINEER REPORT

Signatures:

Name: Signature: Dept.: Date:

Author: Luis Salas Nunez MAE 8/25/2016

Seen: Dr. Bernd Chudoba MAE

Summary:

The following is the Chief Engineer report for the group Aurora Concepts. The report summarizes the work performed on the conceptual design of the Mars Aerial Research and Environmental Laboratory (MARVEL). Such is an aircraft capable of flying on Mars while carrying a human plus its equipment as payload (250 kg total). The aircraft has been designed to have a high aspect ratio wing, with a wing area of 110m2, a mass of 1467 kg, and a power of 60 kW. The aircraft uses three hydrazine rockets (~10kN of thrust) for its takeoff and landing phases, whereas cruise flight is powered by two battery-driven propellers. Operationally, the aircraft is capable of flying on the speed range between 65 to 125 m/s, in altitudes from 0 to 2700 m on Mars, on locations where little to no infrastructure is available. Given its large dimensions, the aircraft would have to be launched on the Space Launch System, Block 2B, and fitted into an entry, descent, and landing system for later deployment on the Martian surface.

This report presents an extensive literature review on Mars, space missions, and previous aircraft concepts. This project has taken inspiration from concepts such as the Astroplane (1978) and, more recently, the ARES (2003). There is also a detailed analysis of the business case that an aircraft such as the MARVEL would provide for companies such as Space X or governmental organizations like NASA. It is important to highlight that this project was conceived taking into account the current limitations imposed by the current industry capability, which is critical in the areas of space travel and propulsion systems.

There is also an extensive analysis on mission, design, and technology trade studies that are pertinent to this design project. This led to a multi-disciplinary analysis for the MARVEL on the areas of aerodynamics, structures, propulsion, energy, geometry, performance, stability, and control. A conceptual design methodology was created and implemented to allow the individual design disciplines to arrive to a converged design, the MARVEL. The most important parameters of such disciplines have been summarized on this report and are used to explain how the resulting aircraft meets the mission requirements.

Distribution: Institution: Dept.: Name:


MAE

Dr. Bernd Chudoba




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: 8/25/2016

Signature: Luis Salas Nunez




Table of Contents Work Disclosure Statement .......................................................................................................................... 2

Table of Contents .......................................................................................................................................... 3

List of Figures ............................................................................................................................................... 6

List of Tables ................................................................................................................................................ 9

Nomenclature .............................................................................................................................................. 10

I. Introduction ......................................................................................................................................... 13

A. Project Overview ............................................................................................................................. 13

B. Request for Proposal ....................................................................................................................... 13

C. Motivations ..................................................................................................................................... 15

D. Business Case .................................................................................................................................. 16

II. Team Organization .............................................................................................................................. 18

A. Team Structure ................................................................................................................................ 18

B. Activity Log .................................................................................................................................... 19

C. Gantt Chart ...................................................................................................................................... 22

III. Literature Review ............................................................................................................................ 24

A. Mars Description ............................................................................................................................. 24

B. The Mars Project ............................................................................................................................. 25

1. History of Missions to Mars ........................................................................................................ 25

2. The Future of Mars Exploration .................................................................................................. 26

C. Previous Mars Aerial Platforms Concepts ...................................................................................... 27

1. Fixed Wing .................................................................................................................................. 28

2. Inflatable Wings .......................................................................................................................... 34

3. Rotorcraft .................................................................................................................................... 35

IV. Methodology ................................................................................................................................... 36

A. Multi-Disciplinary Analysis (MDA) ............................................................................................... 36

B. Individual Responsibilities and Scope ............................................................................................ 39

V. Conceptual Design Analysis ............................................................................................................... 41

A. Mission Studies ............................................................................................................................... 41

1. Transportation and Delivery to Mars .......................................................................................... 41

2. Mission Specifications ................................................................................................................ 43

B. Design Studies ................................................................................................................................. 44

1. Initial Considerations .................................................................................................................. 45




2. Parametric Sizing ........................................................................................................................ 46

C. Technology Studies ......................................................................................................................... 53

1. Propulsion Systems for the Mars Aircraft ................................................................................... 53

2. Rocket-Assisted Takeoff-Landing .............................................................................................. 58

VI. Design Disciplines .......................................................................................................................... 60

A. Synthesis ......................................................................................................................................... 60

1. Aircraft Configuration Highlights ............................................................................................... 60

2. Cost Analysis .............................................................................................................................. 61

B. CAD/Geometry ............................................................................................................................... 63

C. Aerodynamics ................................................................................................................................. 65

1. Wing Characteristics ................................................................................................................... 65

2. Full Body Aerodynamics ............................................................................................................ 67

D. Propulsion and Energy .................................................................................................................... 68

1. Vertical Takeoff-Landing Analysis ............................................................................................ 68

2. Cruise and Maneuvering Flight ................................................................................................... 70

3. Power Plant ................................................................................................................................. 71

4. Drivetrain .................................................................................................................................... 72

E. Structures and Weights ................................................................................................................... 72

1. Fuselage Shape ............................................................................................................................ 72

2. Mass Estimation .......................................................................................................................... 73

3. Load Estimation .......................................................................................................................... 74

F. Stability and Controls ...................................................................................................................... 75

1. Tail Configuration ....................................................................................................................... 75

2. Control Surface Sizing ................................................................................................................ 75

3. VTOL Stability Analysis ............................................................................................................ 77

4. Static Stability Analysis .............................................................................................................. 77

G. Performance .................................................................................................................................... 77

1. Takeoff and Landing ................................................................................................................... 77

2. Cruise and Maneuvering Flight ................................................................................................... 78

3. Operational Flight Envelope and Mission Analysis ................................................................... 80

VII. Discussion ....................................................................................................................................... 82

VIII. ABET Requirements ....................................................................................................................... 83

A. Outcome C: Design System, Component or Process to Meet Needs .............................................. 83




B. Outcome D: Function on Multidisciplinary Teams ........................................................................ 83

C. Outcome F: Understand Professional & Ethical Responsibility ..................................................... 84

D. Outcome G: Communicate Effectively ........................................................................................... 84

E. Outcome H: Understand Impact of Engineering Solutions in Global & Societal Context ............. 85

F. Outcome I: Recognize the Need & Ability to Engage in Lifelong Learning ................................. 85

IX. Conclusion ...................................................................................................................................... 86

X. Appendix ............................................................................................................................................. 87

A. Loftin’s Design Methodology, Propeller-Driven Aircraft .............................................................. 87

1. Aircraft Speed Prediction ............................................................................................................ 87

2. Airport Performance ................................................................................................................... 88

3. Climb Performance ..................................................................................................................... 91

4. Matching Procedure .................................................................................................................... 93

5. Weight, Range and Fuel Fraction ............................................................................................... 94

B. Databases ........................................................................................................................................ 96

C. Disciplines Relevant Figures .......................................................................................................... 98

1. Aerodynamics ............................................................................................................................. 98

2. Propulsion and Energy ................................................................................................................ 99

D. Mars Aircraft Concept Information Tables ................................................................................... 100

E. Individual Discipline Analysis Diagrams ..................................................................................... 107

F. MATLAB Scripts .......................................................................................................................... 109

1. Loftin’s Parametric Sizing Method ........................................................................................... 109

1. Vertical Takeoff/Landing Analysis [53] ................................................................................... 111

2. Thrust and Power Requirements ............................................................................................... 118

3. Cost Estimation Function .......................................................................................................... 119

XI. Acknowledgments ......................................................................................................................... 120

XII. Bibliography .................................................................................................................................. 120




List of Figures Figure 1. Si2 3D CAD Model [1] ............................................................................................................... 13

Figure 2. Mars Platform Comparison Chart [4] .......................................................................................... 14

Figure 3. SpaceX’s Falcon 9 Rocket Vertical Landing on Drone Ship [14] .............................................. 16

Figure 4. Space Launch System Payload Volume Comparison [18] .......................................................... 17

Figure 5. Team Structure Chart ................................................................................................................... 18

Figure 6. "Mars Direct" Base Camp Concept [29] ..................................................................................... 27

Figure 7. DSI Cruiser Configuration Layout [31] ....................................................................................... 29

Figure 8. DSI Landing Maneuver [31] ........................................................................................................ 29

Figure 9. DSI Takeoff Maneuver [31] ........................................................................................................ 29

Figure 10. Artistic Representation of Long-Endurance Aircraft [34] ......................................................... 30

Figure 11. ARES Final Configuration (left) and Prototype [7] .................................................................. 31

Figure 12. Examples of Marsplane Resulting Configurations [41] ............................................................ 32

Figure 13. Zephyr Layout and Propeller [42] ............................................................................................. 33

Figure 14. Prototype of Aircraft with Inflatable Wings [44] ...................................................................... 34

Figure 15. MARV Design [48] ................................................................................................................... 35

Figure 16. Parametric Sizing Diagram Overview ....................................................................................... 37

Figure 17. Detailed MDA Diagram ............................................................................................................ 38

Figure 18. IDA Template ............................................................................................................................ 39

Figure 19. CAD Model of SIAD [54] ......................................................................................................... 42

Figure 20. Lazair Ultralight Airplane [62] .................................................................................................. 46

Figure 21. Loftin Sizing Process Diagram .................................................................................................. 48

Figure 22. Loftin-based Preliminary Matching Chart ................................................................................. 49

Figure 23. Max. Speed vs Power Index for Light Aircraft ......................................................................... 50

Figure 24. Useful Load Fraction vs. Power Loading for Light Aircraft ..................................................... 50

Figure 25. Loftin-adapted Final Matching Chart ........................................................................................ 51

Figure 26. Mass Breakdown of Combustion Engines for an Endurance of 2 Hours [36] .......................... 55

Figure 27. Rocket-Assisted Launch Fighers, F-100 (left) and F-84 (right) [68] [69] ................................. 58

Figure 28. NASA's Flying Bedstead and Rolls-Royce's Pegasus Engine [72] [73] ................................... 59

Figure 29. Isometric View of MARVEL .................................................................................................... 60

Figure 30. Cost per Pound for Different Rockets and Launch Frequencies [79] ........................................ 62

Figure 31. MARVEL Fuselage ................................................................................................................... 64

Figure 32. MARVEL Cabin ........................................................................................................................ 65




Figure 33. MARVEL Wing ........................................................................................................................ 65

Figure 34. MARVEL Empennage .............................................................................................................. 65

Figure 35. Variation of Lift for Various Re and AOA for the S1223 and 63-137 Airfoils [4] ................... 66

Figure 36. FX 63-137 Airfoil Shape [4] ..................................................................................................... 66

Figure 37. MARVEL Drag Polar, Analytical and Computational Results [4] ........................................... 67

Figure 38. MARVEL Lift Coefficient vs. AOA [4] ................................................................................... 67

Figure 39. MARVEL Span Loading [4] ..................................................................................................... 67

Figure 40. MARVEL Lift-to-Drag [4] ........................................................................................................ 68

Figure 41. MARVEL Pitching Moment [4] ................................................................................................ 68

Figure 42. Rocket-Assisted Landing and Takeoff Profiles [63] ................................................................. 69

Figure 43. Thrust and Power Requirements for Different Altitudes at Mars ............................................. 70

Figure 44. Tapdole CAD Model [17] .......................................................................................................... 73

Figure 45. Aircraft Mass Distribution [17] ................................................................................................. 74

Figure 46. Wing Loading Distribution [17] ................................................................................................ 74

Figure 47. Shear and Bending Moment Distribution [17] .......................................................................... 74

Figure 48. Piper Super Cub [86] [87] ......................................................................................................... 76

Figure 49. Rocket-Assisted Takeoff and Landing Profiles [63] ................................................................. 78

Figure 50. Thrust Analysis at Cruise Altitude (1000 m) [89] ..................................................................... 79

Figure 51. Power Analysis at Cruise Altitude (1000 m) [89] ..................................................................... 79

Figure 52. Range and Endurance at Cruise Altitude (1000 m) [89] ........................................................... 79

Figure 53. MARVEL Flight Envelope [89] ................................................................................................ 80

Figure 54. MARVEL Mission Scenario [89] .............................................................................................. 81

Figure 55. Variation of Aircraft Maximum Speed with Power Index for Class I Airplanes [51] .............. 87

Figure 56. Boundaries of Speed with Wing Loading Parameter [51] ......................................................... 88

Figure 57. Variation of Stall Speed with Wing Loading Parameter, flaps-up/down [51] .......................... 89

Figure 58. Landing Ground Run Distance vs. VS2 [51] .............................................................................. 89

Figure 59. Landing Distance over a 50-ft Obstacle vs. Landing Ground Run Distance [51] ..................... 90

Figure 60. Takeoff Ground Run Distance vs. Takeoff Parameter [51] ....................................................... 90

Figure 61. Takeoff Field Length vs. Takeoff Ground Run [51] ................................................................. 91

Figure 62. 𝑪𝑳𝟑/𝟐/𝑪𝑫 Vs Zero-lift Drag Coefficient [51] ......................................................................... 92

Figure 63. Power-to-Weight Ratio vs. Wing Loading [51] ........................................................................ 93

Figure 64. Example of Matching Chart for Propeller-Driven Aircraft [51] ............................................... 93

Figure 65. Useful Load Fraction vs. Power Loading for Propeller Driven Aircraft [51] ........................... 94




Figure 66. Lift Characteristics Comparison for S1223 and FX 63-137 Airfoils [4] ................................... 98

Figure 67. Drag Characteristics Comparison for S1223 and FX 63-137 Airfoils [4] ................................. 98

Figure 68. Aircraft Model Examined Under Computational Flow Simulations [4] ................................... 98

Figure 69. Propeller Design and Outputs [84] ............................................................................................ 99

Figure 70. Tradeoff Between Efficiency and Power [84] ........................................................................... 99

Figure 71. Aerodynamics IDA .................................................................................................................. 107

Figure 72. Structures and Weights IDA .................................................................................................... 107

Figure 73. Propulsion and Energy IDA .................................................................................................... 108

Figure 74. Performance IDA ..................................................................................................................... 108

Figure 75. Stability and Control IDA ........................................................................................................ 109




List of Tables Table 1. Space Launchers Comparison [17] ............................................................................................... 17

Table 2. Activity Log, June 2016 ................................................................................................................ 20

Table 3. Activity Log, July 2016 ................................................................................................................ 21

Table 4. Activity Log, August 2016 ........................................................................................................... 21

Table 5. Gantt Chart .................................................................................................................................... 23

Table 6. Group Research Topics [19] ......................................................................................................... 24

Table 7. Earth and Mars Comparison [20] .................................................................................................. 24

Table 8. Template for Mars Aircraft Concept ............................................................................................. 28

Table 9. Space Launchers Comparison [17] ............................................................................................... 41

Table 10. Historical Aeroshell Specifications [53] ..................................................................................... 42

Table 11. Wing Span and Wing Area Limitations ...................................................................................... 42

Table 12. Aircraft Parameters That Yield a Feasible Configuration .......................................................... 52

Table 13. System Mass Breakdown for Electric Propulsion [36] ............................................................... 54

Table 14. Combustion System Summary .................................................................................................... 55

Table 15. Propellant Candidates with Isp>200 s [36] .................................................................................. 56

Table 16. NASA's Budget per Fiscal Year [81] .......................................................................................... 63

Table 17. Wing and Fuselage Geometric Parameters [17] ........................................................................ 64

Table 18. Empennage Geometric Parameters [17] ..................................................................................... 64

Table 19. Batteries Characteristics [84] ...................................................................................................... 71

Table 20. Drivetrain Design Parameters [84] ............................................................................................. 72

Table 21. Detailed Mass Description [17] .................................................................................................. 73

Table 22. MARVEL and Super Cub Control Surface Comparison ............................................................ 76

Table 23. Static Stability Criteria ................................................................................................................ 77

Table 24. Ideal Flight Velocities [89] ......................................................................................................... 78

Table 25. Mission Summary ....................................................................................................................... 81

Table 26. Summary of Missions to Mars .................................................................................................... 96

Table 27. Light Aircraft Database ............................................................................................................... 97




Nomenclature Symbol DESCRIPTION

AR Aspect Ratio

b Wing Span

c Chord Length

CD Drag Coefficient

CD,0 Zero-Lift Drag Coefficient

CD,i Induced Drag Coefficient

CL Lift Coefficient

CL,c Lift Coefficient at Maximum CL3/2/CD

CL,m Lift Coefficient at Maximum L/D)max

𝐶!!"# Maximum Lift Coefficient

CL3/2/CD Climb Parameter

c.g. Center of Gravity

g Gravity

ℎ Climb Rate

h Height

Ip Power Index

Isp Specific Impulse

K Constant for Power Index Analysis

K1 Constant for Climb Performance Analysis

ll,g Landing Ground Roll

lT,g Takeoff Ground Roll

L/D Lift to Drag Ratio

L/D)max Maximum Lift to Drag Ratio

mb Battery Mass

mT Total Mass

PA Power Available

PR Power Required

R Range

S Wing Planform Area

T/W Thrust Loading

TA Thrust Available




TR Thrust Required

𝑈 Useful Load Fraction

V Volume

Vcruise Cruise Speed

Vmax Maximum Speed

Vstall Stall Speed

W/P Power Loading

W/S Wing Loading

We Empty Weight

Wf Fuel Weight

Wg Gross Weight

Wp Payload Weight

Acronyms DESCRIPTION

ABET Accreditation Board for Engineering and Technology

ARES Aerial Regional-Scale Environmental Survey of Mars

ASDS Autonomous Spaceport Drone Ship

CAD Computer Aided Design

CD Conceptual Design

CE Chief Engineer

CEF Cost Estimation Factor

CEO Chief Executive Officer

DBF Design It, Build It, Fly It

DMMH/FH Direct Maintenance Man-hours per Flight Hour

DSI Developmental Sciences Inc

EMU Extravehicular Mobility Unit

EOS End of Segment

ERV Earth Return Vehicle

FRC Flight Research Center

GEO Geostationary Equatorial Orbit

HALE High Altitude-Long Endurance

HCO Heliocentric Orbit

IC Internal Combustion




IDA Individual Discipline Analysis

ISS International Space Station

JPL Jet Propulsion Laboratory

LEO Low-Earth Orbit

MARV Martian Autonomous Rotary-Wing Vehicle

MoM Measure of Merit

MTOW Maximum Takeoff Weight

NASA National Aeronautics and Space Administration

PS Parametric Sizing

RFP Request for Proposal

RPM Revolutions per Minute

RTG Radioisotope Thermoelectric Generator

SEI Space Exploration Initiative

Si2 Solar Impulse 2

SIAD Supersonic Inflatable Aerodynamic Decelerator

SLS Space Launch System

TAC Tail-aft Configuration

TLI Trans-lunar Injection

TMI Trans-Mars Injection

TO Take-off

TRL Technology Readiness Level

VTOL Vertical Takeoff-Landing

ZEL Zero-length Launch

ZELMAL Zero-length Launch / Mat landing

ZLTO Zero-length Takeoff Systems

Subscript DESCRIPTION

ampr Aeronautical Manufacturers Planning Report

prot Prototype

Greek Symbols DESCRIPTION 𝜀 Oswald’s Efficiency Factor

𝜌 Atmospheric Density




I. Introduction

A. Project Overview This project was established in order to fulfill the requirements for the senior aerospace capstone

design class at the University of Texas at Arlington. In general, undergraduate education consists of analytical processes, where the students learn the necessary skills to be proficient in the industry or in academia. However, most of the work is constricted to the classroom environment. The purpose of the senior design course is to integrate all that knowledge and apply it through a two-semester multi-disciplinary design project.

The first semester (spring of 2016) consisted on the conceptual design (CD) of High Altitude Long Endurance (HALE) solar-powered airplanes; with special emphasis on reverse engineer the Solar Impulse 2 (Si2) aircraft. At the conclusion of the project, the Si2 airplane was fully described from the perspective of multiple design disciplines (propulsion, structures, aerodynamics, and others). Additionally, a Loftin-based CD methodology was adapted for solar powered-HALE airplanes and applied.

Figure 1. Si2 3D CAD Model [1]

From such project, the main lessons learned where in the areas of aerodynamic efficient design, unconventional power plant/propulsion options, weight optimization, and system integration. Furthermore, the students were introduced to the design methodology, work load, and quality expectations.

For this project (summer 2016), the students engage in industry-type dynamics where they respond to a Request for Proposal (RFP) provided by the industry, a national design competition, or faculty members. Some universities have adopted Design It, Build It, Fly It (DBF) competitions as alternatives for the capstone project. Even though they might be engaging, as students obtain hands-on experience when building and designing aircraft, their actual educational value is low. In fact, such competitions are often used by individuals as pastimes, as they do not propose a real challenge for a group of senior aerospace students. Therefore, DBF competitions were not considered for this project.

The group has shifted into the conceptual design of an Aerial Platform for Mars. The main purpose is to demonstrate that flying in Mars is possible within the constraints imposed by the technology readiness level (TRL) of space exploration technologies. More details about this project can be found on the next section of this report.

B. Request for Proposal Historically, Mars exploration has been conducted through orbital (satellites) and terrestrial (rovers)

platforms. Thanks to those, it has been possible to fully characterize the Martian atmosphere and map its surface. As a result, an interactive map of the Martian surface known as Mars Trek [2] is available to the public, as well as a detailed climate and atmospheric database [3].

Preliminary studies had been conducted to determine the chemical and mineralogical composition of its soil. Furthermore, these projects have served as technology demonstrators for future space systems. However, aerial platforms have not been used yet in this endeavor. The following chart, created by Yasir Rauf [4], illustrates this issue, as it compares the missions to Mars conducted since 1997.




Figure 2. Mars Platform Comparison Chart [4]

At first glance, the main advantage that a Mars Aircraft offers is its mobility. Such aircraft could cover large regions in a short time. Moreover, it can be used to access locations where the surface features impede using conventional surface platforms (e.g. Rovers). Additionally, it offers a better resolution than orbiters, allowing taking detailed atmospheric/terrain measurements. Other mission scenarios could be devised, such as an aircraft used to deploy scientific payload at different locations [5]. In conclusion, a Mars aircraft offers high versatility to fulfill a wide variety of scientific missions with evident advantages over conventionally used platforms.

However, there is no support infrastructure available on Mars (i.e. no takeoff – landing areas, re-fueling stations, etc.) and Martian atmosphere proposes a challenge due to its low density and lack of oxygen for conventional air-breathing propulsion systems. In fact, Martian atmosphere resembles the Earth’s at 100,000 km of altitude, which is out of reach for general aviation aircraft. Moreover, how to send the aircraft to Mars in the first place? Current space capsules impose a limit on the weight and size of the payload that can be delivered to Martian surface.

The idea of using aerial platforms for Mars exploration is not new. In fact, one of the most complete works ever performed dates back to 1978 [6]. In such, they considered using up to 12 small –unmanned – remotely controlled aircraft. Two versions of the aircraft were considered: a lander and a cruiser. The cruiser was to be deployed from an aeroshell during atmospheric entry, fly while collecting data, and crash once it ran out of energy. The lander version of such aircraft had a similar design; however it incorporated the option of takeoff and land using vertical rockets. Data were to be sent to an orbiter and then re-directed back to Earth for its analysis.

Lately, the most remarkable project was the Aerial Regional-Scale Environmental Survey of Mars (ARES) [7]. It followed similar mission architecture as the cruiser version of the aircraft mentioned above (e.g. deploy, cruise, and crash). Although a mock-up model was built and tested, it has not been included in any mission yet, as NASA decided to pursuit other options. Section III of this report contains a detail description of these aircraft and others that were used as inspiration for this current proposal.

Unlike previous works, this project will explore the feasibility of having an aircraft that works as a technology demonstrator for future manned missions to Mars. As such, the aircraft has to prove autonomous capability (deployment and flight), with manned controls as an alternative. In addition, this aircraft has to be re-usable, capable of landing, taking off, and surviving the Martian environment. It has

1997 1999 2001 2003 2005 2007 2009 2011 2013 2016

Orbital Platforms

Global Coverage

Low Resolution

Aerial Platforms

Regional Coverage

Medium Resolution

Landed Platforms

Local Coverage

High Resolution

Atmosphere

Surface

Pathfinder

Mars Global

Surveyer

Mars Recon. Orbiter

Altitude

Space

ExoMars

Unexplored Regime

Aeronomy Scout

Mars Science Orbiter

Trace Gas Orbiter

Mars Exploration

Rover Pheonix Curiosity

Mars Odyssey

ESA Mars Express




been assumed as a baseline for this project that there is not significant infrastructure available in Mars. This puts the project closer to today’s TRL but also limits the solution space. The resulting aircraft could be described as a bush plane: an aircraft used to provide transport alternatives to remote (no infrastructure or ground transportation) areas.

Also within the scope of this project is the assessment of Mars atmosphere and topography, as well as the evaluation of available space vehicles payload capabilities, size limitations, and space travel options.

C. Motivations As stated in NASA Authorization Act of 2010, the long-term goal for space exploration is to expand

permanent human presence beyond low-Earth orbit (LEO). Within this goal, one of the major milestones would be to send the first manned mission to Mars and, later, the establishment of space camps [8]. Barack Obama, during the speech given at Kennedy Space Center on April 15, 2010, announced plans to send crewed missions to Mars by mid-2030s [9].

Dr. Robert Zubrin is one of the main advocates for the cause of having human missions to Mars. Back in 1990, Dr. Zubrin proposed his mission architecture for sustained human presence on Mars, which he called “Mars Direct”. In July 10 of 2014 during NASA Ames Research Center Director’s Colloquium, Dr. Zubrin presented the three main reasons why humanity should reach the red planet [10].

1) Science: is/was there life on Mars? The search for extraterrestrial life is one of the main drivers for space exploration. Preliminary observations made using rovers indicate that currently there is not life. However, there is some evidence that water was once present on the planet. This presents multiple scenarios as life, as we know it, depends on liquid water to flourish.

In the first scenario, if it is discovered that Mars in fact had liquid water and, additionally, there is evidence of past forms of life, it could be concluded that life is a direct development from liquid water, other elements, and sufficient time (as it is a chemical process) and therefore life is a general phenomenon on the universe. Alternatively, a second scenario is that water is found on Mars, but no signs of past or present life are found. This gives more credit to the theory that life is not a natural development at all, but it rather depends on chance, and therefore it is possible that life is unique to Earth. The final case is if past/present life on Mars is discovered, but there are no signs of water on the planet. A biochemical examination of such forms of life could determine if those forms of life are Earth like or not. The last scenario would change the conception of life as we know it, and would open the door for infinite opportunities for life in the universe.

2) Challenge: as Dr. Zubrin mentions on his speech, everybody remembers that in 1492 Christopher Columbus arrived to America, although other important historic events also occurred during that year. This marked the beginning of a new era in history, as it was regarded as the discovery of a new world. More recently, the Apollo 11 mission marked a high point for space exploration, as the United States were the first and only country able to land humans on the Moon. Although the Soviet Union were the first to send a human to space, the landing on the Moon is considered much more important.

In a similar way, landing on Mars would be a milestone for the next generations. This will greatly influence the society, besides the obvious scientific and political implications. Mars proposes a challenge that will increase the interest of people for space exploration. This provides motivation for future generations of engineers and scientist. Such influx of new people would quickly return the investment contributing in the development of many areas of science.

3) Future: humanity’s future could lie on space exploration. Whether it is due to human activity or an external factor (e.g. meteor impact), there are many scenarios where Earth will no longer be suitable for




holding life. Hence, humanity will have to translate, and Mars is the closest planet. Therefore, it is necessary to take the first step to open more options for the future.

D. Business Case This aircraft described in the RFP is conceived as an aid for human planetary exploration, and should

be appealing to promising commercial (such as SpaceX) or scientific (e.g. NASA) endeavors. SpaceX, founded in 2002, was conceived to make manned planetary exploration a reality. For such purpose, the first task was to facilitate access to space by having frequent and reliable launches. Following that plan, in 2006 NASA awarded $278 million to SpaceX to demonstrate the ability to send cargo to the International Space Station (ISS) [11]. Following the success of Falcon 1, NASA contracted SpaceX for 12 more robotic supply flights, for a cost of around $1.6 billion [12].

Simplicity, reliability, and cost effectiveness are part of the company’s philosophy. For such reason, one of their goals is to produce reusable rockets. In April 8 of 2016, they successfully landed Falcon 9 rocket vertically on an autonomous spaceport drone ship (ASDS) [13]. Such event was streamed online as a public relations strategy. Previously, four other rockets tried the vertical landing but failed to stay vertical. Their next rocket, the Falcon Heavy, is expected to have its first demo flight in December of this year. Crewed missions, though, are still far in the horizon for SpaceX. The company will have to go through a very extensive process of testing before they are cleared to conduct such missions.

Figure 3. SpaceX’s Falcon 9 Rocket Vertical Landing on Drone Ship [14]

This year, SpaceX’s CEO Elon Musk will announce his Mars colonization plan. On interviews, Musk has stated that the first crewed mission to Mars will be in 9 years, way before NASA’s plans [15]. However, Musk’s plan is to send a one-way mission, with no apparent possibility for the crew to return to Earth. On the other hand, NASA’s wants to land, execute a mission (a sample return, maybe), and come back.

NASA is currently working on the development of the Space Launch System (SLS). The goal of SLS is to provide access to deep space objectives. SLS Block 1, which is capable of delivering up to 70 tons of payload to LEO, 1 is scheduled for its first launch on 2017 [16]. Further developments will have greater capabilities. The following table, courtesy of Juan Lopez [17], compares previous launch systems with SLS expected performance. Then, figure 4 shows how the SLS stands up with respect to previous rockets in terms of payload volume.




Table 1. Space Launchers Comparison [17]

Figure 4. Space Launch System Payload Volume Comparison [18]

Once completed, Block 2B will be the biggest space rocket in the world, and it will be probably used for the mission to Mars. However, one of NASA’s “problems” is on bureaucracy. Since it is a federal agency, it has to rely on the government to obtain funds, and it must report back to the government on everything they do. Ever since the conclusion of the Apollo program, as it will be explained in section III of this report, bureaucracy has slowed down NASA’s progress.




II. Team Organization

A. Team Structure Two groups of senior aerospace students were created and they engaged in industry-type dynamics.

This report presents the work of Aurora Concepts’ group. Aurora Concepts works as a “fictional” corporation, with a corporate logo and mission statement. As such, it has been organized in a multi-disciplinary structure, with a Chief Engineer (CE) and discipline leaders. The team is guided by the class instructor, which acts as CEO and consultant. The positions have been assigned as follows:

• Chief Engineer: Luis Salas Nunez.

• Synthesis Lead: Ian Maynard and Ryan Manns (green).

• Propulsion Lead: Ismael Sanabria (red).

• Aerodynamics and CAD Lead: Yasir Rauf (blue).

• Structures/Weights Lead: Juan Lopez (orange).

• Performance Lead: Nic Dwyer (purple).

• Stability and Controls Lead: Justin Kenna (yellow).

Figure 5. Team Structure Chart

To respond the RFP, the team engages in a conceptual design (CD) phase, where the possible solution space is analyzed. The team analyzes mission, design, and technology trades, and then arrives to a preliminary aircraft configuration. Such configuration is then passed for further analysis by each one of the design disciplines as mentioned above. In detail, the responsibilities of the team are:

Synthesis: The synthesis team and the Chief Engineer define the conceptual design methodology and the way it is going to be implemented by each discipline. To start off the process, they will implement a parametric sizing (PS) process where they create a solution space for each discipline to work. Within this PS process, they define the airplane category, performance metrics, and measure of merit (MoM). Additionally, the PS phase includes mission, design, and technology trade studies. Once the individual disciplines perform an individual analysis, the synthesis team will integrate each discipline into a converged solution (i.e. single spacecraft layout), which will be evaluated against the MoM previously defined. If necessary, the design will be re-evaluated (iteration) until the final product meets the requirements satisfactorily.




CAD/Geometry: As the name indicates, this team is in charge of translating the work of each discipline into a visible model. They will provide any measurements that other areas need for their analysis. In doing so, they have to determine the appropriate computer aided design (CAD) tools to be used. Finally, one of the project deliverables will be a 3-D printed model of the final configuration.

Propulsion & Energy: This discipline is subdivided in two areas. First, for the scope of this project, the group will explore the energy options for flying in Mars. This constitutes one of the greatest challenges of this project, as there is no infrastructure available and so the resources have to either be readily available or have to be transported on the space capsule. In second place, this team will explore propulsion alternatives that theoretically would work on Mars while providing enough thrust to fulfill the mission.

Aerodynamics: Given the aircraft configuration, this team will analyze the aerodynamic properties of the aircraft components in detail. It will also work on validation methods for the results obtained. As there has never been an airplane flying on Mars, nor there exits flight data, it is necessary to have a “sanity” check. Other responsibilities of this team are to define the characteristics of the components required to meet the mission requirements. Some of those characteristics are: selection of airfoil, use of high lift devices, range of wing loading, drag profile, pitching moment determination, and others. The process will use analytical and computational tools for this purpose.

Structures & Weights: This discipline determines the aircraft weight (component by component) by analyzing structural loads and performing a preliminary material selection. In specifics, this team determines the location of the aircraft center of gravity (c.g.), which is the great importance for the assessment of the aircraft. Another deliverable will be an aircraft’s structural envelope diagram (V-n diagram), which will be used to assess if the current configuration is capable of surviving the different stages of the mission (launch, space travel, atmospheric entry, deployment, etc.). It is noted that for the scope of this work, an analysis on the stresses/strains requires high level of detail and computational capability, and therefore it is more appropriate for further stages in the design process.

Performance: This group evaluates the airplane operational capabilities under different flight conditions. Based on the weight, power available, and aerodynamic characteristics, this discipline determines important factors such as the rate of climb, glide ratio, landing/takeoff field distances, speed for maximum range and endurance, and maximum and stall speeds. This group’s work will be highly important when performing the configuration evaluation, as its results will be used to assess if this aircraft can, indeed, be a feasible alternative for planetary exploration.

Stability & Control: Stability can be broken down into two main categories: dynamic and static. The dynamic stability analysis is frequently used to assess the airplane’s handling qualities, and therefore it requires extensive flight test data that is unavailable at this point. On the other hand, static stability can be determined by considering the airplane’s configuration and weights. It can be subdivided into three categories: longitudinal, lateral, and directional. Within this context, this discipline determines the tail volume and control surface area required to provide stability.

This project is to take place during the summer of 2016, which is roughly 11 weeks. Due to the size of the team, the time frame, and the limited technical information available, this project will be maintained within the scope of the conceptual design phase, which will be explained in section IV of this report.

B. Activity Log The following table is used to record the team’s activity. In detail, it has been used to record official

team meetings and presentations. The table displays information about the date, activity performed,




description of what was done, and deliverable (if any). This table is useful to keep track of the work performed and represents the amount of work done over the course of the project. It is important to recall that these are only official meetings, as the group members constantly met to work together, but such meetings are not included in this table.

Table 2. Activity Log, June 2016

Date Activity Description Deliverable

6-JunIntroduction by

professor Bernd Chudoba

Project Launch -

8-Jun BrainstormingThe team got together and defined research topics as well as

preliminary mission scenarios-

10-Jun Team Meeting The team defined the baseline characteristics for the Mars Aerial Platform.

-

12-Jun Team Meeting The team finished the RFP presentation and practiced. -

13-Jun RFP PresentationTeam presented the senior design class, teaching assistants, and professor their request for proposal. Both senior design groups

got together to finalize the mission.RFP Presentation

15-Jun Class Meeting The mission was oficially defined between both teams.Mission and

baseline requirements

20-Jun Report Writing

The first report addressed the request for proposal. It demonstrated the research performed by each one of the

members and the creation of a preliminary database-knowledge base. Each indicated their solution strategy and future work.

Report 1 Due

23-Jun PresentationTeams presented their concept addressing the requirements

outlined in the proposal. Group members discussed and chose a winning configuration to start off with the design process.

Proposal Presentation

25-Jun Team Meeting

After some initial calculations, it was noticed that the chosen configuration was not feasible. The main issues were: propulsion

system, landing/takeoff ground roll, and folding techniques. In addition, the mission still needs more refinement.

Configuration Layout

27-Jun Team Meeting

Team obligations are defined. Each discipline started building their IDAs. The analysis of the configuration starts. The

synthesis team worked on adjusting the parametric sizing process and creating a work schedule for the remainder of the project.

IDA + Work Schedule

29-Jun Meeting with TA

The team meet with Loveneesh Rada (TA) to review the multi-disciplinary methodology (MDA-IDA) and the work schedule.

Based on his comments, the team reviewed the analysis methodology as it needed to be refined. The synthesis team

started to work in a new parametric sizing diagram.

MDA-IDA Presentation




Table 3. Activity Log, July 2016

Table 4. Activity Log, August 2016

Date Activity Description Deliverable

6-Jul Report WritingReport 2 is due. The initial layout will be presented. The design

disciplines will present a preliminary analysis, as well as a discussion on design trade studies and measure of merit.

Report 2 Due

9-Jul Team MeetingThe work performed by the individual disciplines was presented

to the rest of the group so everybody was on the same page. The plan for the next week was defined.

-

13-Jul MDA PresentationPresentation to Dr. Chudoba with our MDA and IDA. The

professor's input was used as a preparation for the upcoming midterm presentation.

MDA-IDA Presentation

14-Jul Team MeetingBased on the comments from Dr. Chudoba during the previous

presentation, the team prepared a format to standarize our MDA/IDA methodologies.

IDA Format

16-Jul Team Meeting

The team kept working on their IDA. The synthesis group defined a format/outline for the midterm presentation so other design disciplines can structure their slides to follow a general

plan. At this point, the disciplines finalized their IDA.

IDA and Midterm Presentation Layout

18-Jul Report WritingThe first iteration will be completed. A preliminary analysis of the configuration will be presented, including some numbers about the

aircraft performance per each discipline.

Midterm Report Due

21-JulMidterm

Presentation Preliminary Slides

The team will meet to plan the midterm presentation. Each discipline should show the content to be presented. This meeting will ensure everybody is on track and prepared for the midterm.

Slides for Midterm Presentation

25-JulMidterm

Presentation

This presentation will summarize the work performed so far. It will cover from the DB/KB, the mission selection, and the

evolution of the design. I will also explore on design trades and preliminary aircraft characteristics.

Midterm Presentation

30-Jul Team MeetingDiscussion of future work, final design decisions, and preparation

for final presentation.-

1-Aug Report Writing

Based on the numbers obtained and presented during the midterm report, the desing will be iterated as needed. A more detailed

business case analysis will be included, as there is now a better idea of how the final configuration will look like.

Report 4 Due

4-7August

Final Presentation Preparation

Final design decisions. Tasks were delegated to group members in order to prepare for the final presentation.

Slides for Final Presentation

8-Aug Final PresentationIt presents the work performed during the entire project. It will

present the final aircraft configuration and a detailed design analysis.

Final Presentation

10-Aug Team MeetingThe team shared their thoughts on the presentation, and shared

the necessary information so that everybody can prepare the final reports.

-

15-Aug Report WritingThe final report will illustrate the final configuration as well as a

complete analysis from each discipline. Final Report Due




C. Gantt Chart Gantt charts are used as a visual aid to track the status of a project. They include information about

the tasks performed and the time when such tasks were performed. Furthermore, it allows to easily visualize when a task overlaps with other. For this project, the Gantt chart is used by the Chief Engineer to assign tasks to the design disciplines, as well as keep track of the work completed. Group members also use it to adjust their schedules, as it gives them clear direction on what they are expected to complete and the time frame for such goal.

For this project, the Gantt chart can be found below. It has task descriptions on the left, with dates on top. Color bars are used to display the state of the task in the following way: green indicates completed, blue in progress, and red future work. Next to the color bars there are the names of the design disciplines/individuals assigned for such task. The final version of the Gantt chart has only green bars, as all tasks have been completed.




Table 5. Gantt Chart

Project Launch RFP Presentation

Report 1 Due Report 2 Due Midterm Report Midterm Presentation

Report 4 Due Final Presentation

Final Report Due

06/06-06/12 06/13-06/19 06/20-06/26 06/27-07/03 07/04-07/10 07/11-07/17 07/18-07/24 07/25-07/31 08/01-08/07 08/08-08/14 08/15-08/21ID TaskName 1 2 3 4 5 6 7 8 9 10 111 Literature Review2 Mars Description Ian Maynard, Ismael Sanabria, Ryan Manns

3 Space Travel and Technology Available Juan Lopez, Nic Dwyer

4 Previous Mars Aircraft Concepts Justin Kenna, Yasir Rauf, Luis Salas

5 Past Missions and Motivation Luis Salas

6 Project Proposal

7 Mission Definition All

8 Configuration Proposals All

9 Tradeoff Studies Yasir Rauf, Ismael Sanabria, Ryan Manns, Luis Salas

10 Initial Configuration Layout All

11 Conceptual Desing Methodology

12 Initial Size and Weight Estimations Yasir Rauf, Ian Maynard, Nic Dwyer, Luis Salas

13 Deployment, Landing, and Takeoff Yasir Rauf, Ryann Manns, Luis Salas

14 Multidisciplinary and Individual Analysis All

15 Aerodynamics/CAD

16 Initial CAD Model Juan Lopez and Yasir Rauf

17 Planform and Airfoil Selection Aerodynamics Team

18 Lift, Drag and Moment Estimations Aerodynamics Team

19 Wing characteristics Aerodynamics Team

20 Fluid Simulations Aerodynamics Team

21 Final CAD and 3D Printed Model Nic Dwyer and Juan Lopez

22 Propulsion and Energy

23 Propulsion System Selection Ismael Sanabria, Yasir Rauf, Ryan Manns24 VTOL Detailed Analysis Ryan Manns

25 Propeller Design Propulsion and Energy Team

26 Selection of Power Source Propulsion and Energy Team

27 Engine Description Propulsion and Energy Team

28 Structures and Weights

29 Gross Structural Configuration Structures and Weights Team

30 Detailed Aircraft Weight Analysis Structures and Weights Team

31 Deployment and Folding Mechanism Structures and Weights Team

32 Stress, Strain, and Load Analysis Juan Lopez and Nic Dwyer

33 Material Selection Structures and Weights Team

34 Center of Gravity Determination Structures and Weights Team

36 Stability and Control

37 Tail Location and Sizing Stability and Control Team

38 Fuselage Sizing (Wing-Body Integration) Stability and Control Team

39 Static Margin Estimation Stability and Control Team

40 Control Surface Sizing and Location Stability and Control Team

42 Performance

43 Stall, Cruise, and Maximum Speed Performance Team

44 Landing and Takeoff Maneuvers Performance Team

45 Aeroshell Fitting Performance Team

46 Range and Endurance Performance Team

47 Operational Envelope Performance Team

48 Synthesis

49 Convergence and Design Tradeoffs Synthesis Team

50 Parametric Sizing Synthesis Team

51 Adaptation of Loftin Synthesis Team

52 IDA - MDA Format Synthesis Team

53 Risk Assesment Synthesis Team

54 Mission Analysis Synthesis Team and Nic Dwyer




III. Literature Review An extensive literature review was conducted during the first stages of this project. Such review

intended to achieve a complete characterization of Mars, space exploration initiatives and technology, and previous Mars aircraft concepts. This work was used to identify a mission scenario and having a rough estimate of a concept of operations for this project. Ian Maynard [19] created the following table to summarize the research tasks:

Table 6. Group Research Topics [19]

A. Mars Description Mars is the fourth planet of the Solar System. It has an elliptical orbit with a period of 687 days,

aphelion and perihelion of 249.2 million Km and 206.7 million of km respectively. With respect to the Earth, it is at 78.34 million Km apart during their closest point. The following table compares Mars with respect to the Earth to understand the differences.

Table 7. Earth and Mars Comparison [20]

As for its weather, there are four main considerations to be taken to account [21]:

1) Wind speed: winds from 4.5 to 31.3 m/s. 2) Temperature variation: at the equator, from 195 K to 270 K on the same day.

Name Research TopicIan Maynard Weather

Ismael SanabriaAtmosphere

Orbit characteristics

Juan LopezAvailable launchers Payload capability

Payload fairing dimensionsJustin Kenna Previous aircraft concepts

Luis SalasMission motivations Mars mission history

Nic DwyerRFP constraints and

requirements Mission concepts

Ryan Manns AtmosphereYasir Rauf Previous aircraft concepts

Earth MarsDay Duration (hr:min) 23:56 24:40

Orbital Period (Earth Days) 365.25 687DiameteratEquator(Km) 12756 6794

Gravity (m/s^2) 9.81 3.71Average Temperature (K) 287.15 210.15Atmospheric Pressure (Pa) 101325 700

Atmospheric CompositionN2: 78.1% O2:20.9% Ar: 0.93%

CO2: 95.32% N2:2.7% Ar:1.6%




3) Pressure variation: caused by the Carbon Dioxide Cycle, still really small compared to Earth’s. 4) Dust storms: widely known for being able to completely cover the planet and take months to

clear. They can be up to 20 km high, and its particles are micrometers in diameter, so they might enter inside any systems and affect its performance.

B. The Mars Project

1. History of Missions to Mars

The fascination with the red planet comes from early times in humanity. As Hogan mentions in his book, “Mars Wars: The Rise and Fall of the Space Exploration Initiative”, Mars was observed by civilizations such as the Egyptians, the Babylonians, the Greeks, and the Romans [22]. Later, astronomers and scientists such as Copernicus, Kepler, Brahe, Huygens, Cassini, and others made important observations not only about Mars but the Solar System in general. In 1877, the U.S. Naval Observatory headed by Asaph Hall discovered Mars’ two moons and named them Phobos and Deimos. The same year, italian Vrigino Schiaparelly observed that Mars surface seemed to have canal systems. Later, American Percival Lowell concluded, among controversy and withough fully convincing the scientific community, that those canalas were the proof of the existence of intelligent beings on Mars. This greatly incresed the public interest for exploring Mars.

Before the end of the 19th century, H.G. Wells publishes “The War of the Worlds”, the first important science fiction book about Mars. This was later followed by the appearance of more books and movies that related to the red planet. By the time the space race was on, Mars was already inmersed on American popular culture.

The first Mars missions were launched in 1960 by the Soviet Union. Both, the Marsnkin 1 and 2 were conceived as Flyby missions, however the third stage of the rocket launcher failed and the missions never reached space [23]. Later, missions Sputnik 22, 23, and 24 (1962) were also unsucessful. Sputnik 23 in fact did reached Mars, but communications were lost and no data was recovered. The first successful mission was Mariner 4 by NASA. It returned the first images ever known of Mars, which proved that Mars was actually lifeless. Later, Mariner 6 and 7 (1969) also reached Mars and sent more photos.

In the 70s, Viking 1 and 2 were the first U.S. mission to successfully land a spacecraft on the surface of Mars. The mission consisted of two systems, an orbiter and a lander. These missions collected data that allowed to fully characterizing Martian atmosphere. Furthermore, they performed some biological experiments aimed to look for signs of life [24].

In the 80s there was not much activity related to Mars. Both the Soviet Union and NASA pursued other space objectives. NASA was particularly busy working on the Space Shuttle [25], whereas the Soviets launched multiple missions to Venus [26]. The 90s, on the other hand, marked a revival of the interest for Mars. Other nations, such as Japan, launched their own missions to the red planet. The most remarkable of all those missions was the Mars Pathfinder (1996). It was conceived as a technology demonstrator. It consisted of a solar-powered station and a mobile rover equipped with an X-ray spectrometer, three cameras, and other equipment. More interestingly, the mission used inflatable decelerators to cushion the landing [27].

Finally, during the last 16 years, the missions to Mars have consisted mainly on sending surface rovers and high-tech orbiters. Thanks to those, it has been possible to create a complete map of Mars and its surface features. In addition, such rovers are in the search for possible signs of life and prepare future human missions. One of the most important is the Mars Science Laboratory, or as commonly known, the Curiosity rover (2011). A complete list of all Mars missions has been included on the appendix of this report.




2. The Future of Mars Exploration

Following the Apollo missions, where six missions successfully landed in the moon, Mars was regarded as the next step for the American Space Program. Since then, crewed missions to Mars have been considered in three different occasions [22]. Most remarkably, the Space Exploration Initiative (SEI), launched in 1989 by President George W. Bush, focused on the construction of the space station, sending humans back to the Moon, and, as a long term goal, sending humans to Mars for the first time. Following President Bush announcement, multiple organizations were tasked to determine the requirements for the Mars mission in terms of technology and budget. However some of them have more science fiction than actual technology.

As an example, let’s look back to Wernher von Braun plan. In 1952, von Braun publishes “The Mars Project”, a book that outlines his vision for human travel to Mars. According to him, it was required to have ten spacecraft capable of carrying 400 tons of payload and a crew of 70 to Mars. A glider would descend from the spacecraft and land on the planet’s poles. Then, the crew would trek to the equator to build landing sites for future airplanes. The crew would set up an inflatable base camp, and conduct a 400 day survey of the planet. It is important to recall that von Braun published his book before the Mariner mission was accomplished, and so there was still the hope that Mars would be habited by intelligent beings.

In the 90’s, a study from the National Space Council concluded that the estimated cost of the SEI program would be 500 billion dollars. Both industry and academia proposed more than 2000 mission architectures that responded SEI’s requirements. One of the most complete scenarios was proposed by Dr. Zubrin in his “Mars Direct” plan [28]. Such plan received lots of media attention, although its technological demands were out of NASA’s capabilities [29].

For the Mars Direct plan, a heavy-lift rocket would be launched in 1996 carrying a 40 metric tons propellant factory (for in-situ propellant production) and an unmanned (reusable) Earth Return Vehicle (ERV). This approach was innovative, as no other project relied so heavily on using Martian resources for its sustainability. The ERV would use a SP-100 nuclear reactor as the power source.

Three years after the initial launch, NASA would send two more rockets: one similar to the propellant factory/ERV and other consisting of a habitat lander, capable of housing a crew of four. While on space, the two systems would be joined by a 1400 m long cable. Then, the two systems would spin in order to generate some gravity for its crew. The mission contemplated a six month minimum-energy trajectory for each trip, with the crew staying around 500 days on Mars. The return trip accounted for 100 kg of Mars samples.

In the long term, the two rocket mission would be send with every Mars opportunity (every two years). In this way, multiple bases would be established at different locations in the planet, so that a 100-person Mars settlement could be ready before 2020. The following is an artistic concept of how the basic settlement would look like.




Figure 6. "Mars Direct" Base Camp Concept [29]

By 1994, SEI studies were finalized, and NASA shifted its focus to space and Earth science [30]. Recently, the excitement for Mars has arisen again, especially as SpaceX has plans for crewed missions within the next decade. NASA, on the other hand, has a more indirect approach for sending humans to Mars. Their plan is to use the SLS and the Orion capsule to first send an unscrewed mission to a, still to be defined, deep space objective. Later, a crewed mission would be sent to an asteroid, something that has never been attempted. Finally, a mission to Mars would be send around 2030, using the experience gained from the previous missions [18].

C. Previous Mars Aerial Platforms Concepts Although no aircraft has ever flown on Mars, multiple design projects have been conducted for this

purpose. The first complete mission analysis and aircraft design dates back to 1978. Such project was conducted by Developmental Sciences Inc (DSI). The idea for a Mars airplane was conceived in a meeting between David Scott (former director of NASA Dryden Flight Research Center) and Bruce Murray (director of the Jet Propulsion Laboratory) in January of 1977 [31].

One of the breakthroughs that enabled the conception of such aircraft was the invention of the hydrazine engine by James Akkerman. Such engine was developed for conditions where conventional air-breathing engines are unusable. In specific, the engine was used to power NASA’s Mini-Sniffer, an aircraft which successfully proved to be flyable for cruise altitudes over 70,000 ft [32]. DSI’s final design used a combination of Akkerman’s engine with electric propellers. They were expecting their design to be included in a hypothetical mission to Mars during the 1980s. However, as the previous section explained, there was no mission at all.




Justin Kenna [33] and Yasir Rauf [4] have created a database of all previous Mars aircraft concepts. The following are works that have contributed something to the current project. In order to classify and organize the information, the following format has been created:

Table 8. Template for Mars Aircraft Concept

The purpose of this table is to have a fast an easy way to reference a specific project. For such reason,

the template’s upper section includes data about the author, organization, publication date, and citation. Then, there is a brief description of the report/project, including specific details about the aircraft (geometry, propulsion system, and others), the mission/design/technology trades, and related areas. For the “applicability” section, the template will summarize how such project contributed to the current design. Finally, in order to classify the information, the final section presents similar works that are either written by the same author or use the same mission/aircraft concept. The complete tables can be found on the appendix of this report.

The aircraft concepts of interest have been classified into three categories: 1) fixed wing, 2) inflatable wings, 3) rotorcraft (vertical takeoff).

1. Fixed Wing

The term “fixed wing” is really broad, as it allows for an unimaginable amount of possible configurations. Here, it is used here to differentiate the aircraft concepts that conceive using a rigid wing while flying.

a) Developmental Science Inc Cruiser/Lander Aircraft [31]

The first concept that is worth reviewing is the Developmental Science Inc proposal that was introduced before.

The baseline configuration, shown below, has similar features as a regular glider. It was going to use a large propeller to sustain the aircraft while flying. DSI designed a variation of this model that had takeoff and landing capabilities. The configuration retained its propeller located in front of the nose, but it also included six hydrazine rockets that gave it enough power to perform a vertical lift off and to slow down the vertical landing. The takeoff and landing maneuvers are also shown below.

Related Works

Design Discipline 1 Design Discipline 2 Design Discipline 3Use of bullet points that illustrate how this project was used on the

current project

Publication Date:

Summary: a brief description of the report and the resulting aircraft. [Figure: Vehicle Illustration]

Applicability

Author(s): Organization: Mars Aircraft Concept Overview

Reference:




Figure 7. DSI Cruiser Configuration Layout [31]

Figure 8. DSI Landing Maneuver [31]

Figure 9. DSI Takeoff Maneuver [31]

DSI’s aircraft was conceived as a versatile platform that could perform a variety of missions such as delivery of science packages, sample collection, or aerial survey. It was expected to have cruise speeds between 60-100 m/s, with a maximum range of 10,000 km, a payload capability of 40-100 kg, and a flight time between 17-31 hours.

Probably what is more interesting about this design is its folding mechanism. Using up to six folds, they were expecting to fit a 21 m wing span, 20 m2 of wing area, into a 3.8 m diameter aeroshell. Within their mission, they also perform an analysis of the loads expected during the vertical landing/takeoff maneuvers. Additionally, they talk in detail about the power requirements for each one of the configurations, as well as the power available from various power plant options.

b) Anthony Colozza, Long-Endurance Mars Aircraft [34]

Colozza’s work is to study the feasibility of a Mars long-endurance (about a year) aircraft that carries 100 kg of scientific payload. He considers two energy production systems: radioisotope/heat engines and solar cells. The design is driven by the goal of achieving the maximum endurance while keeping the minimum wingspan possible.




The study concludes that it is indeed possible to have such aircraft, although the required wing spans and areas are not feasible for space transportation.

For a 25% efficiency solar cell, the required wing span would be 47.5 m, for a wing area of 118.75 m2 and a wing loading of 3.69 kg/m2.The total power required would be over 3 kW for a 438.2 kg aircraft. On the other hand, if a radioisotope (Cm 244) is used, the resulting aircraft would have a wing span of 37.97 m, a wing area of 103 m2, and a wing loading of 2.99 kg/m2. The total aircraft mass would be roughly over 300 kg, and the power required is 7.36 kW.

Figure 10. Artistic Representation of Long-Endurance Aircraft [34]

When comparing the two systems, the author concludes that the radioisotope one is more versatile as it is not limited by the required solar irradiance to sustain flight. Additionally, the sizes and weights are smaller, which increases its feasibility.

Finally, the study concludes with recommendations for future Mars aircraft design projects [34]:

1) A definitive flight plan should be designated in order to model the environment in which the aircraft is expected to fly. 2) To increase the collected solar energy, movable or variable geometry wings should be considered. 3) The radiator system for both, radioisotope and solar powered aircraft, must be fully described in order to obtain accurate figures of merit for the system. 4) A deployment scheme for the aircraft should be devised. This scheme depends mostly on the

mission type.

The biggest contribution obtained from this work is on the design methodology. In order to perform the analysis and arrive to wing, power, and weight estimations, the author adapted HAPP design methodology for the Mars aircraft concept. Such methodology is used for the design of Long-endurance aircraft; however most of its assumptions are still applicable to this project. Furthermore, the appendix contains detailed weight estimation equations. Finally, the publication contains detailed propulsion system diagrams, which will aid on the design of the propulsion system for this project.

The author has other publications that are also relevant to this project. The “APEX 3D Propeller Test Preliminary Design” [35] contains design specifications and performance data for the design of propellers on Mars-like atmospheric conditions. Finally, the “Comparison of Mars Aircraft Propulsion Systems” [36] includes detailed performance data for multiple propulsion/power plant options for future Mars aircraft concepts. Such paper will be used not only as a reference, but also as a point of comparison for the estimations and assumptions made by the propulsion/energy discipline.

c) Aerial Regional-scale Environmental Survey of Mars (ARES)

The last work that is worth mentioning at this point is the Aerial Regional-scale Environmental Survey of Mars (ARES). ARES is a scientific aerial platform that, unlike the previous works, has proven

SENIOR DESIGN:

MAE 4350 Mini Project

Ref.: MAE 4350-001-2016 Date: 25. Aug. 2016 Page: 31 of 127 Pages Status: In Progress Status: In Progress


to be feasible and flyable. The concept has been tested on Earth and it was considered to be included in one of the latest NASA’s missions to Mars.

The paper “Evolution of a Mars Airplane Concept for the ARES Mars Scout Mission” [37] contains a detailed description of the mission/design/technology trade studies conducted while designing the ARES. It is possible to observe how the design evolved from the baseline configuration to the end result. The mission design is also very specific, and can be summarized as follows [37]:

1) It must fit into an aeroshell with a maximum internal diameter of 2.48. 2) Must survive G, radiation, and thermal environments associated with launch, space travel,

and atmospheric entry inside an aeroshell. 3) It must be a stable platform for science experimentation. 4) It has to be autonomous, with capability to follow a specified ground track, flying at altitudes

1-2 km above the ground and with a range around 500 km. 5) The final requirement was for the data to be simultaneously transmitted to a communication

satellite without interruption.

Figure 11. ARES Final Configuration (left) and Prototype [7]

The design philosophy behind ARES is to meet the science requirements; hence, the risk, complexity, and cost of the system should be minimal. The aircraft mission architecture consisted on: 1) autonomous deployment from aeroshell, 2) pullout maneuver to transition from atmospheric entry to cruise flight, 3) fly while collecting data, 4) crash land once its energy was over.

The ARES configuration was selected after considering options such as non-rigid wings, flying wings, and even canard-wings. The final configuration, as shown above, is an inverted v-tail which can be folded forward in order to fit within the aeroshell. As for the propulsion system, the glider option was discarded after some calculations proved that it would not be possible to meet the range requirements. On the other hand, rocket propulsion was considered as the lowest risk option. Moreover, there is a long tradition of using rockets for space exploration, thus adding more certainty to the design.

In summary, the final design has a wing span of 6.25 m, an aspect ratio of 5.58, a CL between 0.52-0.71, a L/D of 14-14.4, a T/W~0.1, a range of 500 km, and an endurance of 60 min.

There are many things that this project can learn from the ARES design, as well as the entire data library available that discusses detailed aspects about the aircraft design. Pretty much all of the design

SENIOR DESIGN:




disciplines have obtained something from ARES, and for such reason it will be used as a reference for the remainder of this project.

Other relevant works are “Mars Airplane Airfoil Design With Application to ARES [38]”, “High Altitude Drop Testing in Mars Relevant Conditions for the ARES Mars Scout Mission” [39], and “Simulating the ARES Aircraft in the Mars Environment” [40].

d) The Mars Airplane Revisited

This project took place during the spring of 1988 at the University of Illinois at Urbana-Champaign. It was a senior design project for aerospace engineering students. Every semester, the faculty presented students two project alternatives: an aircraft and a spacecraft. This project proved to be the perfect match of both areas, as it explored the Mars aircraft (“Marsplane”) itself and the system required to deliver it [36].

For the aircraft design project, the goal was to design a system capable of: 1) carrying two persons (including the life support systems) for a payload weight of 1200 N (in Mars), 2) the airplane must be able to takeoff/land in a prepared surface no larger than 1000 m in length, 3) it must have an endurance of at least eight hours, 4) a rescue scenario must be conceived, and 5) the airplane and aeroshell had to be compatible. Finally, it was assumed that some basic infrastructure was available on Mars in order to provide the necessary fuel/assembly for the aircraft operation.

Due to the large size of the student body, the group was divided in eight teams, each one performing an independent aircraft design. In each team, every member was appointed as a discipline lead, with one person serving as the group coordinator. The disciplines were: aerodynamics, performance, power and propulsion, stability and control, structures and materials, surface operations, and weights and balance. In addition, Loftin’s design methodology was used for the parametric sizing process. In many ways, the Marsplane design challenge resembles the structure and purpose of the current project, and therefore it is highly relevant.

In total, eight aircraft were designed, each one with different sizes, weights, and performance. The paper presents tables that show the minimum, average, and maximum parameters for each design discipline. Therefore, it can be used as a comparison to evaluate where the current configuration stands. The following figure shows two of the resulting designs as an example.

Figure 12. Examples of Marsplane Resulting Configurations [41]

It was possible to identify three common design drivers as follows: 1) low atmospheric density, which results in low wing loadings, 2) need for lightweight propulsion/energy systems, 3) the takeoff/landing performance which required some type of aid or VTOL capability otherwise it was not possible.

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Of the eight resulting designs, two employed canards, two had joined wings, and four used the conventional tail-aft configuration. The gross weights ranged from 4600-7500 N, with wing spans of 37.5-72.0 m, cruise speeds between 60-109 m/s, aspect ratios of 8-25, and wing loadings from 19.1 to 100 N/m2. As for the propulsion system, they had power loadings from 71-222 N/kW, and propeller diameters between 3.0-8.7 m.

In general, all design disciplines can learn something from the Marsplane project. However, the most important contribution comes from the fact that they used Loftin for their parametric sizing process. However, in doing so, it was necessary to adapt it, as it does not have data for low wing loading aircraft, nor it accounts for other propulsion systems besides propellers and jets. Furthermore, it is debatable whether Loftin’s sizing process is accurate for Mars atmosphere, as the only way it accounts for the low atmospheric density is through the use of the density ratio.

e) The Zephyr: Manned Martian Aircraft

The Zephyr is an aircraft designed at the University of Toronto in conjunction with Canadian Forces in 1999. This study assumed the existence of multiple human settlements in Mars, and therefore proposed using an aircraft as a quick and versatile system to provide transport solutions between those base camps. In addition, the Zephyr could be used to conduct scientific studies and participate in search-and-rescue operations. Hence, it had to be designed to support a crew of two persons with a range of 200 nm.

The sizing process the aircraft should weight about 2000 Mars pounds, with a wing area of 2000 ft2. Such structure was both large and light, and therefore an important focus was placed on the structural design. This was the main reason why the biplane configuration was chosen at the expense of the drag. The design was based on powered sailplanes and hydrogen-oxygen fuel cell specifications.

Figure 13. Zephyr Layout and Propeller [42]

A propeller was designed and optimized to operate within Mars atmosphere. The resulting design had ten 5.9 ft diameter blades. At 900 RPM, each propeller provided 54 lb at cruise speed, and therefore 3 were required to meet the thrust requirements. The structural analysis concluded that the design could withstand loads of +/- 2g for different maneuvers. The static margin was determined to be around 6%, with acceptable static longitudinal, lateral, and dynamic stability. A performance analysis determined the stall/maximum speeds, climb rates, turn, service ceiling, airport performance, and V-n diagram.

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Finally, the takeoff runway was determined to be around 20,000 ft, with 6,100 ft more required to meet the 50-ft ground clearance requirement. For this project, such length is not an issue, as it assumes the existence of base camps in Mars and clear runway areas.

The main contributions for this project are on the areas of structures, performance, stability and control, and propulsion. This project stands over the rest (with exception of the ARES project) as it contains a detailed stability analysis, including dynamic stability derivatives determination. In addition, it estimates the required control surface deflection to meet the stability requirements.

On the other hand, the paper might lack enough data to sustain its findings. Moreover, the excessive takeoff/landing field required makes the Zephyr design not appropriate for this study.

2. Inflatable Wings

The University of Kentucky has conducted extensive research on the use of inflatable wing aircraft. Professor Jamey D. Jacob from Oklahoma State University has led multiple research projects on the design of HALE aircraft, which closely resembles the atmospheric conditions that would be encountered in Mars. These concepts offer the advantage that is easier to fit into the space capsules or aeroshells, and they have a considerable less weight than other options (see references [43] and [44]). However, there might not be enough data or theoretical-analytical tools for this project.

In detail, the University of Kentucky and Oklahoma State University worked together on the flight testing and simulation for a Mars aircraft that employed inflatable wings. A prototype was built and tested under various flight conditions. The results of such test were compared against the analysis performed using different computational software (such as “UNCLE”). The following figures show the actual system with its wings folded (deflated) and a top view of its design.

Figure 14. Prototype of Aircraft with Inflatable Wings [44]

One of the most important areas of interest was to compare the effect of having “bumby” or “smooth” wing surfaces. Depending on the system used to inflate the wings, the resulting airfoil will have discontinuities (“bumps”), and at first glance it would be desirable to cover it to make it smooth. However, it was concluded that at low Reynolds numbers (<100,000), the “bumpy” profile had a better aerodynamic performance, as its surface reduced flow separation. However, for moderate (~250,000) and high (~500,000) Reynolds number, the smooth profile performed better than the bumpy one [45].

In addition, reference [46] contains an extensive review of all possible deployable wing mechanisms: folding, inflating, nesting, extending, morphing, and hybrid. Moreover, it analyzes the aerodynamic and structural design requirements for each one of the previous options. If inflatable wings were to be used for this project, these would be the main sources of reference.

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3. Rotorcraft

In response to the 2000 American Helicopter Society Student Design Competition, a team of aerospace students from The University of Maryland conducted the study “The Martian Autonomous Rotary-Wing Vehicle (MARV)” [47]. They evaluate the feasibility of using a vertical lift platforms (solar powered) to study different locations on the Martian surface. Such aircraft provide a good alternative for planetary exploration due to its maneuverability and precision, thus combining the benefits of a rover and an air vehicle.

The mission requirements for this project were: 1) maximum takeoff mass of 50 kg, 2) autonomous deployment, 3) 25 km range, 4) 30 min of controlled flight, 5) 1 min of hovering flight, and 6) optional restart capability. The critical design points are no different than a Mars aircraft. That is, low atmospheric density, low Reynolds number operation, lack of oxygen for conventional propulsion systems, etc. Folding and deployment from an aeroshell was also part of the design requirements.

The study concludes that such vertical platform is feasible within the constraints imposed. The downside of using rotorcraft devices is on the high power requirements and the limited payload capabilities. On the MARV, ~20% of the gross takeoff weight is used for the power supply system. On RHOVER (explained below), power train is ~40%, with a payload of roughly 15%. Reference [47] contains a detailed analysis of each design aspect behind the MARV. As an illustration, the following is the final configuration.

Figure 15. MARV Design [48]

On the other hand, students at Cornell University also studied the feasibility of using a rotary-powered aircraft for Mars exploration (see reference [48]). Their study is mostly focused on determining the power and torque required for hovering and analyzing its feasibility under the current technology. Such analysis would be tremendously useful in case this project decided to employ a rotorcraft.

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IV. Methodology

A. Multi-Disciplinary Analysis (MDA) In general, the design process is an iterative procedure, which can be broken down into three phases:

first, the conceptual design phase, were the major parameters of the artifact (the airplane) are outlined. It is in this stage where the designers define what the purpose behind the construction of this airplane is. One key factor to remember is that “form follows function”, which implies that the airplane will be designed in order to fulfill a given challenge. Henceforth, some mission requirements are established, along with a preliminary sketch of what the airplane will look like. Other key parameters, such as the desired operational specifications (speed, altitude, range, and payload) are set based on what the engineers consider that is necessary to fulfill the mission. In most projects, the conceptual design phase is carried out by a small group of lead engineers, which represent the key design parameters to consider during the following design stages.

Overall, airplane conceptual design can be sub-divided into some key aspects such as: aerodynamics, propulsion, performance, controls, stability, materials, CAD & geometry, systems, business case (costs), and others. Some of these areas might not appear in the design process at all, it depends on the aircraft category and the mission profile. However, it is important to have a balance between these design parameters, and that is where the synthesis group and the chief engineer role come to play. As a guide, the design process needs to follow Concept of Operations (ConOps) document, which describes how the airplane is to be operated [49]. Towards the end, design trade studies will indicate if the initial requirements can be achieved under the current configuration. If not, it is necessary to conduct a study on the possible trade-offs that yield a plausible solution.

The following design stages are the preliminary and detailed design stages. In the first, the conceptual design results are validated through extensive testing. More detailed studies are conducted to determine the load distributions, stress on the structure, engine-airframe coupling, among others. Finally, some manufacturing considerations are taking into account, which serve as a start off point for the detailed design stage [49]. For the last one, the configuration is fixed, and no major changes should be applied to the structure. In this point, the engineers will design all the small components and mechanisms that are necessary to operate the aircraft. Sometimes it will be convenient to design mock-ups of the actual system. At this point, the production of the actual airplane starts.

Again, this project will be kept within the conceptual design phase. For conventional aircraft design, multiple authors have formulated generic processes for the parametric sizing process. In general terms, parametric sizing consists of arriving to a general aircraft configuration based on mission/technology/design trades. The PS phase, then, analyzes the solution space to define if the proposed mission is feasible within the current industry capability. Then, the PS phase will output an estimate of the size/scale of the vehicle, as well as a configuration that best suits the mission [50].

The information obtained from the PS phase will be used as a start point for the design process, with outputs that will be used by the individual disciplines for detailed analysis. Such analysis is used to determine the single design point that converge all disciplines to meet the mission and market requirements. In this process, the design might need to be iterated, where some parameters on the aircraft configuration are changed and the resulting effect on the performance is evaluated. For such reason, this process must be parametric.

There is not a single formula that yields a more accurate answer. Loftin, for example, implements a matching procedure that depends on performance requirements and historical correlations for propeller and/or jet-driven aircraft [51]. As a result, Loftin arrives to a converged weight, size, and power

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combination that yield a feasible design point that meets the mission requirements (more details of Loftin’s methodology will be presented in section V.B of this report). Gary Coleman’s dissertation [41] contains a full review of other PS processes and their implementation on real design case studies.

The synthesis team and the chief engineer have worked on adapting a parametric sizing program for the design of Mars aircraft. The following diagram shows an overview of the MDA.

Figure 16. Parametric Sizing Diagram Overview

The parametric sizing process starts with the initial concept research, which allows defining a clear and feasible mission for this design project. Then, such requirements will be used by the synthesis team during the parametric sizing stage, where they are to determine the aircraft parameters in the areas of weight, size, propulsion, performance, and aerodynamics. Then, these results are used to determine the aircraft configuration (airplane geometry box), which will be further analyzed by the structures and weight discipline.

In this process, other disciplines appear as sanity checks for the results at each stage. Iteration loops are represented by red arrows for clarity. The final step occurs with performance, which is going to analyze the resulting design to check if it meets the mission concept. As it was mentioned before, the previous figure is just an overview, which is used to facilitate the understanding of the process. A more detailed diagram is shown below:

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Figure 17. Detailed MDA Diagram

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During the initial concept research, it was found that there are two main drivers: weight and configuration. Weight is constrained by the launcher selection and the deployment system used. Those will be discussed in more detail in the next chapter of this report. As for configuration, the initial research was driven by the selection of a proper propulsion/energy system, the Mars environmental/topographic conditions, and the takeoff/landing strategy.

After this point, a refined mission was specified as the diagram shows. Such requirements were passed through the parametric sizing process, which outputs variables for the design disciplines.

It is really important that all design disciplines understand their position within the design methodology. Each one of them will adapt their methods to serve this process, as their individual discipline analysis will show.

B. Individual Responsibilities and Scope Every discipline has developed an Individual Discipline Analysis (IDA) chart. In such, the discipline

determines the internal information flow. A template of such diagram was created by the synthesis team in order to standardize it across the disciplines. Such template displays what a simple aerodynamic IDA would look like for the general goal of analyzing the wing.

Figure 18. IDA Template

The IDA flows from left to right, and it is subdivided into four columns: 1) tasks, 2) known, 3) analysis, and 4) output. For tasks, the discipline defines its major responsibilities in the design process. This gives the process clear guidance and provides clarity for the reader. The “known” column consist of the variables needed for the analysis, as well as the discipline where the variable is obtained from. A color matching system is employed to make it clearer.

The analysis is the most important portion of the IDA. It includes information on critical design points, assumptions, analysis methods, and calculations. Red arrows indicate the moment where iterations take place. This portion should clearly tell the story of how to fulfill the task in hand. Finally, the output column shows the variables that result from the analysis.

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In order to avoid the cluster on this report, IDAs have been moved to the appendix of this report. However, they are still important to understand the design process and the group methodology. In addition, each discipline has developed a personal time line that helps them to keep track of their project and guide their work.

The author of this report did not have an IDA for his work. Instead, the Chief Engineer work consists on giving direction to the design disciplines and ensuring that the project is going to the right direction and within schedule. Furthermore, the Chief plays an important role on the synthesis and evaluation stages, as he is the one who takes decisions when conflicting designs come up. For such purpose, it is necessary that the Chief Engineer has knowledge on all categories and keeps up to date with all of the information gathered by the teams.

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V. Conceptual Design Analysis The conceptual design analysis has been subdivided in three sections: mission, design, and

technology studies. The first section analyses the possible mission profile of the Mars aircraft. This is fundamental as the mission is the main design driver for the aircraft. Remember that one must always “design to mission” [49]. Then, based on the specific mission profile developed, the second section presents a parametric sizing study, as well as an analysis of the critical design challenges that the mission imposes. Finally, the technology studies section addresses, in specific, the technology readiness level and feasibility of the solution space.

A. Mission Studies

1. Transportation and Delivery to Mars

This aircraft is conceived within the limitations imposed by the fact that there is no infrastructure available on Mars (as outlined in the RFP). As a result, anything that is designed to operate in Mars must be transported and remotely controlled from Earth. Therefore, the first part of the mission analysis starts with the selection of the launch system, as this imposes a limitation on the size, weight, and volume of the aircraft. This process is coupled with the consideration for the delivery (i.e. aeroshell) and deployment (i.e. folding and packing) of the aircraft onto the Martian surface.

As a preliminary consideration, the team has decided that the system should be delivered using a single rocket and fitted within and aeroshell, as any proposal consisting of more than one rocket launch would be turn down immediately. Table 1, repeated here for convenience, compares the current and future launcher options.

Table 9. Space Launchers Comparison [17]

The latest systems, Atlas V-541, Delta IV, and Falcon Heavy (in development), offer similar payload

capabilities, with diameters ranging from 5.0-5.4 m. The Space Launch System (SLS), on the other hand, will provide payload fairings ranging from 8.4 m to 10 m, and up to 130 tons of payload to LEO (SLS Block 2B). Using the Falcon Heavy, which is expected to have its first demonstration in December of 2016, would put this project closer to the current state of technology [12]. SLS Block 2B allows for having a larger system, with more than three times the payload capability, but it will not be complete until mid-2020 [52].

In second place, it is necessary to consider the aeroshell that will be used to deliver the system to Mars’ surface. The following table summarizes the specifications for historic aeroshells:

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Table 10. Historical Aeroshell Specifications [53]

So far, the largest aeroshell was used for the Mars Science Laboratory (MSL), or as commonly

known, the Curiosity rover. It had a diameter of 4.5 m and the possibility to land almost 1000 kg of weight. Currently, there are projects developing a Supersonic Inflatable Aerodynamic Decelerator (SIAD) with a diameter of 4.7 m and fit to be used to deliver payload to Mars [50]. Now, once SLS Block 1B and 2B are ready, there will be research programs aimed to design an aeroshell that better suits those rockets. It would be reasonable to assume that such systems will have diameters over 9 m.

Figure 19. CAD Model of SIAD [54]

In DSI’s project, the aircraft was designed with a 21 m wing span and 20 m2 wing area and folded it into a 3.8 m aeroshell [27]. If a similar folding scheme is used, and assuming there is a linear correlation between wing span/wing area and aeroshell diameter, it is possible to determine hypothetical combinations for future aeroshells. The calculations on table 10 will be used to assess if the sizing results are feasible within the constraints imposed by the aeroshell size. Another possible constraint, on the other hand, would be the aeroshell mass delivery capability. The aeroshell shown in figure 18 can carry up to 3160 kg [50]. The resulting Mars aircraft will probably have low wing loadings, and so this limit seems reasonable.

Table 11. Wing Span and Wing Area Limitations

Aeroshell Diameter [m]

Wing Span [m] Wing Area [m^2]

3.8 21.0 20.04 22.1 22.24.5 24.9 28.05 27.6 34.65.5 30.4 41.96 33.2 49.96.5 35.9 58.57 38.7 67.97.5 41.4 77.98 44.2 88.68.5 47.0 100.19 49.7 112.29.5 52.5 125.010 55.3 138.5

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2. Mission Specifications

The Request for Proposal section on this report states “this project will explore the feasibility of having an aircraft that works as a technology demonstrator for future manned missions to Mars. As such, the aircraft has to prove autonomous capability (deployment and flight), with manned controls as an alternative. In addition, this aircraft has to be re-usable, capable of landing, taking off, and surviving the Martian environment.” In specific, the aircraft must be capable of carrying a human. This requirement will be used to determine the payload mass.

a) Selection of Payload Mass

The selection of the payload mass is not trivial and plays an important role on the design process, especially since any increase on the payload weight results in an even larger increase on the aircraft gross weight (and size as well). A more detailed weight sensitivity analysis, on later sections, supports this claim.

For this project, the payload mass has been determined based on the weight of an astronaut plus its suit, life support systems, and leaving some allowance for any tools/equipment that he might need during the mission. Looking at the past, the Apollo suit weighed around 180 lb (~82 kg), whereas the Space Shuttle suit, with the life support system, weighed 310 lb (~141 kg). The difference is due the Apollo suits being used for one mission and the requirements for them to be light so astronauts could easily move on the moon [55]. Although the next generations of space suits are more likely to be lighter, it is better to be conservative on this aspect, as any extra weight could be used to carry more equipment. References [56] and [57] contain a detailed analysis of the weight and energy requirements for the space suits. Hence, the payload mass has been determined as follows:

• Pilot: 75 kg • Space Suit + Life Support: 145 kg (enhanced EMU) [55] • Miscellaneous: 30 kg (for tools and other external equipment) • Total payload mass: 250 kg

This payload mass is larger than what any other previous Mars aircraft has conceived, and as a result, this aircraft will be considerable larger (see section III for previous concepts). As this is a technology demonstrator, the first missions will probably be unmanned. Still, the aircraft must prove that it can carry such a large weight before being considered for a manned mission.

b) Operating Speeds

The determination of the stall and maximum speeds are necessary for the parametric sizing process. Due to the low atmospheric density on Mars, the aircraft will need a very low wing loading. Then, a preliminary performance analysis demonstrated how the flight speeds are larger than Earth’s general aviation aircraft. Hence, the following have been determined:

• Stall Speed: 70 m/s (~155 mph) • Maximum Speed: 150 m/s (~335 mph)

c) Short Takeoff and Landing

Since it is assumed no infrastructure is available on Mars, the aircraft will have to takeoff and land from the planet’s natural ground. Previous observations (using rovers and satellites) have determined that Mars’ surface is highly variable, with its surface being either rocky or covered with “sand” dunes [58]. Therefore, it has been set as a requirement that the takeoff/landing ground roll should be as short as

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possible. If such requirement is not feasible, any type of vertical takeoff-landing (VTOL) system would have to be designed.

• Landing and Takeoff ground roll: 92 m (~100 ft) or VTOL

d) Range and Endurance

The range/endurance requirement is used to evaluate the propulsion system and the amount of fuel to be consumed during the mission. These variables are used during the sizing process, as it will be explained later. It is important to differentiate between range and endurance. The first refers to the distance to be covered during the mission, whereas the second specified the flight time. Then, depending on the type of mission to be performed, one parameter becomes more important than the other. For instance, if it was desired to have an aircraft to perform atmospheric sampling, the design would rather be driven by the endurance of the aircraft. On the other hand, if the aircraft is needed just to transport the payload from one site to another, then range becomes the driving requirement.

Mobility is one of the reasons why having an aircraft on Mars is attractive. Mars Rover Opportunity set a record when it covered 42.2 km on the Martian surface. During the Apollo 17 mission, the Lunar Rover traveled 35.7 km, which is the longest distance a manned vehicle has covered [59]. On the other hand, an aircraft flying at a flight speed of 100 m/s for an hour would cover a distance of 360 km. This is just an example of how the aircraft would greatly offer an increased operating environment for future Mars missions.

There is, however, a limitation that must be taken into account. The EMU offers life support for up to ~8 hours [60]. A hypothetical mission profile for this aircraft would consist on taking off from the Mars base (or landing site) to a site of interest. Then, the astronaut exits the aircraft to perform his tasks. Once completed, the astronaut takes the aircraft and flies back to the base. Therefore, and as a security measure, the total flight time (round trip) should be kept to less than 2 hours. This leaves enough time for the astronaut to perform his mission (2-3 hours) and still leaves some time for any unexpected event. Based on that, the following requirement was determined:

• Range: 400 km

Based on the previous mission requirements, a parametric sizing process was performed in order to obtain a solution space for the aircraft design. This process, as well as the critical design points, are addressed in the following section.

B. Design Studies The Request for Proposal (RFP) section fully describes the goal of this design project. The idea is to

provide a simple but robust aircraft to be used for manned Mars exploration. Therefore, the first questions that one must ask are, first, is it even possible to fly an aircraft on Mars? Second, are Earth’s aerodynamic principles still applicable to Mars?

As explained in the Mars description section, Mars has a thin atmosphere, just on the limit to be suitable for fly. Its average atmospheric density is 0.015 kg/m3, about 100 times less than Earth’s. However, Mars’ low gravity (1/3 of Earth’s) is an advantage. Additionally, the low temperatures (~250K) yield a sonic speed around 250 m/s. Hence, the resulting Mars cruise conditions are of low Reynolds number, low atmospheric density, but high subsonic Mach number. Finally, the low density environment requires aircraft to fly at high speeds (high stall speed) in order to produce enough lift. This, combined with the low subsonic Mach number, really limits the operational envelope. As a result, and based on the mission parameters defined above, the following critical design points have been defined:

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1. Landing and takeoff: most of the works referenced in section III had something in common: they were to be deployed from the aeroshell during atmospheric entry, fly until they were out of energy and crash once the mission was fulfilled. However, such mission is not appropriate for this project, as the manned requirement induces the necessity to land and takeoff. Previously, it was determined that either a short takeoff-landing (STOL) was necessary due to the lack of infrastructure. If such requirement could not be met, then the system requires VTOL capability. This introduces challenges in every one of the design disciplines. What are the loads that the structure must withstand during such maneuver? What propulsion system would be needed to achieve the necessary thrust for a vertical takeoff? This leads to the second design point.

2. Propulsion: the selection of the propulsion system is paramount, and will be addressed with more detail on the technology studies section of this report. A preliminary assessment concludes that due to the lack of oxygen, conventional air-breathing engines cannot be used. The low atmospheric density and low sonic Mach number critically affects the amount of thrust that can be obtained from a propeller. A rocket system would need fuel to be transported from Earth, which dramatically increases the cost of the project. Finally, a hybrid propulsion system (rocket + propeller, for example) is really complex and would need extensive research and testing.

3. Energy: either the system will use in-situ resources, or those will have to be transported to Mars in the spacecraft. Mars atmosphere is mostly composed of CO2, which could be used as propellant. In fact, Zubrin’s Gashopper has demonstrated that CO2 can be used as propellant for a vertical takeoff-landing (VTOL) vehicle [45]. However, the concept is not mature enough to be considered for this project. It is rather conceived to be used for small aircraft. Other options, such as batteries, fuel cells, and solar panels, can also be considered, although the amount of energy available from those systems might not be enough.

4. Aircraft transportation and deployment: aircraft transportation was previously addressed as part of the mission studies. Deployment, on the other hand, is a critical design point. It is necessary to specify if the aircraft is to be deployed on air, if it will be assembled on the ground, or other. From a structures point of view, it is necessary to consider the folding mechanism (if there is any), the aerodynamic loads during atmospheric entry and deployment, and system integrity in general. These considerations, however, seem to be too complex for the conceptual design stage. The final design will try to be as simple as possible, but no detailed folding or deployment mechanism will be described.

1. Initial Considerations

Initially, the team was split into four different groups. The idea was to independently propose an aircraft configuration that addressed the critical design points. From such exercise, the groups coincided that ultralight aircraft could serve as inspiration. Generally speaking, ultralights are small, simple, and efficient. The main advantage was on its inherent light weight structure, which facilitated its transportation. In specific, the Lazair Ultralight was chosen as a baseline due to its configuration.

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Figure 20. Lazair Ultralight Airplane [62]

Based on the Lazair style, the goal was to adapt it to make it flyable on Mars. The inverted v-tail was attractive as it resembled the ARES design. Furthermore, the use of a tail boom simplified the folding of the aircraft into the aeroshell. Its empty weight is less than 100 kg, while providing a wing loading of 15.4 kg/m2. As a comparison, DSI’s lander had an estimated wing loading of 14.3 kg/m2. Hence, the maximum takeoff weight (MTOW) was estimated to be around 990 N on Mars (~450 kg).

With such a low weight, this design would provide allowance for fuel to be included within the spacecraft. Therefore, the initial propulsion system was the AMPAC-ISP 22-N thruster. A quick estimation determined that about 80-N would be enough to provide enough power to maintain cruise conditions. Therefore, around 5-6 rockets would be sufficient. Finally, as for the landing/takeoff, an erroneous estimation predicted stall speeds around 30-40 m/s, and therefore it was thought that a conventional short landing/takeoff would be feasible.

With this configuration in mind, the team started their analysis. A detailed performance analysis showed the error in the stall speed calculations, as the actual stall speed for an aircraft of such characteristics would be around 80 m/s. For such speeds, and with the current available thrust, the takeoff/landing ground rolls would be too large, thus some type of VTOL system had to be added to the design. However, an early estimation determined that about 300 kg of fuel were necessary for the VTOL. This value was deemed too high considering that the maximum takeoff mass had been estimated to be 450 kg.

In order to have a better visualization of the solution space, a Loftin-based parametric sizing study was conducted, based on the mission requirements specified on the previous section of this report. The goal is to obtain estimates for the weight, power, and size combination that yield a feasible aircraft to meet the mission. Such study is the paramount importance for the design disciplines to start their analysis.

2. Parametric Sizing

The Chief Engineer, in consultation with the Chief Executive Officer, has decided to use Loftin methodology for the sizing phase. Such method is explained in detail in the appendix, section A, of this report. In summary, this method consist on determining the aircraft size (wing area S), weight (W), and power (P), starting from a set of performance requirements and some baseline (assumed) aerodynamic parameters. It makes use of historical correlations obtained from a pool of three general aircraft configurations for both propeller and jet-driven aircraft.

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It is noted, though, that Loftin does not have data for low wing loading aircraft. From the literature review, it is concluded that a Mars capable of flying in Mars would inherently have wing loadings < 10 lb/ft2. Loftin method can still be used, but its results might not be accurate at all. Moreover, the low atmospheric density combined with the low Reynolds number flow field has an effect on the power requirements. Finally, this Mars aircraft is probably going to use some sort of VTOL or rocket aid, which is not accounted for in Loftin.

The following diagram illustrates how the parametric sizing process works. It starts with five mission requirements which are the stall speed, maximum speed, takeoff/landing field lengths, range, and payload mass. Those were set to 70 m/s, 150 m/s, 300 ft, 400 km, and 250 kg respectively. The field length was chosen so that the resulting design would have an extremely short takeoff/landing, which would be acceptable under the assumption that no infrastructure is available at the planet.

Using those requirements, Loftin’s method determines a power index (Ip) based on the maximum speed (Vmax), a wing loading (W/S) parameter as a function of the stall speed (Vstall) and a guessed maximum lift coefficient (CLmax), and curves of power loading (W/P) as function of wing loading to meet the climb and field length requirements. The power index and the wing loading parameter are combined to produce an additional curve. The result is a matching chart that illustrates the combination of wing/power loading that meets the mission requirements.

After this point, the power loading is used to determine a useful load fraction (𝑈), and a fuel fraction (𝑊!/𝑊!) is obtained from the range requirement. Such fuel fraction depends on the type of propulsion and power plant system used. Then, using the payload weight (𝑊!), which is also obtained from the mission requirements, the aircraft gross weight (𝑊!) is found. Finally, the wing area (S) and power (P) are found using such weight and the wing/power loading. The following diagram illustrates the entire process, with the variable flow shown at each step.

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Figure 21. Loftin Sizing Process Diagram

This process was implemented using a MATLAB script, which has been included in the appendix of this report. As a result, the following chart was produced:

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Figure 22. Loftin-based Preliminary Matching Chart

From this chart, it is noted how the takeoff/landing field length requirements are not feasible. In other words, some sort of rocket aid or VTOL system has to be devised, which agrees with the initial assessment. The stall speed line (blue) is obtained from an analysis of conventional aircraft with horizontal takeoff, so such line also needs adjustment. In conclusion, the initial sizing did not throw satisfactory results.

Based on this, it was decided to collect a database of light aircraft in order to obtain more accurate correlations. A set of 67 aircraft were studied, including general aviation aircraft, powered gliders, and ultralights. The data collected consisted on maximum takeoff weight, wing area, engine power, empty weight, and operating speeds. Other characteristics, such as the type of engine used, the aircraft configuration (braced wings, type of landing gear, and others) were also included.

From the data obtained, the following two curves were plotted: maximum speed vs. power index, and useful load fraction vs. power loading. The power index is a variable defined by Loftin and corresponds

to !/!!/!

! . The useful load fraction, on the other hand, is defined as 1 −𝑊!/𝑊!. The result is shown in the

following figures:

0 1 2 3 4 5 6 7 80

10

20

30

40

50

60

70

Wing Loading Parameter (W/S), lb/ft2

Powe

r Loa

ding

(W/P

), lb

/hp

Matching Chart

Max. Speed RequirementTakeoff RequirementClimb Req.Stall Speed Req.Landing Wing Loading

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Figure 23. Max. Speed vs Power Index for Light Aircraft

Figure 24. Useful Load Fraction vs. Power Loading for Light Aircraft

While there is a correlation between the maximum speed and the power index, the useful load fraction data does not allow obtaining a useful equation. Therefore, only the data between maximum speed and power index were used. As for the stall speed curve, it was determined in the following way (calculations shown in international units for convenience):

𝑉!"#$$ =2𝑊/𝑆𝜌𝐶!!"#

y = 226.59x - 70.69R² = 0.77444

0

20

40

60

80

100

120

140

160

180

200

0.5 0.6 0.7 0.8 0.9 1 1.1

Max

. Spe

ed, m

ph

Power Index, Ip

Max. Speed Curve Linear (Max. Speed Curve)

y = 0.0135x + 0.2435R² = 0.27395

00.10.20.30.40.50.60.70.80.9

1

5 7.5 10 12.5 15 17.5 20 22.5 25 27.5 30

Use

ful L

oad

Frac

tion,

U

Power Loading, lb/hp

Useful Load Fraction Curve

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𝑊𝑆=12𝜌𝐶!!"#𝑉!"#$$

! = 0.5 0.015 1.8 70! = 66.15𝑁𝑚! = 17.83 𝑘𝑔/𝑚!

From these adjustments, a matching chart was build using three curves: a maximum speed curve, a stall speed curve, and a climb requirement curve. The later was found using the climb criterion defined by FAR 23 for aircraft whose stall speed is larger than 70 mph as follows:

𝑃𝑊=

7.784 𝑊/𝑆𝐶!!"# 𝜎

+1736.8 𝑊/𝑆𝐶!!/!/𝐶! !"#

𝜎

33,000𝜂

Figure 25. Loftin-adapted Final Matching Chart

For convenience, the power loading and wing loadings were converted to international units, easing the data flow between disciplines. The curves intercept at two points. The shown point corresponds to the combination that yields the larger wing loading while meeting the climb requirements. However, it is noted that any point under the red and black curves and to the left of the blue curve results in a feasible combination.

Prior finding the gross mass, it was necessary to estimate the fuel mass fraction. Again, this mission will require a vertical takeoff/landing. For such, Ryan Manns [63] developed a MATLAB script that determines the amount of fuel used by rockets during this maneuver given a liftoff altitude, rocket Isp, and nozzle mass flow. It determines that the fuel fraction used ranges from 0.20-0.30 for the landing / takeoff, it depends on the fuel’s Isp. On the other hand, if batteries are used, the mass fraction of the batteries is found using the following equation, obtained from Ref. [64]:

Feasible Design region

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𝑚!

𝑚!=

𝑅𝐸𝜂(1/𝑔)(𝐿/𝐷)

Which, for a range of 400 km, and assuming a propulsive efficiency of 0.85, a L/D of 24, and an energy density of 300 Wh/kg, yields that the resulting mass fraction is 0.0674. Then, the fuel fraction is found by adding the mass of fuel used for the VTOL maneuver and the battery mass fraction:

𝑊!

𝑊!= 𝑊!!"#$ +𝑊!!"##$%& = 0.3 + 0.0674 = 0.3674.

Then, using the gross mass prediction method (as outlined on the appendix, section A), and setting the wing loading constant at 17.87 kg/m2, the following parameters were found:

Table 12. Aircraft Parameters That Yield a Feasible Configuration

The chosen parameters have been highlighted in yellow. In summary, the resulting aircraft will have a

gross mass of ~1500 kg, a wing area of 84 m2, and a power of 150 kW. As for the landing and takeoff, the aircraft will be assisted by rockets, whereas it will use propellers for cruising flight. The disciplines will start their analysis aiming to confirm the previous characteristics.

An alternative sizing process (range-based) was developed by the synthesis team. It was modified to best suit electric aircraft. In broad terms, it determines cruise parameters such as the thrust/power required, the L/D, maximum lift coefficient, and total aircraft mass for different ranges of wing span and wing area. Such results will be used as a validation of the results previously obtained. More details about this process can be found on Ryan Manns report [63].

Power Loading (kg/kW)

Useful Load Fraction (-)

Gross Weight (kg)

Empty Weight (kg)

Wing Area (m^2)

Power (kW)

15 0.740 670.07 173.92 37.58 44.6714 0.699 752.87 226.30 42.22 53.7813 0.658 859.01 293.45 48.18 66.0812 0.617 999.99 382.63 56.08 83.3311 0.576 1196.32 506.85 67.10 108.7610 0.535 1488.60 691.75 83.49 148.869 0.494 1969.85 996.21 110.48 218.878 0.453 2910.92 1591.58 163.26 363.877 0.412 5573.71 3276.17 312.60 796.246 0.371 65386.35 41116.25 3667.17 10897.735 0.330 -6719.24 -4500.88 -376.85 -1343.854 0.225 -1756.15 -1361.02 -98.49 -439.043 0.200 -1493.82 -1195.05 -83.78 -497.942 0.175 -1299.67 -1072.23 -72.89 -649.831 0.150 -1150.18 -977.66 -64.51 -1150.180 0.125 -1031.54 -902.60 -57.85 -

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a) Weight Sensitivity Analysis

Jan Roskam [62] defines functions to determine the gross weight sensitivity as functions of the payload weight, empty weight, and other variables such as range, endurance, or propulsive efficiency. The author defines the takeoff weight as follows:

𝑙𝑜𝑔𝑊!" = 𝐴 + 𝐵𝑙𝑜𝑔(𝐶𝑊!" − 𝐷)

Where A and B are growth factors defined by the author. For a composites-made twin engine propeller aircraft, A = 0.1130 and B = 1.0403. C and D, on the other hand, are defined as:

𝐶 = {1 − 1 +𝑀!"# 1 −𝑀!! −𝑀!"#}

𝐷 = 𝑊!" +𝑊!"#$

Where Mres is the reserve fuel fraction at the end of the mission, Mff is the overall fuel fraction, and Mtfo is the trapped fuel fraction. From the mission specification and the parametric sizing process, it is defined that Mres = 0, Mff = 0.37, and Mtfo is assumed to be 0.005. On the other hand, WPL and Wcrew are the payload and crew weight. As a result, C = 0.365 and D = 550.

The sensitivity analysis determines how changing one parameters affects the aircraft gross weight. It is possible to determine that change by taking the derivative of the gross weight with respect to any variable “y” in the following way:

𝜕𝑊!"

𝜕𝑦= 𝐵 𝑊!"

! 𝜕𝐶𝜕𝑦

− 𝐵𝑊!"𝜕𝐷𝜕𝑦

/{𝐶 1 − 𝐵 𝑊!" − 𝐷}

As it was mentioned on the mission studies section, the selection of the payload weight plays an important role on the MARVEL design. Therefore, it was pertinent to determine the sensitivity of the gross weight with respect to payload weight. Replacing y by WPL:

𝜕𝑊!"

𝜕𝑦= 𝐵𝑊!" 𝐷 − 𝐶 1 − 𝐵 𝑊!"

!!

Note that !"!!!"

= 1.0 and !!!!!"

= 0. Replacing B, C, D, and Wto (from the sizing process) on the previous equation:

𝜕𝑊!"

𝜕𝑦= 1.0403 3275.8 550 − 0.365 1 − 1.0403 3275.8 !! = 5.697

This means that any pound increased on the payload weight results in an increase of 5.697 pounds on the gross weight. This shows how critical is to be careful on the selection of the payload weight, as it is fundamental on keeping the aircraft structure light and within the region of convergence.

C. Technology Studies This section will present technology studies in the areas of propulsion for Mars aircraft and the use of

VTOL systems as a solution to the limited infrastructure problem.

1. Propulsion Systems for the Mars Aircraft

Providing enough propulsion and energy is one of the critical challenges behind the design of a Mars aircraft. In this field, Anthony Colozza has conducted extensive research throughout his career. In detail, reference [33] analyses using three different propulsion systems: 1) electrical propulsion, 2) combustion engine systems, and 3) rocket systems. Three different mission durations were considered (1, 2, and 4

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hours). All systems were designed to produce 35N, which is the approximate thrust required for an aircraft with a 5 m wing span.

The systems analyzed by Colozza depend on Earth resources. Such resources must be transported to Mars, which greatly increases the cost of the mission. In addition, if the aircraft was to be re-usable, it would depend on future supply missions, which makes the operational cost unbearable. A most feasible option would be to use martian resources. As an example, Dr. Zubrin proposes using martian CO2 as propellant. This alternative will also be studied.

a) Electrical Propulsion

The author concludes that the total drivetrain mass (motor controller + motor + gearbox + propeller) is 18.1 kg. This value will be kept constant for the electric propulsion analysis. Then, both fuel cells and battery-based systems are considered.

Fuel cells can be used to generate electrical power. The system consists of a fuel cell stack, pressure regulators, filters, hydrogen and oxygen pressure tanks, and other components [36]. A conversion efficiency of 50% was assumed to calculate the amount of hydrogen/oxygen. The fuel cell performance is estimated to be 1 Kg/kW. Then, as the mission duration increases, so does the specific energy, as only the tank and reactant masses have to increase, while all other systems remain unchanged.

On the other hand, a battery propulsions system is simpler, as it does not require active controls nor uses mechanical components. In addition, hydrogen/oxygen tanks might require a significant amount of volume, whereas batteries can be allocated throughout the aircraft. The two issues with batteries are their energy densities and discharge rates. From all, the lithium sulfur chloride and lithium manganese dioxide seem to be the most appropriate options. Besides, heating becomes a concern when using batteries.

In summary, the following table summarizes the masses of fuel cell-battery propulsions systems. Recall that drive train mass is fixed at 18.1 kg, and systems have been sized to produce 35 N of thrust.

Table 13. System Mass Breakdown for Electric Propulsion [36]

* Silver Zinc ** Li-Sulfur Chloride

b) Combustion Propulsion System

As oxygen is not available in Mars, conventional combustion engines cannot be directly used for the Mars aircraft. However, such systems can still be applied if the oxidizer is either carried with the aircraft, or the combustion is done by using less energetic fuels that decompose through a catalytic reaction [36]. Therefore, the performance of a combustion engine will depend on the fuel and engine used.

Moreover, the fuel / oxidizer combination is also critical for its performance. Such combination was chosen based on: 1) ability to perform in Mars’ environment and during space transit, 2) storage within the vehicle, 3) ability to meet the performance requirements. As a result, the combination must have a low

Component 1 Hour 2 Hour 4 HourPower System, Dry

Mass (kg) 25.47 27.44 31.44Power System, Wet

Mass (kg) 29.83 36.15 48.87Specific Energy (W-

hr/kg) 239 395 584Total Mass (kg) 47.93 54.25 66.97

Battery (kg) 56* 113* 141**Wiring + Controller (kg)

Total Mass (kg) 77.1 134.1 162.1

Fuel Cell

Battery System 3

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freezing point (below -40°C) to insure that it remains liquid. This is also really important to kept the fuel volume low and reduce its complexity.

For fuel options, the author considers the following: hydrogen, ammonia, hydrazine, monomethyl hydrazine, unsymmetrical dimethyl-hydrazine, RP-1, methane, propane, and diborane. For oxidizer, the options considered are: oxygen, fluorine, nitrogen tetroxide, chlorine trifluoride, inhibited red fuming nitric acid, and oxygen difluoride. On the other hand, monopropellants could not be used for internal combustion systems. As a result, the author has selected the following combinations, which offered the best performance.

Table 14. Combustion System Summary

Other options considered were piston expanders, akkerman engines, and internal combustion engines.

Piston expanders, which are normally used on torpedo propulsion, needs to be cooled and scaled down to meet the power and weight requirements. The Akkerman engine, on the other hand, was considered by DSI on their Mars aircraft design. Its specific energy is 1.62 kW/kg, and the fuel consumption is around 2.7 kg/kW-hr. Finally, internal combustion (IC) engines are normally used for small model aircraft. However, they could be scaled up to produce enough power.

For a two hour mission duration, the following is a mass breakdown of the piston expander, and 2-4 cycle IC engines. It is evident how a 4 cycle IC engine yields the lowest mass of all options considered.

Figure 26. Mass Breakdown of Combustion Engines for an Endurance of 2 Hours [36]

Fuel Oxidizer Heat of

Combustion (kJ/kg)

Oxidizer/Fuel Ratio

Density Fuel/Oxidizer

Hydrogen (H2) Oxygen (O2) 144000 8 18.35/375.46

Monomethyl-Hydrazine(MMH)

(N2H6C)

Nitrogen Tetroxide: 25% mixed oxides

(N2O4)

28347 1.75 874/1450

Hydrazine (N2H4) NA 19429 NA 1008

Unsymmetrical dimethyl-Hydrazine

(UDMH) (N2H8C2)

Nitrogen Tetroxide: 25% mixed oxides

(N2O4)

39960 2.25 792/1450

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c) Rocket Propulsion System

Rocket propulsion has been historically used for space applications. From all options considered, rocket have the largest technology readiness level. However, their high fuel consumption imposes a limit on the endurance.

The highest specific impulse propellant should be chosen to maximize performance. The thrust generated is calculated by: 𝑇 = 𝑀×𝑢!, where 𝑢! is the exhaust exit velocity and M is the mass flow. The exhaust velocity is calculated as 𝑢! = 𝐼!"𝑔, assuming an ideal nozzle and that the flow is fully expanded at the nozzle exit.

The best propellant (fuel/oxidizer) combination is MON25, which has an Isp of ~295 s. The highest Isp occurs at a mixture ratio of 1.55. Using the equations shown above, the exhaust velocity is 2893 m/s, with a mass flow of 0.0121 kg/s (for 35 N of thrust). If a monopropellant was to be used, the best option would be hydrazine. The following table summarizes the propellant combinations with Isp > 200 s [36].

Table 15. Propellant Candidates with Isp>200 s [36]

When using rocket systems, there are two options: simple blowdown systems or regulated systems.

For the first, the thrust decreases steadily throughout the flight, as it depends on the fuel pressure. As a result, this system has a non-uniform thrust, which is not appropriate for this application. The second system is more complex and, at the same time, heavier. However, it enables having a constant thrust throughout the entire mission.

d) Conclusion: System Comparison

From the author point of view, the main characteristic of comparison is the propulsion system mass. However, it is noted that the differences in technology readiness level and operational risks made the systems not comparable at all. In summary, the following conclusions are reached:

Propellant Oxidizer / Fuel Ratio

Specific Impulse (s)

Fuel: Monomethyl Hydrazine Oxidizer: Nitrogen Tetroxide

(MON25)1.625 295

Fuel: UDMH Oxidizer: Nitrogen Tetroxide

2.7 286

Fuel: Monomethyl Hydrazine Oxidizer: Chlorine Tetroxide

3 283

Fuel: UDMH Oxidizer: Chlorine Tetroxide

2.85 278

Fuel: Monomethyl Hydrazine Oxidizer: IRFNA

2.5 274

Fuel: RP-1 Oxidizer: IRFNA

4.9 263

Monopropellant: Nitromethane - 245Monopropellant: HPB-2517 - 220

Monopropellant: n-Propyl Nitrate - 209

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1. Rocket systems have the lowest performance but the highest TRL. Due to the high fuel mass, rocket systems should not be considered for mission durations greater than 2 hours. Yet, for short flight times, a bipropellant (MON-25) rocket would be the best choice.

2. Combustion systems are lighter than rockets, but they require a significant amount of development and research, and therefore they are not feasible for short-term missions.

3. Fuel cell systems are a relatively new technology, and it has never been used as the main power system for an entire aircraft, thus they still have to be designed, tested, and developed.

4. Batteries have two complications: first, the only batteries that offer a good enough energy density have slow discharge rates, thus they might not provide enough power. Second, some sort of external cooling system would need to be implemented, as convection cooling with the surrounding airflow is not enough.

5. For propellers, it is necessary to develop a special blade that can operate at low Re and high subsonic flight Mach number. Besides, the propeller has the complication that it must be folded and then deployed.

e) A Different Approach: Using CO2 as Propellant

The idea of using CO2 as propeller comes from Dr. Zubrin’s proposal, the Mars Gashopper (see Ref. [66]). Such system can be considered as an alternative to explore Mars given its high mobility and, more importantly, because it uses Martian CO2 rather than transporting fuel from Earth.

For this system four options were considered, one unheated and three heated CO2 rockets. After the analysis, it was concluded that a heating system, which uses a hot pellet bed, is the best option. The idea is to use a pump to acquire CO2, which is then stored in liquid form in a tank at 150 psi. Then, the gas is warmed until it reaches a pressure of around 1000 psi. At this point, a valve is opened allowing the liquid gas to flow over a hot pellet bed (previously heated to 700 C). This process turns the liquid CO2 into gas, which is then expelled out through a nozzle to produce thrust.

There are main advantages provided by this hot pellet bed over the other heated/unheated concepts. First, is rate of energy transfer depends on the gas flow rate, and thus it can be easily controlled. Second, it is a robust system, optimal for repeated use. Third, and more importantly, its energy density is higher than any current battery. For instance, a lithium bed heated to 1000 K would provide the same energy as a hypothetical 750 Wh/kg battery. Higher temperatures provide even more benefits.

The specific impulse (Isp) is the most important performance metric when evaluating rocket systems. The specific impulse can be plotted as a function of chamber temperature. The analysis used One-Dimensional Equilibrium software and assuming a chamber pressure of 1000 psi and a nozzle expansion ratio of 400. Ideally, specific impulses over 100 s can be achieved with temperatures around 600 K. A 2000 K provides an Isp above 200 s. Other performance parameters (velocity and range), though, are found not to be dependent of chamber temperature.

Dr. Zubrin asserts that production of liquid CO2 can be done at an energy cost of 84 Wh/kg, which can be provided by solar cells. The problem comes with the pumping part of this system, as the most efficient (at the time) were only 2% efficient. Then, a 180 W compressor could acquire 50 g of CO2 per hour. The two alternatives would be to, first, use thermal energy from an RTG (demonstrated in 1994), and second, to acquire the CO2 by freezing (demonstrated in 1998). The first requires around 50 W of energy to produce 500 g of CO2 per day. The second requires 30 W to produce the same 500 g per day.

Dr. Zubrin’s report was elaborated in 2000. The current technology improvements make using CO2 a more viable option. This option, in the opinion of the author of this report, is the necessary technology to enable using aircrafts (manned or unmanned) in Mars, simply because of the elevated costs and risks of sending resources from Earth.

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2. Rocket-Assisted Takeoff-Landing

Rockets have been used primarily on military aircraft to produce a significant increase in the thrust for short periods of time. Given the nature of their missions, sometimes it is necessary to reduce the takeoff distances to the minimum amount as possible. There are also scenarios where it was necessary to deploy and aircraft but no airfields were available. This led to the development of zero-length takeoff systems (ZLTO, or ZEL for zero-length launch) during the cold war, which seemed more feasible than a VTOL system.

The first program, launched in 1953, was named “zero length launch / Mat landing” (ZELMAL). It used F-84G Thunderjet fighters, but was cancelled after 28 tests [67]. This program, although cancelled, proved the feasibility of using rockets for takeoffs. Later, the Air Force returned to the concept of using rocket-assisted takeoff systems. This time, the aircraft chosen for tests was a F-100. In this case, a disposable booster would be used to instantaneously produce up to 130,000 lb of thrust, which would launch the aircraft to an altitude of 120 m at a speed of 450 km/h [67].

Figure 27. Rocket-Assisted Launch Fighers, F-100 (left) and F-84 (right) [68] [69]

Even though most of the tests were successful, the concept had many difficulties if it was to be used during real combat situations. The logistics of transporting the launch platform into areas of difficult access, the security concerns with the pilot, and the necessity of having a landing field ultimately led the concept to be deemed unviable. Soviet’s ZEL program, which used a MiG-19, also went through the same issues [67]. Rockets were further considered for other scenarios. One of the most striking ones was the addition of rockets in a C-130 Hercules. The engineers expected these additions would allow the Hercules to land and takeoff in a distance less than 100 m [70].

Rolls Royce made two developments that changed the way rockets were used on aircraft. The first was the Thrust Measuring Rig (1954), which was later turned into the NASA’s Flying Bedstead. Such technology was the foundation for the Lunar Landing Research Vehicle. Its use of rockets for hover control was innovative and opened a whole world of possibilities. On the other hand, the Pegasus Engine (1959) was the enabling technology of the Harrier’s VTOL system. Such engine is able to produce up to 23,800 lb of thrust both forward and downward [71]. Nowadays, this technology is known as thrust vectoring.

The Harrier was the first and only VTOL aircraft during the 60s. Although its performance was significantly below other fighter aircraft, it was considered a viable option by both the United States and the British government. In fact, Harriers were fundamental for the British during the Falklands War [70].

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Figure 28. NASA's Flying Bedstead and Rolls-Royce's Pegasus Engine [72] [73]

The previous is only a brief review of some rocket-assisted systems for takeoffs and landings. The reality is that there have been a lot of experimental aircraft that have used any variation to the conventional takeoff and landing maneuvers. Examples like using tiltrotors (see the USAF CV-22 [74]), tail-sitters (see ref. [75] for a brief review of the history of these aircraft), and more recently the F-35 Lightning II and its multiple variants just show how much knowledge is available on the field. Looking to the future, DARPA has been working on developing alternative VTOL concepts that could, potentially, combine the advantages provided by helicopters and conventional fighter aircraft [76].

In regards to this project, it is important to emphasize on the extensive legacy of using rocket-controlled systems for spacecraft. That, combined with the knowledge achieved by the previous mentioned projects, makes using a rocket-assisted takeoff / landing for this Mars aircraft a feasible idea.

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VI. Design Disciplines This section is used to present the final design configuration discipline by discipline. The reader is

encouraged to read the individual disciplines reports for a more complete analysis of the results presented. As an aid, the individual discipline analysis diagrams (IDAs) have been placed on the appendix of this report (section D).

This chapter starts with the synthesis discipline followed by the aircraft CAD/Geometry. This is done so that the reader has a visualization of the aircraft from the starting point. This facilitates the comprehension of the details presented in subsequent sections. Performance, on the other hand, has been moved to the end of this section, as it presents a mission analysis to demonstrate that the aircraft does, indeed, meet the mission requirements. Therefore, the disciplines have been arranged in the following manner: synthesis, CAD/Geometry, aerodynamics, propulsion and energy, structures and weights, stability and control, and performance.

A. Synthesis This section is used to present a summary of the final aircraft configuration after the disciplines have

performed their analysis. A cost analysis is also presented at the end of this section.

1. Aircraft Configuration Highlights

The team has decided to call the aircraft MARVEL, which stands for Mars Aerial Research and Environmental Laboratory. From now on, the aircraft will be referred by its acronym MARVEL. An isometric view of the aircraft is shown below.

Figure 29. Isometric View of MARVEL

The aircraft has been designed with a high aspect ratio wing (17.6), with a slight dihedral at the tips for improved lateral stability. The wing has a wing area of 110 m2 and a wing span of 44 m. This guarantees proper aerodynamic performance on the low Re number flow that characterizes Mars. The fuselage was designed as a tadpole, sized to properly fit a full-suited astronaut, while leaving enough space for the payload. Such shape guarantees good aerodynamic performance while minimizing the volume. The tail boom extends for 13.27 m, which gives the tail enough moment arm to provide longitudinal stability. Finally, the tail was designed with twin vertical tails and a horizontal stabilizer. A

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single vertical tail would be so large that it would touch the ground, and therefore the twin structure was designed.

The aircraft uses a hybrid system for its propulsion. The landing and takeoff are powered by hydrazine rockets, whereas two four-bladed propellers are used during cruise flight. This guarantees that the aircraft can operate virtually everywhere on Mars without requiring any special infrastructure. The propellers have also been sized to provide enough thrust to operate up to 2700 m, with velocities ranging from 65 to 125 m/s. Such propellers are powered by Lithium-Sulfur batteries. With this configuration, for a cruise speed of 85 m/s, the range and endurance are 779 km and 2.55 hour respectively.

In total, the aircraft has a mass of 1468 kg, with the engines producing about 60 kW of energy. This results in a wing loading of 13.34 kg/m2, and a power loading of 24.47 kg/kW. With this dimensions in mind, the aircraft requires a space rocket capable of carrying an entry, descent, landing (EDL) system at least a 9m on diameter, and capable of landing a mass of 1500 kg.

2. Cost Analysis

Cost is probably one of the main constraints for space exploration technologies. Historically, space agencies have failed to predict the cost of their projects. This complicates finding a valid business case and obtaining enough support to finance the project. Cost analysis is divided in two areas: first, the cost of producing the aircraft, and second, the cost of transporting and delivering the aircraft to Mars.

a) Cost of the Aircraft

For the MARVEL aircraft, the cost prediction method was based on Roskam’s method for estimating prototype program cost [61]. As the author states, prototypes are used as technology demonstrators, such as NASA’s X-planes. This method is valid given the character and the intention behind the MARVEL aircraft. It is assumed, though, that no more of 4 prototypes will be produced. Similarly, this cost is only for the aircraft development. The cost of the total project must include the space transportation, aeroshell development, and operational cost.

Roskam defines the prototype cost (Cprot) to be a function of the Wampr (Aeronautical Manufacturers Planning Report), the number of prototypes to be built (Nprot), and the Cost Escalation Factor (CEF). This relationship is illustrated as follows:

𝐶!"#$ = 1115.4 10! 𝑊!"#$!.!" 𝑁!"#$

!.!!(𝐶𝐸𝐹!!!" !"#$)/(𝐶𝐸𝐹!"#$)

Where Wampr, in pounds, is found as:

𝑊!"#$ = 𝑖𝑛𝑣𝑙𝑜𝑔 0.1936 + 0.8645 𝑙𝑜𝑔𝑊!"

Roskam applies his weight calculation method to the Grumman X-29 prototype program. The cost of such program, although no official number has been disclosed, is estimated to be around US $100,000,000. Using the equations shown above, for a takeoff weight of 17,800 lbs, two built prototypes, and for the year 1982, the cost of the program is estimated to be US $109,724,907, which agrees well with the initial estimation.

Then, using the takeoff mass of 1467.7 kg (~3228.94 lb) as determined by the structures department, assuming only one prototype will be build, and with CEF2016 = 6.27 and CEF1973 = 1.14 [62], the following calculations are performed:

𝑊!"#$ = 𝑖𝑛𝑣𝑙𝑜𝑔 0.1936 + 0.8645 log 3228.94 = 1687.26

𝐶!"#$ = 1115.4 10! 1687.26 !.!" 1 !.!!(6.27)/(1.14) = 𝑼𝑺 $𝟖𝟐.𝟔𝟔𝟐 𝑴𝒊𝒍𝒍𝒊𝒐𝒏

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If four prototypes were to be built, the cost would escalate to US $336 million. This cost must be added to the space launch, aeroshell development, and operational cost. In addition, this method is more appropriate for aircraft expected to be flown on Earth. A Mars aircraft will have a higher level of sophistication and would have to undergo extensive testing before being approved.

b) Space Launch, Transportation, and Delivery

The selection of the space launch system depends on the technology readiness level of such system and the MARVEL dimensions. As it was shown earlier, the rockets to be considered are SpaceX’s Falcon Heavy and NASA’s Space Launch System. Due to the MARVEL dimensions, it would need to be either transported on a single rocket launch using SLS Block 2B (10 m payload fairing), or using multiple launches on the Falcon Heavy (5 m payload fairing).

Estimating the cost of the rocket launch and operation is not an exact science. Since these rockets are still in development, the only option is to refer to NASA’s and/or SpaceX quotes. John Strickland has developed an estimated for SLS costs based on the Space Shuttle launch and operating costs [66]. In the past, a 2011 study determined that the Space Shuttle costed about $1 billion per launch ($1.5 billion if development costs are included). On the other hand, the annual operating costs were between $3-5 billion.

Strickland quotes that the development costs of SLS and the Orion capsule is of $3 billion for ten years. Then, since SLS will not have any re-usable parts (as the Shuttle did), the operating costs would be reduced to about $2 billion. However, such costs do not account for the development of any payload (rover, space exploration vehicle, or others). For this project, such payload would be the MARVEL aircraft, along with its EDL system.

The cost per launch depends on how often the rocket will be used. There are many factors that affect the frequency of the launch, such as how often can NASA prepare a new payload. Strickland produces the following figure, illustrating the cost per pound for various launch frequencies for both, SLS block 1 and block 2.

Figure 30. Cost per Pound for Different Rockets and Launch Frequencies [79]

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For this project, it is going to be assumed that one SLS rocket is launched every two years (every Mars opportunity window) with re-supply missions. Therefore, using Strickland’s estimates, it is found that the cost per pound would be $31,100 to LEO. Then, assuming that about one-fourth of the payload can be taken to Mars (as stated by Dr. Casey Handmer) [80], it is found that the cost of taking the MARVEL to Mars would be:

MARVEL weight: 1468 𝑘𝑔 × 2.2 !"!"

~3230 𝑙𝑏

Cost to LEO: 3230 𝑙𝑏× $31.1 !!"

~ 𝑈𝑆$100.44 𝑀𝑖𝑙𝑙𝑖𝑜𝑛

Cost to Mars: 𝑈𝑆$100.44 𝑀𝑖𝑙𝑙𝑖𝑜𝑛 ×4 ~ 𝑼𝑺$𝟒𝟎𝟏.𝟕𝟔 𝑴𝒊𝒍𝒍𝒊𝒐𝒏

The final step on the analysis is to estimate the cost for the EDL system. Unfortunately, there is not a model that allows determining how much of the total cost of a project is dedicated for the EDL. In fact, such systems depend a lot on the needs of the payload and the landing location. For such reason, that analysis is not included in this report.

As a result, the total cost of the project would be the summation of the space launch cost and the aircraft manufacturing: 𝑼𝑺 $𝟒𝟎𝟏.𝟕𝟔 𝑴𝒊𝒍𝒍𝒊𝒐𝒏 + 𝑼𝑺 $𝟖𝟐.𝟔𝟔𝟐 𝑴𝒊𝒍𝒍𝒊𝒐𝒏 = 𝑼𝑺 $𝟒𝟖𝟒.𝟒𝟐 𝑴𝒊𝒍𝒍𝒊𝒐𝒏. Such cost must be conceived within NASA’s current budget. The following table shows the budget for fiscal years 2015 to 2020.

Table 16. NASA's Budget per Fiscal Year [81]

For the current fiscal year (2016), the total budget is US $18.53 billion. From that figure, planetary

science consists of 26% of the total science budget (around US $1.37 billion). Note that the development of the MARVEL aircraft will probably take more than a single year. Also, SLS Block 2B is expected to be ready by mid-2020. Hence, it is not possible to determine how much of the total budget would be used for this project. But still, it is seem to be well within NASA’s budget.

B. CAD/Geometry The CAD and geometry was taken care of by two group members. Yasir Rauf, from one side, is the

person in charge of putting together the final CAD model to illustrate the aircraft configuration. Juan Lopez, on the other hand, kept track of a geometry file that was shared with all group members. Such geometry file contains the necessary information not only to create a CAD model, but also to perform and stability and control analysis. It is also used to determine the weights of each aircraft component. As a result, the following tables show the complete geometric parameters:

FY2015 FY2016 FY2017 FY2018 FY2019 FY2020Science $5,244.70 $5,288.60 $5,367.90 $5,448.40 $5,530.20 $5,613.10Aeronautic $571.40 $571.40 $580.00 $588.70 $597.50 $606.40Space Technology $724.80 $724.80 $735.70 $746.70 $757.90 $769.30Exploration $4,505.90 $4,505.90 $4,482.20 $4,298.70 $4,264.70 $4,205.40Space Operations $4,003.70 $4,003.70 $4,191.20 $4,504.90 $4,670.80 $4,864.30Education $88.90 $88.90 $90.20 $91.60 $93.00 $94.40Safety, Security, & Mission Services $2,843.10 $2,843.10 $2,885.70 $2,929.10 $2,973.00 $3,017.20Construction & Envrmti Compl Restoration $465.30 $465.30 $436.10 $442.60 $449.30 $456.00Inspector General $37.40 $37.40 $38.00 $38.50 $39.10 $39.70Total $18,529.10 $18,529.10 $18,807.00 $19,889.20 $19,375.50 $19,666.10

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Table 17. Wing and Fuselage Geometric Parameters [17]

Table 18. Empennage Geometric Parameters [17]

As the lead of the Structures and Weights discipline, Juan Lopez has developed volumetric estimates

for the cockpit, based on the required volume to accommodate one fully suited astronaut and the standard human measurements (95th percent). For details on such calculations, refer to his report [17]. Consequently, he has also taken the job of designing the CAD model for the fuselage, including the cockpit. This model was then integrated with the tail (designed by Justin Kenna, Stability and Control lead) and the wing (designed by Yasir Rauf, Aerodynamics lead). Such models are shown below:

Figure 31. MARVEL Fuselage

Variable Description Value UnitsSw Wingplanformarea 110 m^2ARw Wingaspectratio 17.6 -bw Wingspan 44 mTRw Wingtaperratio 0.8 -CRw Wingrootchord 2.5 mCTw Wingtipchord 2 mCMACw Wingmeanaerodynamicchord 2.5 mALLEw Leadingedgesweep - degAL25w Quarterchordsweep - degTCw Averagechordsweep - -TWISTw Wingtwistangle 8 deg

Variable Description Value UnitsALFUS Fuselagelength 13.27 mHFUS Fuselagemaxheight 1.92 mWFUS Fuselagemaxwidth 1.2 mDMAX Fuselagemaxequivalentdiameter 1.92 m

WINGDESCRIPTION

FUSELAGEDESCRIPTION

Variable Description Value UnitsSh Horizontaltailarea 15.21 m^2ARh Horizontaltailaspectratio 3 -Bh Horizontaltailspan 6.755 mCRh Horizontaltailrootchord 2.25 mTCh Horizontaltailthicknessratio 0.08 -ALCh DistancefromwingtoHTMAC 10 m

Variable Description # Value UnitsSv Verticaltailarea 2 5.88 m^2ARv Verticaltailaspectratio 1.5 -Bv Verticaltailspan 2 2.613 mCRv Verticaltailrootchord 2 2.25 mTCv Verticaltailthicknessratio 0.101 -ALCv DistancefromwingtoVHTMAC 10 m

HORIZONTALTAILDESCRIPTION

VERTICALTAILDESCRIPTION

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Figure 32. MARVEL Cabin

Figure 33. MARVEL Wing

Figure 34. MARVEL Empennage

C. Aerodynamics This discipline will analyze the aerodynamics of the aircraft to characterize its behavior under

different flow conditions which correspond to the conditions which the aircraft is expected to operate. In detail, the aerodynamics team starts its analysis by defining the wing parameters such as airfoil, planform shape, and others. The analysis also studies the entire aircraft to determine the total lift, drag, and pitching moment, as well as the Reynolds number regime.

1. Wing Characteristics

For this discipline, it is important to remind that Martian atmosphere is extremely thin and cold. In fact, it can be said it resembles the Earth’s at 100,000 ft. Therefore, it is necessary to find an airfoil that has a high lift coefficient for a high Mach Number–low Reynolds number (<150,000) flow. Multiple

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airfoils were considered for this. At the end, the decision was between the S1223 – used by the Zephyr Concept (see section III, Previous Mars Aerial Platform) – and the Wortman 63-137.

In order to select the definitive airfoil, the lift at various Reynolds number and angles of attack was analyzed. This was done using Mark Drela’s XFLR5 software. The results, shown below, show how the S1223 airfoil produces more lift than the 63-137. However, the S1223 has an unsteady performance at low Re and high 𝛼 (see red boxes), whereas the 63-137 seems to have a smoother performance. Therefore, the 63-137 was chosen as the final wing airfoil. A more detailed analysis shows how the 63-137 airfoil also provides more L/D and a lower drag. Such comparisons can be seen on the appendix of this report, section C.

Figure 35. Variation of Lift for Various Re and AOA for the S1223 and 63-137 Airfoils [4]

Figure 36. FX 63-137 Airfoil Shape [4]

The total wing analysis brings together aerodynamic, stability, and structural design tradeoffs. Furthermore, there is the limitation imposed by the aeroshell, which constraints the wing span and wing area. Based on these, and after analyzing the wing performance, the aerodynamics decided for a wing area of 110 m2, a wing span of 44 m, and a mean aerodynamic chord (mac) of 2.5. These parameters yield a wing with an aspect ratio of 17.6. This value was corroborated by the Structures and Weights and the Stability and Control disciplines. The planform shape was inspired by high altitude gliders, such as the Perlan 2 Glider.

The wing was chosen to be placed at a high position with respect to the fuselage. Such configuration ensures the wing is as far away as possible from the exhaust gases from the rocket. It also provides more visibility for the pilot when scanning for landing sites. These benefits, though, come with a drag penalty. At the wing tips, the wing was chosen to have a small (5.79°) dihedral and a small sweep angle as well. These features improve lateral stability, increase the efficiency, and results in less induced drag.

Note that the sizing process had determined a wing area of 84 m2. This new area is indeed larger, but it will result in a lower wing loading, which is still within the solution space. A detail weight analysis, shown later, will confirm that the wing loading is satisfactory. It is also worth noting that the wing span is within the capabilities of a theoretical 9 m aeroshell.

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2. Full Body Aerodynamics

The full aircraft aerodynamic analysis consisted on determining the aircraft drag, lift, and pitching moment. This analysis was done using analytical and computational tools, such as the drag buildup method and the vortex lattice method (XFLR5) respectively. As a result, the MARVEL’s drag polar, lift curve, span loading, lift-to-drag, and pitching moment were obtaining. For a more detailed aerodynamic analysis, please refer to Yasir Rauf’s report [4].

Figure 37. MARVEL Drag Polar, Analytical and Computational Results [4]

Figure 38. MARVEL Lift Coefficient vs. AOA [4]

Figure 39. MARVEL Span Loading [4]

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Figure 40. MARVEL Lift-to-Drag [4]

Figure 41. MARVEL Pitching Moment [4]

D. Propulsion and Energy Propulsion and Energy is one of the most critical design areas for this aircraft. Mars’ low density

affects the amount of thrust that could be generated if a propeller was to be used. The atmospheric composition (95% CO2) makes it impossible to use conventional air-breathing engines. If a rocket was to be used, the propellant would have to be transported from Earth, making the operational cost unbearable. The distance between Mars and the Sun also complicates using solar cells as its energy source. Finally, alternative propulsion systems (internal combustion engines, CO2 based rockets, fuel cells, and others) still need more development before being considered for an aircraft of this magnitude. In conclusion, each system comes with a tradeoff, and therefore a careful analysis must be done to select the most appropriate system.

Therefore, this discipline will select the type of propulsion system to be used. It is important to remind the reader that, from the parametric sizing process, it was concluded that the aircraft needs a VTOL system. Such system will be analyzed independently. On the other hand, the design of the propulsion system for the cruising and maneuvering stages will be shown. The analysis will be completed with the power plant and drivetrain selection.

1. Vertical Takeoff-Landing Analysis

The idea of using a VTOL system for a Mars aircraft has been conceived before. One of the designs presented by Developmental Science Inc, in 1978, proposed a landing and takeoff maneuver that can be used for a preliminary analysis. Basically, the ground rolls are so large that is simply not feasible to have an aircraft horizontally takeoff or land. As an illustration, the MARVEL aircraft would need a takeoff/landing field of 17.25 km and 10.4 km respectively.

Ryan Manns [63] created a MATLAB script to determine the amount of fuel consumed during a vertical ascend/descend given that the altitudes are provided, along with an aircraft mass, and a nozzle mass flow. The results obtained from this program showed that the fuel fraction ranged from 0.2 to 0.3 of the total aircraft mass, depending on the type of propellant and its specific impulse. This code can be found on the appendix (section E), and is described in more detail in Ryan Manns’ report.

After discussing it with the other design disciplines, it was noted how the ascend/descend profile was assumed to occur completely vertical at a zero speed. In reality, once the aircraft ascends to a given altitude, it must gain enough speed to sustain flight. Similarly, when landing, the aircraft must come to a complete stop before descending.

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For the takeoff phase, one strategy considered was to ascend to an altitude of 2000 m, then fall to 1000 m while gaining speed. Once the aircraft reached the stall speed, then it would perform a pull up maneuver to transition into cruise flight. From a potential-kinetic energy conversion standpoint (mgh=1/2mV2), a drop of 1000 m would provide a speed of 86.1 m/s, which is well above the stall speed. For the landing, a deep stall maneuver could be used to slow down the aircraft. Then, a rocket-controlled descend would guarantee a safe landing.

The issue with the previous maneuvers is that they might be too dangerous for both the human and the aircraft itself. Humans can, naturally, sustain accelerations around 49 m/s2 (5G’s on earth). By using special suits and after some training, this limit can be pushed to 88.3 m/s2 (9G’s on earth) [82]. There are many more variables involved in the effect of acceleration on the human body. Thus, extensive flight-testing is needed to ensure that these limits are never exceeded. The aircraft structure, similarly, would need to be tested against such loads. Here, there is an evident contradiction between the desire to keep the aircraft structure as low as possible and the necessity to make the aircraft structure robust enough to withstand this maneuver.

Another alternative is to use a rocket/jet-assisted takeoff. These systems, discussed on the technology studies section of this report, have been employed in many cases to reduce the takeoff field lengths, particularly on military aircraft. This technique allows keeping the takeoff length as short as possible (this system is sometimes termed as zero-length takeoff) while gaining enough speed to sustain flight. Also, this maneuver puts less stress on the structure and the astronaut. Finally, it is more feasible to be certified by NASA or any other space agency, as it has been used on Earth before. As for the landing, the deep stall maneuver could still be used to slow down the aircraft, followed by a controlled rocket descend onto Martian surface.

For the takeoff and descent system, it was decided to use Curiosity rockets, which provide up to 3300 N of thrust. Three rockets would be more than enough to provide enough thrust (almost 10 kN). From those, only 57% is used to lift the aircraft, while 12% is used to stabilize the aircraft during the takeoff/landing maneuvers. The remaining 31% can be used in case of an emergency. The performance analysis concludes that approximately 250 kg of fuel is required to power a round-trip mission (two takeoffs, two landings), which is only 17% of the total aircraft mass. Note that such fraction is well within the initial 0.3 mass fraction used during the sizing procedure. Therefore, this rocket-assisted takeoff and landing maneuver is feasible. The following two figures show the takeoff and landing profiles maneuvers.

Figure 42. Rocket-Assisted Landing and Takeoff Profiles [63]

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2. Cruise and Maneuvering Flight

For cruise flight, the propulsion and energy team has decided to use propellers. Most of the previous design projects use this type of system due to its simplicity. Moreover, Aurora Concepts’ team has experience on analyzing propellers, as well as the necessary tools for such analysis. Again, the two downfalls of using propellers are, first, the resulting low thrust due to the low atmospheric density and, second, the need to fold and pack the propellers in order for them to fit into the aeroshell. Given Mars atmospheric conditions, the resulting propeller design must have a large blade diameter, but at the same time it must rotate at a small RPM to avoid the tip falling into the transient regime.

The propeller design will be depend on the amount of thrust required to sustain cruise flight and perform some basic maneuvers (turning and descent). From FAR requirements, as outlined in Loftin [51], aircraft with stall speeds larger than 70 mph need to have enough power so that the climb rate is at least 0.02Vstall

2. Then, assuming the thrust required is equal to the necessary force to overcome the drag, and that power required is simply the thrust multiplied by the flight speed, the following plots can be created.

Figure 43. Thrust and Power Requirements for Different Altitudes at Mars

The previous plots were created with a script that can be found on the appendix of this report. They show the amount of thrust and power that the propeller/engine must supply (red line), depending on the flight speed. The plots above were obtained by assuming the flight speed is 150 m/s (maximum speed). Therefore, if climbing/maneuvering was to be required at the maximum speed, the propellers must provide a thrust around 580 N and a power of 87 kW.

For the analysis, a propeller code was found on Internet that allows changing the atmospheric conditions to Mars atmosphere. It has airfoil shapes built-in, and it allows inputting the number of blades, the propeller diameter, and the forward flight speed. For more details on such code and its functioning, please refer to the Propulsion and Energy report.

At this point, the most optimal conditions consist of using two four-bladed propellers, 4.25 m of diameter, at a RPM 800. The thrust output, at sea level, is around 511.25 N. It is found, using the code above, that the maximum achievable speed is 152 m/s (for straight-leveled flight). This thrust, on the other hand, is enough to provide climbing/maneuvering for a flight speed of 138 m/s. However, those flight speeds would result on regions of the propeller blade running into the supersonic regime, which is undesirable. It is important to note that the current program uses conventional propeller blades for its calculations. An advanced propeller, designed for high-altitude flight, would solve this issue. For now, the resulting effect on the aircraft operational profile is shown on the performance section of this report. A CAD model and a summary of the propeller design are shown on the appendix of this report.

0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000350

400

450

500

550

600

Height, m

Thru

st, N

Thrust Available and Thrust Required

Min. Thrust RequiredThrust Available

0 500 1000 1500 2000 2500 3000 3500 4000 4500 500055

60

65

70

75

80

85

90

Height, m

Powe

r, kW

Power Available and Power Required

Min. Power RequiredPower Available

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3. Power Plant

From the energy production point of view, the goal is to provide enough energy to power the propellers during different flight conditions. From the experience of the team, and given the considerations presented on the technology studies section, it was decided to use rechargeable batteries as the power plant for this Mars aircraft.

The battery selection process depends on the mission profile and the thrust requirements, both previously addressed. The mission range yields an estimate for the mass of the batteries with respect to the aircraft total mass. Such calculation (equation derived by Hepperle, Ref. [64]) is repeated here, for convenience. The mass fraction of the batteries is found using the following equation:

𝑚!

𝑚!=


=400000

300(3600)(0.85)(1/3.71)(24)= 0.0674

When using batteries, it is necessary to address the existing tradeoff between the battery’s life cycles and its energy density. At the same time, there are issues related with their discharge rates and the heating they produce.

For the selection of the battery, the propulsion and energy team has contacted battery companies to obtain full data on state of the art batteries. Oxis [83] was kind enough to provide us with information about their ultra-light Li-S batteries. As June of 2016, these batteries proved to have energy densities of 325 Wh/kg, with an expected life of 250 cycles, operating temperatures as low as -30C˚, and a charge time of 3 hours. Such batteries are still on evaluation stage. In the future, high performance batteries will provide an energy density of 450 Wh/kg, a life of 1500 cycles, and a charge time of only two hours. These are expected to be ready for evaluation in June of 2018, and for production on September of 2019.

The energy analysis was done assuming high performance batteries are used. Then, the necessary battery mass and volume were found given the performance requirements. The following table summarizes these parameters. A detailed description on how these were calculated can be found on Ismael Sanabria’s report [84].

Table 19. Batteries Characteristics [84]

As for recharging the batteries, two options were studied: solar cells and radioisotope thermoelectric

generators (RTGs). Both options are proven solutions to generate power in space. Satellites commonly use solar cells to power their systems during their long missions (various years), whereas Mars rovers used RTG. Solar cells have the limitation that the amount of solar irradiance is only 44% of Earth’s. Furthermore, they are limited by the weather conditions. RTGs, on the other hand, provide a steady power output for multiple years and require minimal maintenance. However, the amount of power generated by these systems is considerable less than solar cells. The propulsion and energy team estimated that, for the MARVEL, solar cells and RTGs would need 6 hours and 30 days for a full charge. Regardless of which system is used, there would need to be an auxiliary power plant that needs to be included into the mission.

Voltage 200 NACapacetence 487.39 VCellVoltage 2.1 Ampere-hoursCellCapacity 50SystemPower 97477.91 Watt-hoursSystemWeight 216.62 kgSystemVolume 194.96 LEndurance 2.2 Hours

BatteyInformation

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4. Drivetrain

The design of the drivetrain is pretty straightforward, as there is an extensive knowledge base on this aspect. Basically, the drivetrain is designed to make sure the engine power is transformed into mechanical energy on the propeller in the most efficient way as possible. A high efficiency guarantees that the propulsion system can maintain the desired endurance/range.

The propulsion and energy team has determined that the mass of the drivetrain is 78.72 kg for each propeller. This includes the engine, controllers, and propeller. Hence, for two propellers, the total mass would be 157.44 kg. The overall system efficiency is 68%. The following table summarizes the drivetrain design parameters. For more details on these calculations, refer to Ismael Sanabria’s report [84].

Table 20. Drivetrain Design Parameters [84]

E. Structures and Weights

The structures and weights discipline determines the weight of the aircraft depending on the aircraft dimensions and the materials selected. In this process, it is necessary to make sure that the aircraft’s resulting weight is within the acceptable range as defined by the synthesis discipline. The structure must also be robust enough to withstand aerodynamic and inertial loads during flight. One case of special interest, for example, is the analysis during the takeoff and landing maneuvers. Then, the engineer designs the structure to be strong enough to sustain such loads. Finally, the structure must be simple and small so that it can be fitted within the EDL system.

It is important to remind the reader that this project is kept within the conceptual design phase. Therefore, it is not necessary to define the entire internal configuration of the aircraft (number of ribs/wings/spars on the wing, for example). Rather, this discipline will define the external features of the aircraft. As a result, there will be a strong connection with the aerodynamics and stability disciplines, as they are the ones who have major influence on the wing and the tail, respectively.

1. Fuselage Shape

This discipline has determined that the most optimal fuselage shape is a tadpole. It provides 30-40% less wetted area than a frustum (hence, less drag). Its shape is also beneficial to sustain a laminar boundary layer over its surface. The only downside is that the fuselage volume is limited compared to other options. This, however, might be beneficial for saving mass.

Mass (kg) EfficiencyGear Box 18.1 0.95Propeller 9.9 0.75Controller 9.52 0.99Motor 45.93 0.97Total 157.44 0.68

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Figure 44. Tapdole CAD Model [17]

2. Mass Estimation

The mass estimation is done using Nicolai [49] and Roskam [85] methods, giving some allowance for the use of composites materials. This method has been implemented using a MATLAB script. Then, using the geometry inputs mentioned before, it was determined that the maximum gross mass is 1467.7 kg, with an empty mass of 625.7 kg. Compared with the sizing results (1489 kg and 691.75 kg), the difference is 1.43% and 9.55%. Recalling the wing area is 110 m2, the resulting wing loading is 13.34 kg/m2, which is lower than the 17.83 kg/m2 predicted by the sizing process. Still, this value within the convergence zone. The detailed mass estimation and mass fractions are shown in table 20 and figure 43, respectively.

Table 21. Detailed Mass Description [17]

Variable Description Value UnitsWCRW Crewweight 75 kgWPAY_MAX Maxpayloadweight 250 kgTOGW Take-offgrossweight 1467.7173 kgWFUEL Fuelweight 592 kgAMZFW Maxzerofuelweight 625.72 kgOWE Operatingweightempty 625.72 kgOEW Operatingemptyweight 625.72 kgWSTR Structureweight 476 kgWSYS Systemsweight 77.07 kgWWING Wingstructuralweight 259.46 kgWFUS Fuselagestructuralweight 105.32 kgWNACC Nacelleweight - kgWHT Horizontaltailweight 41.37 kgWVT Verticaltailweight 11.5 kgWP Engineweight 150 kgWLG Landinggearweight 59.7 kgWFC Flightcontrolsystemsweight 77.1 kgWLS Lifesupportsystemsweight 145 kgWFUR Furnishingweight 17.5 kgFF_TOTAL Fuelfraction 0.35 -WSYS_TOGW Systemsweightfraction 0.06 -WSTR_TOGW Structuralweightfraction 0.33 -WR Weightratio(TOGW/OWE) 0.48 -

WEIGHTDESCRIPTION

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Figure 45. Aircraft Mass Distribution [17]

3. Load Estimation

The analysis of cruise loads was determined based on the wing loading. For this stage, the analysis was simplified by assuming the wing could be taken as a cantilever beam. During the aerodynamic analysis, the wing loading distribution was determined to be as shown in the following figure:

Figure 46. Wing Loading Distribution [17]

Figure 47. Shear and Bending Moment Distribution [17]

The maximum shear and bending moment are experienced at the wing support (the fuselage). Their values are, respectively, 260 N and 2800 N-m. Such values are relatively small, and the resulting stresses are well within the limits for the materials commonly used on aircraft structures. A more detailed analysis is presented on the structures and weights report [17].

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Due to the time limitation on this project, and the fact that the main focus was to determine the feasibility of the Mars aircraft, the propulsion and energy team was limited to prove that rocket propulsion was viable during critical phases (takeoff and landing). Recall that this team concluded that, of the total fuel, only 57% is used to lift the aircraft, 12% is used to stabilize the aircraft, and the remaining 31% is saved and only used in case of emergency. Therefore, if a complete load analysis is to be performed, it would be necessary to, first, determine the location and size of the thrusters. This process is left for future stages of this project.

F. Stability and Controls The stability and control is in charge of ensuring that the aircraft is suitable for flight (stable and

controllable). Mars’ low Reynolds number environment decreases the aerodynamic forces, and therefore larger control surfaces are necessary to provide stability. This, however, imposes a challenge, as it is also desired to minimize the drag as much as possible. On the other hand, the aircraft volume is limited by the aeroshell limits. The other challenge is to ensure the aircraft will be controllable during the VTOL stages. The results will be compared with an Earth airplane in order to visualize the challenge of providing control on Mars.

The author apologizes to the reader because the Stability and Control analysis is not complete. The lead of this discipline, Justin Kenna, was not able to complete his work within the time frame due to personal reasons.

1. Tail Configuration

The work of this team started with defining the tail configuration, size, and placement. The team conducted an analysis of conventional configurations such as the canard, flying wing, and tail-aft. The canard and flying wing are not common in general aviation aircraft. Since one of the main design drivers of this project is simplicity, they were automatically disregarded. Then, tail-aft designs were analyzed. Configurations such as the t-tail, v-tail, inverted v-tail, and a conventional empennage were considered.

Preliminary, a conventional tail-aft empennage was selected. Such configuration uses a horizontal and a vertical stabilizer. The first is used to generate pitch, which provides longitudinal stability of the aircraft. The vertical stabilizer, on the other hand, provides directional stability.

For the tail sizing and placement, the team used a tail volume method. Then, a database of sailplanes was created, emphasizing on its control surfaces (see Ref. [33]). Sailplanes were used since they have high AR surfaces, thus they are closer to the current MARVEL design. Such database was scaled to account for Mars atmosphere, using a Reynolds number matching criterion. The results were optimized to minimize the empennage weight.

Initially, the horizontal tail had a span of 6.76 m and a chord of 2.25 m (total area of 15.2 m2). The vertical tail, on the other hand, had a span of 4.2 m with a chord of 2.8 m. As for its location, the distance between the aircraft c.g. and the tail’s quarter chord was determined to be 10 m. However, the vertical tail was so large that it would crash with the ground. This led to the decision of using a twin vertical tail. As a result, the horizontal tail remained constant, whereas the vertical tail span was reduced to 2.61 m and a chord of 2.25 m (each with an area of 5.88 m2). The airfoils selected were the S8025 (inverted) and the SD8020 (symmetric) for the horizontal and vertical profiles, respectively.

2. Control Surface Sizing

The control surfaces must provide enough force to trim the aircraft and respond to the pilot inputs when maneuvering is necessary. The critical stages are the takeoff/landing, and the situation where one engine is inoperative. The sizing process was performed using the roll authority criterion and

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Gudmundsson strip integral method [33]. The latter is used as an aileron sizing method, but it can also be used for elevator and rudder sizing. Alternative control methods, such as control moment generation, were also being studied but ruled out due their complexity.

As a comparison, the team selected the Piper P-28 Super Cub. Such airplane is a bush plane, used in rural areas to provide access to zones with little to no infrastructure. The Super Cub and the MARVEL have similar missions and similar payload capabilities, and therefore it is possible to make a comparison between both systems. The following table compares the dimensions of the ailerons (roll control), horizontal stabilizer (pitch control), and vertical stabilizer (directional control).

Table 22. MARVEL and Super Cub Control Surface Comparison

It is more than evident how the low Re number environment in Mars requires such a tremendously

increase in the size of the control surfaces.

Figure 48. Piper Super Cub [86] [87]

Control Surface Description MARVEL Super Cub

Span (m) 6.76 3.2Chord (m) 2.25 0.509

Surface Area (m^2) 10.64 1.632Moment Arm (m) 10 3.94

Span (m) 2.61 1.35Chord (m) 2.25 0.511

Surface Area (m^2) 4.115 0.39Moment Arm (m) 10 4.075

Chord (m) 0.75 0.367Span (m) 0.3 0.227

Surface Area (m^2) 9 2.54Moment Arm (m) 6.75 0.931

Horizontal Stabilizer

Vertical Stabilizer

Aileron

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3. VTOL Stability Analysis

As it was mentioned above, the takeoff and landing are probably the most critical stages for the MARVEL mission profile. The stability and control team was going to provide an initial sizing of the thrusters in order to maintain altitude during landing. It was also going to determine the thrusters’ location to ensure the maneuver is safe. As a reference, NASA’s design MATADOR also conceived using thruster for the VTOL stages. However, the amount of information available about such design is limited. Again, the author apologizes since the VTOL stability analysis was not completed on time. Anyway, the propulsion and energy team determined that only 13% of the fuel was necessary to control the aircraft during this stage.

4. Static Stability Analysis

Once the entire aircraft geometry and its control surfaces are defined, the final step is to ensure the aircraft meets the criteria for static stability. Dynamic stability, on the other hand, often requires advanced simulation techniques and even real life testing. Therefore, it will not be included in this project. The static stability criteria are summarized in the following table:

Table 23. Static Stability Criteria

The determination of the stability derivatives was to be done using Roskam’s methods, as outlined in

Ref. [88]. However, such calculations were not carried out. Still, the aerodynamic department was able to determine that the pitching moment decreased with angle of attack, which is the requirement for longitudinal stability. On the other hand, the wing was designed with a small dihedral at the wing tips in order to improve lateral stability. Finally, the vertical stabilizer was found to be large enough to control the aircraft on its axes. Even so, a proper stability and control analysis is necessary to confirm these statements.

G. Performance The performance team conducts an analysis of the final aircraft configuration to ensure that it can

complete the mission. The results from performance are used to identify the areas in which the design must be improved. Finally, this discipline will present a flight envelope diagram.

1. Takeoff and Landing

The takeoff and landing maneuvers were previously shown in the propulsion and energy section of this chapter. The profiles are repeated here, for convenience. It was assumed that the rockets used are the same as Curiosity, which provide up to 3300 N of thrust (each). Then, the profile was analyzed from a zero velocity start until the aircraft reaches its stall speed. The following figure shows the aircraft position during such maneuvers. For more details on how this calculations were done, please refer to Ryan Manns report [63].

Stability Derivatives

Desired Output Description Incidence On

<0 Pitching Moment due to AOA Longitudinal>0 Yawing Moment due to Sideslip Directional<0 Rolling Moment due to Sideslip Lateral

𝐶"#

𝐶$%𝐶&%

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Figure 49. Rocket-Assisted Takeoff and Landing Profiles [63]

The main conclusion here is that the aircraft is, indeed, capable of using a rocket-assisted (zero-length) takeoff and landing for its mission. For a round-trip mission (two takeoffs, two landings), the fuel using is less than 30% the total aircraft mass, and so it is feasible. Thus, the mission requirements outlined in the request for proposal are met.

2. Cruise and Maneuvering Flight

For the cruise and maneuvering analysis, the performance team worked closely with the propulsion and energy team to ensure the thrust available was larger than the thrust required. A preliminary analysis was presented in the parametric sizing section of this report. However, the performance calculations are the definitive assessment of these requirements.

Using the latest gross mass, wing size, and aerodynamic parameters, the performance team was able to find the theoretical stall, minimum drag, and minimum power velocities, for altitudes ranging from 0 (ground level) to 3000 m. The results of most interest are those at an altitude of 1000 m, as this is the MARVEL’s cruise altitude. The following table, courtesy of Nic Dwyer, shows these results.

Table 24. Ideal Flight Velocities [89]

Altitude(m)StallVelocity

(m/s)

VelocityminDrag/maxRange

(m/s)

VelocityminPwr/maxEndurance

(Hr)0 62.666 81.949 62.268250 63.342 82.833 62.940500 64.026 83.727 63.619750 64.717 84.631 64.3061000 65.415 85.544 65.0001250 66.121 86.467 65.7011500 66.835 87.400 66.4101750 67.556 88.343 67.1262000 68.285 89.296 67.8512250 69.021 90.260 68.5832500 69.766 91.234 69.3232750 70.518 92.218 70.0703000 71.279 93.212 70.826

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Then, for 1000 m altitude, a detailed analysis showed how the thrust produced by the propellers (orange line) is more than the minimum thrust required to sustain steady flight (blue line) for a range of velocities. This analysis considers not only the parasite drag but also the drag due to lift, and therefore is more reliable than the calculations performed by the synthesis team. Other relevant figures, such as power, range, and endurance at 1000 m are also presented.

Figure 50. Thrust Analysis at Cruise Altitude (1000 m) [89]

Figure 51. Power Analysis at Cruise Altitude (1000 m) [89]

Figure 52. Range and Endurance at Cruise Altitude (1000 m) [89]

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From the previous figures, it is shown how the aircraft can perform in the range of speeds of 65 m/s (stall speed) up to a maximum speed of ~ 125 m/s. The maximum range is 779 km at a speed of 85 m/s, with an endurance of 2.2 hr. Since the mission requirements called for a minimum range of 4000 km and endurance about 2 hours, it is concluded that the aircraft satisfactorily meets the mission. This will be further shown on the next section.

3. Operational Flight Envelope and Mission Analysis

This section presents two important figures: the operational flight envelope and an example of a mission scenario. The first is a figure shows the altitude vs. speed limits, allowing the reader to visualize the capabilities of the MARVEL. It consists of three curves: stall speed, operational ceiling, and maximum thrust. However, this is just a preliminary estimation. A more complete operational envelope takes into account structural limits as well.

Figure 53. MARVEL Flight Envelope [89]

The second figure is a mission scenario. It is an idealization of the aircraft’s profile. It is obtaining by combining the rocket-assisted takeoff/landing phases with the climbing, cruise, and descent phases. To get this figure, it was necessary to analyze the rate of climb and descent at small altitude increments. Such analysis is shown in more detail in Nic Dwyer’s report.

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Figure 54. MARVEL Mission Scenario [89]

Table 25. Mission Summary

Takeoff Speed (m/s) 70

EOS TO Altitude (m) 50

Avg. Rate of Climb (m/s) 5

Cruise Altitude (m) 1000

Cruise Velocity (m/s) 85

Avg. Rate of Descent (m/s) 3.5

Start of Landing Altitude (m) 100

Landing Speed (m/s) 70

Total Range (km) 335Total Endurance (hr) 1.1

Mission Analysis Summary

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VII. Discussion The aircraft presented on the previous chapter proofs that manned flight on Mars is possible. The

resulting aircraft is considerable large as a result of Mars’ low atmospheric density. Since it was assumed that no infrastructure was available, a rocket-assisted landing / takeoff became a necessity. As a result, there are several considerations that must be addressed.

In first place, it was assumed that this aircraft was to be delivered using SLS Block 2B, which has a payload fairing of 10 m. Therefore, the MARVEL would have to be developed once the SLS is ready. Such system is still far from being operationally ready, and so does the MARVEL. An option that would seem more viable would be to send the aircraft by parts, using more than one aeroshell. A rocket such as the Falcon Heavy can accommodate more than one aeroshell. DSI’s proposal, for instance, planned to deliver about 4 aeroshells on a single spacecraft. However, the aircraft would need to be assembled on the planet. This implies that there is already a human base present, with the tools required to assemble such aircraft. Mission architectures, such as Mars One, estimate a human base will be ready by 2029. However, a base capable of assembling and testing an aircraft is far more complex. For this reason, this option pushes back the MARVEL even further in the future.

On the same line of thought, it is necessary to consider the logistics behind the operation of an aircraft. It is necessary to have a power supply station on the ground to recharge the batteries. The propulsion and energy section of this report discussed about using RTG or solar cell systems. The first provides a continuous but small power supply, whereas the second requires large systems that must be transported under special conditions.

On the other hand, the maintenance of the aircraft requires the availability of tools and other service systems on Mars, in addition to people trained for such tasks. This becomes especially important when talking about a manned aircraft, as it must prove to be safe enough for operation. A special figure of merit would then be the maintenance man-hours per flight hour (DMMH/FH). An advance system, such as the F-22, required 10.5 DMMH/FH in 2009 [90]. Such “low” requirement was only achieved after extensive work, process automation, and experience.

As for the landing and takeoff maneuver, there are several operational concerns that would have to be addressed. On Earth zero-length launch systems proved to be feasible but were not incorporated because of the extra challenge that they imply. Also, other technological developments made their use obsolete for military applications. For Mars, however, there does not seem to be another option. Even if it is assumed that humans have already established an advance base on Mars, the airfield lengths would be too large.

In the opinion of the author, the development of aerial platforms for Mars has to start with small unmanned scientific aircraft, such as the ARES. Later, once those aircraft prove to be successful, the next step would be to produce prototypes to the rocket-assisted takeoff and landing concept. A good example would be the Astroplane. In this way, humans on Mars would gain experience with operating and maintaining aircraft. At the same time, more data on Mars aerodynamics would be gained.

With the previous considerations on mind, the author believes that an aircraft such as the MARVEL will not be a reality for the following 20-30 years. The exercise, although interesting, has demonstrated the multiple challenges that have to be overcome for, first, flying an aircraft and, second, designing a spacecraft. The first humans in Mars will probably use vehicles (rovers) to facilitate transportation and exploration. Aircraft, however, will be used as aids on scientific exploration rather than being considered as a serious method of transportation. Finally, it is a necessity to have a fuel production system on Mars. Even with re-supply missions, sending fuel to Mars every two years is not a viable option for any business case.

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VIII. ABET Requirements The Accreditation Board for Engineering and Technology, Inc., is a non-government organization

that accredits universities to guarantee the quality of the programs offered. The organization counts with more than 2200 experts from different scientific and organizational backgrounds [91]. As a result, they have defined a set of outcomes that must be met by each educational program. For instance, for aerospace and related programs, they specify that programs “must prepare graduates to have a knowledge of aerodynamics, aerospace materials, structures, propulsion, flight mechanics, and stability and control. [92]” In specific, the capstone senior design addresses ABET student outcomes C,D, F,G, H, and I, as it will be shown below.

A. Outcome C: Design System, Component or Process to Meet Needs In detail, outcome C reads: “an ability to design a system, component, or process to meet desired

needs within realistic constraints such as economic, environmental, social, political, ethical, health and safety, manufacturability, and sustainability” [92]. The senior design project documented on this report addresses the hearth of such outcome. Before the start of the project, the students researched the most pertinent mission scenarios for a hypothetical Mars aerial platform. As a result, they come out with a request for proposal, which was presented an approved by the course instructor.

Later, the students were divided in groups to respond to such proposal. Such proposal calls for the design of a new system (in this case, a Mars aircraft) within the constraints imposed by the technology readiness level of space exploration and propulsion systems. The project was monitored by the submission of bi-weekly reports. In specific, the midterm and final reports are formal deliverables that show the complete state of the project at different stages throughout the semester. Furthermore, these reports are accompanied by a midterm and a final presentation to a technical audience, a team poster, and a 3D printed flight vehicle configuration.

B. Outcome D: Function on Multidisciplinary Teams The senior design class has been designed to simulate a professional engineering environment. After

the request for proposal has been drafted, the teams were divided into two groups, each one with a Chief Engineer and discipline leads. The teams created a logo and defined a corporate identity. They were organized in a multidisciplinary structure, addressing the major fields of study in aerospace sciences: aerodynamics, propulsion, structures, performance, and stability and control.

The teams had to learn how to properly integrate the disciplines through the design process. In the process, each group created a multi-disciplinary methodology (MDA) that summarizes the team strategy to arrive to an aircraft solution. Such process is not intended to be a “black box” system, where no one knows how it works. On the contrary, it is important that all team members understand their position in the design process.

During this project, there is a discipline named “synthesis”, which is in charge of two persons (all other disciplines had one member only), with the Chief Engineer also playing an important role. The synthesis team was in charge of defining the MDA, and basically providing a guide for the disciplines to perform their analysis. Then, once some results were obtained, the synthesis team combined them and converged into a solution. Often, the outputs from the disciplines would be contradictory. Then, the synthesis and the Chief Engineer would be take a decision based on mission, design, and technology studies and took a decision.

In addition, the team members would meet multiple times throughout the week to work on specific tasks. During the first phase of the project (prior the request for proposal), all members worked and

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presented together their research on Mars: mission opportunities, previous concepts, Mars characteristics, and others. Later, the team members worked together on generating an aircraft idea. During this process, the members went beyond the individual disciplines and worked on multiple tasks at the same time. Then, once the team had an idea of the type of aircraft to be designed, the disciplines started their individual analysis. Finally, after the baseline design was ready, the members got together again to refine their differences and finalize the design.

C. Outcome F: Understand Professional & Ethical Responsibility This is one of the areas were engineering education often needs improvement. It is hard for students

to understand the importance of ethics and professionalism if their work is constrained to the classroom environment. In this project, ethics is addressed in many opportunities. In first place, it is paramount to learn the importance of documenting the work and giving credit to others when necessary. During the research phase it was found that this is not the first time a team attempts to design an aircraft for Mars. Therefore, it was tempting to use some of the previous results. This was addressed by the high standards under which the reports and presentations are evaluated, as well as the professional environment that surrounds the development of the project. All work presented on the reports must be properly documented and backed up with data and a clear design process. Otherwise, such results are not acceptable.

On the other hand, the course instructor made sure to emphasize on professionalism and ethics during his lectures. One of the most fulfilling aspects of the entire senior design course was to hear his experiences when working in both, industry and academia. Different case studies were analyzed, such as the case where a starting company engaged in the design of an aircraft capable of serving the space tourism industry. On that occasion, the Chief Engineer committed a fault when he thought that simply copy pasting an old wing design would be good enough for the project. That was an evident fault against professionalism and ethics, and ultimately led the company to bankruptcy.

D. Outcome G: Communicate Effectively Communication occurs in multiple levels and in different directions. There is communication with the

course instructor / teaching assistant, communication with teammates, and communication with external sources. One important part of this project is to network with professional companies, which required team members to go beyond the university and start sending emails, making phone calls, or rely to any other medium.

It is important that engineers learn how to effectively communicate. This is often one of the areas where engineers fail, as the complexity of their job make it difficult to share with non-technical audiences. In occasions, miscommunications occur even between engineers from the same project. A classic example of such issue is the mishap in units that led to the crash of one space [93]. In that moment, engineers at Lockheed used English units for their calculations, whereas NASA has traditionally using the metric system. As a result, the $125 million Mars Climate Orbiter was lost.

Within the team, the members had to constantly communicate in order to arrange team meetings. It was also important to pass information between the disciplines, as they often required inputs from multiple areas at the same time in order to perform their analysis. Finally, the reports and presentations are an example of the team members documenting their work and sharing it with other people.

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E. Outcome H: Understand Impact of Engineering Solutions in Global & Societal Context This outcome is closely related with the two previous ones: ethics and communications. Engineers

must understand that their work goes beyond the computer, and that it can actually affect the lives of other people. As outcome C addressed, the students must work on solving real life issues, because that is what engineers do. Within that context, an engineer’s work has produced tremendous effects on the past. Sure enough, engineers will also have a lot of impact on the future.

In this project, this work is also addressed on the cost analysis section. The request for proposal stated that this Mars aircraft has to be fitted into promising commercial (Space X) or scientific (NASA) projects. Therefore, there has been an extensive research on those companies, the past, the present, and the future of Mars missions.

F. Outcome I: Recognize the Need & Ability to Engage in Lifelong Learning Finally, one of the most important outcomes from this project is to learn what is like to produce

professional-quality reports and presentations. In this process, the students have learned about the high standards that must be held when being an engineer. In this report, that process started since day one, when the team started to build a data base on Mars and space exploration.

Of this report, a big portion is dedicated to the literature review. Engineering work would be impossible if it was not for the previous works done on the field. It is important that senior or newly graduate students entering into the workforce learn the importance of performing a proper review, and the importance of creating a database (DB) and a knowledge base (KB). It is also necessary to learn the lessons from the past in order to move forward. Also, students must learn how to properly process and organize the information in order to communicate it. This report is a proof of such process.

After the project is finalized, the student has gotten a grasp of what being an engineer is like. In fact, the senior design report is probably one of the best resources to demonstrate the quality of the work of a future work aspirant.

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IX. Conclusion This report is about the conceptual design of a Mars aircraft capable of carrying one person and flying

for at least 400 km in a round trip mission. The design process started with an extensive literature review, which allowed the team to clearly state the purpose and mission of this project. Then, the team was organized into six design disciplines, although the team members have been working on multi-disciplinary tasks. There has been an important effort to ensure all disciplines standardize their methodologies, with the goal of clarifying the information flow.

Mission, design, and technology studies were conducted, and every stage of the design process has been documented. This consists on the screening of possible space launchers that could carry the aircraft and which, at the same time, impose size and weight constraints. The design studies consisted on the identification of the critical design points, the initial aircraft concept considered and, more importantly, the parametric sizing process. Such process was conducted using Loftin methodology, adapted for low wing loading aircraft and with some slight modifications given the special mission requirements behind this design process. Finally, technology studies were conducted on the areas of propulsion systems and rocket-assisted takeoff and landing alternatives.

The sizing process determined the aircraft had to have wing loadings below 17 kg/m2, and power loadings below 20 kg/kW. Such results are similar to Earth’s high altitude gliders. This gave the team an idea of the type of aircraft that had to be designed. The sizing process also showed that a horizontal short takeoff and landing was not feasible, and therefore an alternative had to be designed. Therefore, the team concluded it was necessary to use a hybrid system, using a rocket-assisted takeoff / landing and a propeller-driven cruise flight. The rockets were to be powered by a propellant / bipropellant that offered high specific impulses and a high technology readiness level.

Based on these results, the individual disciplines started their analysis, consisting on: aerodynamics, structures, geometry, propulsion, performance, and stability and control. The final design has a wing area of 110 m2, a mass of 1478 kg, and a power around 60 kW. More detailed aspects of the final design are thoroughly documented on its respective section on the report. The analysis is concluded with a mission scenario that proved that the aircraft does meet the mission specifications.

There are multiple lessons learned from this design project. Beyond the technical aspects behind the design of an aircraft for Mars, this project also touched on the areas of team management, communication, professionalism, and ethics. More specifically, this project properly addressed ABET outcomes C, D, F, G, H, and I. These aspects are equally important and as challenging as the engineering background behind it. Finally, this project has led the author with lessons that will be extremely useful for his professional life, and serves as the perfect closure for its undergraduate experience.

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X. Appendix

A. Loftin’s Design Methodology, Propeller-Driven Aircraft Loftin’s design methodology consist on determining the aircraft size (wing area S), weight (W), and

power (P), starting from a set of performance requirements and some baseline (assumed) aerodynamic parameters. It makes use of historical correlations obtained from a pool of three general aircraft configurations, defined as follows:

• Class I: Internally braced monoplanes with retractable landing gear (the one that is relevant for this project).

• Class II: Monoplanes with fixed landing gear and (1) internally braced wings, (3) wings with a single strut on each side, (3) wire-braced wings.

• Class III: Biplanes and multistrut monoplanes. Then, the design process flows in the following matter:

• Aircraft Speed Prediction. • Airport Performance. • Climb Performance. • Matching Procedure. • Weight, Range, Fuel Fraction, and Sizing.

1. Aircraft Speed Prediction By analyzing data from the aircraft classes defined above, it was found that there is an empirical

correlation between the speed (V) and the power index (Ip), which at the same time allows finding the wing loading (W/S) and the power loading (W/P) without knowing the drag coefficient. The figure below, extracted from Loftin’s Publication, shows the linear relation for previous planes of Category I:

Figure 55. Variation of Aircraft Maximum Speed with Power Index for Class I Airplanes [51]

𝑉 = 𝐾𝐼!

𝐼! = (𝑊/𝑆𝑊/𝑃

)1𝜎

!

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Where K is equal to 170 for Class I airplanes. It is important to remark that these values were based on maximum gross weights. As for the spread of the data, it is within +/- 10% the trendline, which is acceptable at these stages. The differences might be due to a misinterpretation of historical data. Note that an airplane that lies above the mean line has a higher aerodynamic efficiency than the mean.

Drag characteristics can be found by:

𝐶! = 𝜂 77.3 !(𝐼!𝑉)!

𝐶!,! = 0.9𝐶!

These drag equations were used to find minimum, average and maximum drag coefficients for the pool of aircraft used during the analysis. For class I aircraft, the average zero-lift drag coefficient was found to be 0.023. On the other hand, induced drag is around 10% of the total drag coefficient, and can be expressed as a function of the lift coefficient and aircraft parameters as:

𝐶!,! =𝐶!!

𝜋𝐴𝑅𝑆

Then, the speed can be plotted as a function of the wing loading parameter !/!!

for different values

for the induced drag coefficient as follows:

Figure 56. Boundaries of Speed with Wing Loading Parameter [51]

2. Airport Performance

a) Stall Speed The first parameter for the airport performance analysis requires a determination of the stall speed, which can be

determined, in mph, as a function of the wing loading parameter and the maximum lift coefficient:

𝑉! = 19.73 (𝑊/𝑆)1𝜎(

1𝐶!,!"#

)

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Then, for wings with little or no sweepback, moderate taper ratio, airfoils of moderate thickness, conventional design, and simple trailing-edge flap systems, the following correlation exists:

Figure 57. Variation of Stall Speed with Wing Loading Parameter, flaps-up/down [51]

There is a limitation on the values for the maximum lift coefficient. As it was mentioned before, this correlation does not include the use of leading-edge high-lift devices.

b) Landing Field The landing field distance is defined as the horizontal distance from the point at which the aircraft is 50 ft above

the ground to the point of the runway at which the aircraft is brought to stop. According to FAR, part 23, there is no minimum approach speed for aircraft of less than 6000 lb gross weight.

The following two figures must be used simultaneously to determine the ground run distance as a function of the square of the stalling speed (in mph), and the landing field length (in ft) as a function of the landing ground run distance. Then, for a specified landing performance, it is possible to obtain limits for wing loadings.

Figure 58. Landing Ground Run Distance vs. VS

2 [51]

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Figure 59. Landing Distance over a 50-ft Obstacle vs. Landing Ground Run Distance [51]

c) Takeoff Field Performance Similarly to the landing field, the takeoff distance is defined as the horizontal distance from the beginning of

ground roll to the point at which the aircraft reaches an altitude of 50-ft. The field length is proportional to the takeoff parameter by the following relation:

𝑙!,! ∝𝑊/𝑆

𝜎𝐶!,!(𝑇/𝑊)

For this analysis, the thrust-to-weight ratio is assumed to be proportional to the horsepower to weight ratio. Then, the takeoff ground run distance and takeoff field length can be plotted in terms of (!/!)(!/!)

!

and the takeoff ground run distance, respectively:

Figure 60. Takeoff Ground Run Distance vs. Takeoff Parameter [51]

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Figure 61. Takeoff Field Length vs. Takeoff Ground Run [51]

3. Climb Performance

a) Derivations

The rate of climb (ℎ) can be expressed in terms of power available, power required for level (unaccelerated) flight, and aircraft weight. To simplify the equations, the climb parameter 𝑃 is used as follows:

𝑃 =𝜂𝑃𝑊

−𝑊/𝑆

19(𝐶!!/!/𝐶!) 𝜎

ℎ = 33,000𝑃

An analysis of the aircraft in climb performance yields that the rate of climb is a maximum when the aircraft is flown at a speed such that 𝐶!!/!/𝐶! is at maximum (equation 11). At 𝐶!!/!/𝐶!, the induced drag coefficient is three times the zero-lift drag coefficient. However, for (𝐶!!/!/𝐶!)!"#, the induced drag is equal to the zero-lift drag. The lift coefficient at (𝐶!!/!/𝐶!)!"# (𝐶!,!) is expressed by equation 10:

𝐶!,! =3𝐶!,!𝐾

= 3𝐶!,!𝜋𝐴𝑅𝜀

The total drag coefficient is about four times the zero-lift drag coefficient (𝐶! = 4𝐶!,!), and so the induced drag accounts for ¾ of the total drag (𝐶! =

!!𝐶!,!). Then, (𝐶!!/!/𝐶!)!"# is obtained as shown:

(𝐶!!/!/𝐶!)!"# =1.345 𝐴𝑅𝜀 !/!

𝐶!,!!/!

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It is possible to plot the 𝐶!!/!/𝐶! in terms of the zero-lift drag coefficient for different aspect ratios. To demonstrate this relationship, the author assumed an efficiency factor of 0.7 and generated the following plot:

Figure 62. 𝑪𝑳

𝟑/𝟐/𝑪𝑫 Vs Zero-lift Drag Coefficient [51]

The following relations allow finding the 𝐶!!/!/𝐶! at any lift coefficient if (𝐶!!/!/𝐶!)!"# and 𝐶!,! are know:

𝐶!! =

𝐶!𝐶!,!

(𝐶!!/!/𝐶!)(𝐶!!/!/𝐶!)!"#

=4𝐶!

!/!

1 + 3𝐶!!

b) FAR Climb Criteria For aircraft less than 6000 lb gross weight, all engines operating:

• The rate of climb in the takeoff configuration (with flaps) must be at least 10VS. • The rate of climb in the landing configuration must be at least 5VS.

It is important to note that rates of climb are given in ft/min and speeds in mph. The required rate of climb is

expressed in terms of the stalling speed and the factor K1 as:

ℎ = 𝐾!𝑉!

For takeoff K1=10 and for landing K1=5. The required ratio of horsepower to weight to provide such climb criteria can be expressed as functions, (𝐶!!/!/𝐶!)!"#, the propulsive efficiency, and the wing loading parameter:

𝑃𝑊=

𝑊/𝑆𝜎

19.73𝐾!𝐶!,!"#

+ 1736.8(𝐶!!/!/𝐶!)!"#

33,000𝜂

This relation is showed in the following figure:

SENIOR DESIGN:




Figure 63. Power-to-Weight Ratio vs. Wing Loading [51]

4. Matching Procedure

The previous methods are aimed to determine a relationship between wing loading and power loading that allows the aircraft to meet with certain performance objectives. The following figure illustrates the matching chart that can be built using the procedure outlined in the previous sections.

Figure 64. Example of Matching Chart for Propeller-Driven Aircraft [51]

The intersection of the takeoff field length and maximum speed lines (match point), represents the unique wing/power loading combination required to meet the desired speed and takeoff field objectives. For this example, the following objectives were chosen:

• Vmax = 200 mph at sea level. • VS = 65 mph, with field length of 1640 ft.

SENIOR DESIGN:




• LT = 1550 ft over a 50 ft obstacle. • Climb rate = 1100 ft/min.

Then, the chart is build using the following:

• Maximum Speed Line: o Inputs: power index Ip, and maximum speed Vmax. o See: figure 19.

• Stalling Speed Line: o Inputs: Stall speed requirement VS and Maximum Lift Coefficient CL,max. o See: figure 21.

• Landing Field Line: o Inputs: landing field requirement LT. o See: figures 22 & 23.

• Takeoff Field Line: o Inputs: takeoff lift coefficient CL,T. o See: figures 24 & 25.

• Climb rate: o See: figure 27.

The previous procedure can be further simplified if the maximum speed and stalling speeds are the only specified performance objectives.

• Given the maximum speed (Vmax), use figure 19 to find the power index (Ip). • Then, use figure 21 to find the required wing loading (W/S) for such power index and CL,max. • The power loading (W/P) can be estimated by dividing the wing loading by the cube of the

power index.

5. Weight, Range and Fuel Fraction

Based on the relationship between power loading and aircraft’s physical parameters, Loftin performs an analysis of weight estimation techniques for conventional propeller-driven aircraft. Based on historical correlations, it is found that the sum of fuel weight (Wf), propulsion systems weight (Wt), and payload weight (Wp) are about 60% of the total gross weight (Wg). Then, the parameter [1-(We/Wg)], defined as the “useful load fraction”, is related to the power loading as follows:

Figure 65. Useful Load Fraction vs. Power Loading for Propeller Driven Aircraft [51]

SENIOR DESIGN:




1 −𝑊!𝑊!

= 𝑈 =𝑊! +𝑊!

𝑊!

𝑊!𝑊!

= 𝑈 −𝑊!

𝑊!

Then, the gross weight can be determined if the fuel fraction weight is known as follows:

𝑊! =𝑊!

𝑈 − (𝑊!/𝑊!)

The fuel fraction weight is found by analyzing the range requirement specified by the mission requirements. Such calculation depends, then, on the type of propulsion and fuel used. For conventional fuel-driven propulsion systems, range is determined using the Breguet range equation:

𝑅 = 375 𝜂(𝐿/𝐷)

𝑐log!

11 − (𝑊!/𝑊!)

Then, the fuel fraction weight can be solved as: 𝑊!

𝑊!= 1 −

1𝑒!/!

For other types of propulsion system, the fuel fraction weight equation must be properly modified. For instance, if batteries are used, the mass fraction of the batteries is expressed in terms of the energy density (E, in Wh/kg), the range, the propulsive efficiency, the gravity, and the lift to drag ratio [52]:

𝑚!

𝑚!=


Once the gross weight is determined, what is left is to find the size and power using the wing loading (W/S) and power loading (W/P). This concludes the sizing process of the aircraft, which creates a solution space that can be used to start the design process by the disciplines. It is really important, though, that this process is adapted to the changes that flying in Mars imposes. In addition, Loftin does not have data for low wing loading aircraft, and the results are obtained through an interpolation of the data. This increases the uncertainty on the sizing process, which might yield results that will not be feasible. Finally, if some sort of aid is used for landing and takeoff, the airport performance requirements might not be longer useful.

A MATLAB script has been written to apply Loftin’s method. Such script can be found on section D of this report.

SENIOR DESIGN:




B. Databases Table 26. Summary of Missions to Mars

Spacecraft Year Country Mission CarrierRocket MissionOutcome References Objectives/OnboardEquipment LaunchMass(kg) Payload(Kg) Shape Geometry SolarPanels? Power(MarsDistance)Marsnik1 1960 SovietUnion Flyby Molniya FailedtoOrbit 2,3,6 650 10 Cylindrical L:2,D:1.05 YesMarsnik2 1960 SovietUnion Flyby Molniya FailedtoOrbit 4,5,6 640 10 Cylindrical L:2,D:1.05 YesSputnik22 1962 SovietUnion Flyby Molniya DestroyedinLEO 6,7 893.5 Yes

Sputnik23(Mars1) 1962 SovietUnion Flyby MolniyaSpacecraftfailure(Itdid

reachedMars,butcommunicationswerelost)

8Imagethesurface,Mars'

magneticfield,atmosphericstructure\

893.5 Cylindrical L:3.3,D:1.0 Yes

Sputnik24(Mars2) 1962 SovietUnion Lander Molniya LaunchFailure 9 890 Yes

Mariner3 1964 USA Flyby AtlasAgenaD

ProtectiveShieldfailedtoeject;extraweight

preventedittofollowitstrajectory

10,11 Measuresolarwinds,radiation,cosmicrays/dust,TVcamera

260.8 OctagonalFrame H:0.457,Diag:1.27 Yes 300

Mariner4 1964 USA Flyby AtlasAgenaD Returned21Images 10,13 SameasMariner3 260.8 OctagonalFrame H:0.457,Diag:1.27 Yes 300

Zond2 1964 SovietUnion Flyby(Landing?) Molniya Communicationslostbeforeflyby

12,14 Phototelevisioncamera,ultravioletspectometres

890 CylindricalwithSystemsattached Yes

Zond3 1965 SovietUnion Flyby Molniya

LaunchwindowtoMarswaslost,itwasstill

launchedasaspacecrafttest

15 TVCamera,andotheratmosphericsensors

960 CylindricalwithSystemsattached Yes

Mariner6 1969 USA Flyby AtlasCentaur Transmitted75photos 6,16StudyMars'surfaceand

atmosphere,providedataforMariner7

411.8 57.6 OctagonalFrame H:0.457,Diag:1.387 Yes 449

Mariner7 1969 USA Flyby AtlasCentaur Transmitted126photos 6,17 StudyMars'surfaceandatmosphere

411.8 57.6 OctagonalFrame H:0.457,Diag:1.387 Yes 449

Mars1969A 1969 SovietUnion Orbiter Proton-K/D(UR500) FailedtoOrbit 18 TVCamera,andotheratmosphericsensors

4850 Yes

Mars1969B 1969 SovietUnion Orbiter Proton-K/D(UR500) FailedtoOrbit 18 TVCamera,andotheratmosphericsensors

4850 Yes

Mariner8 1971 USA Orbiter AtlasCentaur FailedtoOrbit 19 Takeatmosphericreadings 997.9 63.1 OctagonalFrame H:0.457,Diag:1.387 Yes 500Cosmos419 1971 SovietUnion Orbiter Proton-K/D(UR500) LaunchFailure 20 InterplanetaryProbe 4650 Cylindrical 3 Yes

Mars2 1971 SovietUnion Orbiter Proton-K/D(UR500) PartialSuccess,landerfailed

21 Firstman-madeobjecttoreachMars'surface

2265 1210(lander) Cylindrical H:4.1,D:2,5.9withSolarPanels Yes

Mars3 1971 SovietUnion Orbiter Proton-K/D(UR500)PartialSuccess,

communicationwithlanderwaslost

21Studytopographyofthesurface

andotheratmosphericmeasurements

2265 1210(lander) Cylindrical H:4.1,D:2,5.9withSolarPanels Yes

Mariner9 1971 USA Orbiter AtlasCentaur Orbitedfor349days 22MappingMartiansurfaceandstudychangesinMartian

atmosphere558.8 63.1 OctagonalFrame H:0.457,Diag:1.387 Yes 500

Mars4 1973 SovietUnion Orbiter Proton-K/D(UR500) UnabletoenterMarsorbit 23 Camerasandatmosphericdevices

2270 Cylindrical Yes

Mars5 1973 SovietUnion Orbiter Proton-K/D(UR500) FailedtodepartEarthOrbit 24Atmosphericdevicesand

communicationlinkswithlaterdevices

2270 Cylindrical Yes

Mars6 1973 SovietUnion Lander/Flyby Proton-K/D(UR500) FailedwhiledescendingtoMartianatmosphere

25 Atmosphericdevices 635(lander) Cylindrical Yes

Mars7 1973 SovietUnion Lander/Flyby Proton-K/D(UR500)Landingprobeseparatedprematurelyandmissed

Mars26 Atmosphericdevices 636(lander) Cylindrical Yes

Viking1 1975 USA Orbiter/Lander TitanIIE-CentaurSuccess,mostcomplete

atmosphericcharacterizationtodate

27,28Obtainhighdefinitionimages,sampletheatmosphere,andsearchforevidenceoflife

883(Orbiter)and572(Lander)72(orbiter)and91(lander) Yes 620(Orbiter)and70(Lander)

Viking2 1975 USA Orbiter/Lander TitanIIE-Centaur Successful 27,29Obtainhighdefinitionimages,sampletheatmosphere,andsearchforevidenceoflife

883(Orbiter)and572(Lander)72(orbiter)and91(lander) Yes 620(Orbiter)and70(Lander)

Phobos1 1988 SovietUnion Orbiter/Lander Proton-K/D(UR500) TerminalfailureontheroutetoMars

30,31,32 Atmosphericandinterplanetarymeasurementdevices

6220(Total),2600(Lander) Yes

Phobos2 1988 SovietUnion Orbiter/Lander Proton-K/D(UR500)Partialsuccess,

communicationlostwithlander

30,31,32Atmosphericandinterplanetary

measurementdevices 6220(Total),2600(Lander) Yes

MarsObserver 1992 USA Orbiter CommericalTitanIIIPartialfailure,datawas

collecteduptoorbit,wherecommunicationswerelost

33

Topographicandatmosphericmeasurements,determinequantityandavailabilityof

resources

1018 Corewithtwolargebooms Core:1.1x2.2x1.6,6mbooms Yes 1147

MarsGlobalSurveyor 1996 USA Orbiter DeltaII7925 Success 34 SameasMarsObserver 1030.5 Corewithtwolargebooms Core:1.17x1.17x1.7,2mbooms Yes 980

Mars96 1996 Russia Orbiter/Penetrator Proton-K/D(UR500) Failuretolaunch 35 Studypastandpresentphysical/chemicalprocess

3159 Spacecraftwithtwoindependentstations Yes

MarsPathfinder 1996 USA Lander DeltaII7925 Successful 36Technologydemonstratorfor

low-costlandingsandexplorationofMartiansurface

264(lander) 10(rover) Stationwithsmallrover Yes

Nozomi(PlanetB) 1998 Japan Orbiter M-V(M5) DidnotreachMars-electricalfailures

37 Studytheatmosphere,focusonionosphere

258 33 Prism H:0.58,1.6Square,1mbooms Yes

MarsClimateOrbiter 1998 USA Orbiter DeltaII7425 Burnedintheatmosphere 38 Studytheweather;datarelaysatellite

338 Box,SolarCell,andAntenna H:2.1,W:1.6,D:2 Yes 500

MarsPolarLander 1999 USA Lander DeltaII7425 Failedtoland;communicationwaslost

39 Studythepoles 290 H:1.06,W:3.6 Yes 200

DeepSpace2 1999 USA Penetrator DeltaII7425 Contactwaslost 40Testforthepresenceofwater,studysubsurfacematerial,technologydemonstrator

3.57 Aeroshell,aftbodyandforebodyAft:H:0.1053,D:0.136,Fore:H:0.1056,D:0.035 No

2001MarsOdyssey 2001 USA Orbiter DeltaII7425 Operational 41

Detailedmineralogicalanalysisoftheplanet'ssurface;studyiftheplanetwaseversuitabletoholdlife;studyradiationand

otherenvironmentalhazardsforfutureastronauts

376.3 Box-shapedCorewithsolarpanelsandotherinstrumentsCore:2.2x1.7x2.6 Yes

MarsExpress 2003 ESA Orbiter/Lander Soyuz-FG/Fregat Operational 42HRPhotogeology,mineralogical

mapping,andatmosphericcomposition

666 113 CubedspacecraftwithtwosolarpanelsCore:1.5x1.4x1.68,12mtiptotip Yes 460

Spirit(MER-A) 2003 USA Rover DeltaII7925

Successful;becametrappedandsolarpanelscouldnolongerproduceenough

energy

43Determineiflifeeveraroseon

Mars,prepareforhumanexplorationinMars

185 Rover H:1.5,W:2.3,L:1.6 Yes 140

Opportunity(MER-B) 2003 USA Rover DeltaII7925H Operational 44 SamemissionasSpirit 185 Rover H:1.5,W:2.3,L:1.6 Yes 140

MarsReconnaisanceOrbiter 2005 USA Orbiter AtlasV401 Operational 45 Characterizeclimate,characterizeterrain

1031 Mainbus,twosolarpanels,andantenna Yes 1000

Phoenix 2007 USA Lander DeltaII7925Completedallexperiments,

butsystemranoutofenergyduringMars'winter

46 Searchforenvironmentssuitableforlife

350 Circulardeckwithsolarpanels2.2tallD:1.5(deck),5.5long(solarpanels) Yes

Phobos-Grunt 2011 Russia Orbiter Zenit-2M Spacecraftfailure(didnotleaveLEO)

47,48

CollectandreturnasampleofPhobos(oneofMars'moons);reconstructthehistoryof

Phobos

730 350 Yes 150

Yinghuo-1 2011 China Orbiter Zenit-2M LostwithPhobos-Grunt 49 Studytheatmosphere;surfacemapping

115 Box-shapedwithtwosolarpanelwings 0.75x0.75x0.65,6.85mtiptotip Yes 90

MarsScienceLaboratory(Curiosity)

2011 USA Rover AtlasV Operational 50 Exploretheenvironmentforlife 750 Six-wheeledvehicle L:2.8, Yes

Mangalyaan 2013 India Rover PSLV-XL Operational 51 ExploreMars'surfacefeatures 488 15 Cubeandsolararrays Side:1.5 Yes 800

MAVEN 2013 USA Orbiter AtlasV401 Operational 52,56 ExploreMars'upperatmosphere Cubicprimarycentralstructurewithtwosolarpanelswings2.3x2.3x2.0 Yes

ExoMars 2016 ESA Orbiter/Lander Proton-M/Briz-M Enroute 53,54,57 Technologydemonstrator,geological/chemicalsamples

4332(launch),600(lander) 112 Box-shaped 3.2x2x2,17.5tiptotip Yes

InSight 2018 USA Rover - Indevelopment 55 StudythegeologicalevolutionofMars

10's

60's

70's

80's

90's

00's

SENIOR DESIGN:




Table 27. Light Aircraft Database

NameGross

Weight (lb)

Empty Weight

(lb)

Engine Power

(hp)

Wing Area (ft^2)

Vmax (mph)

Vcruise (mph)

Vstall (mph)

Wing Loading (lb/ft^2)

Power Loading (lb/hp)

Useful Load

Fraction (-)

Power Index (-)

# of Propellers Conf. Engine

Wing Position Braced (Y/N)Landing Gear

550 Ultra 3Xtrim 1212 717 98.6 133.5 136 106 44 9.078651685 12.29209 0.408416 0.903924 1 Tractor Rotax 912 High Y FixedACA Champ 1230 925 100 165 108 94 46 7.454545455 12.3 0.247967 0.846263 1 Tractor Continental O-200-D High Y Fixed

ACA Citabria Aurora 1750 1120 118 165 120 115 52 10.60606061 14.83051 0.36 0.894265 1 Tractor Lycoming O-235-K2C High Y FixedACA Citabria Adventure 1750 1200 160 165 140 135 52 10.60606061 10.9375 0.314286 0.989795 1 Tractor Lycoming O-320-B2B High Y FixedACA Citabria Explorer 1800 1250 160 171.86 132 128 52 10.47364134 11.25 0.305556 0.976446 1 Tractor Lycoming O-320-B2B High Y Fixed

ACA Denali Scout 2150 1340 210 180 150 136 36 11.94444444 10.2381 0.376744 1.052727 1 Tractor Lycoming IO-390-A1B6 High Y Fixed

ACA High Country Explorer 1950 1300 180 171.86 142 132 49 11.34644478 10.83333 0.333333 1.015545 1 Tractor Lycoming O-360-C4P High Y FixedACA Scout 2150 1320 180 180 140 130 43 11.94444444 11.94444 0.386047 1 1 Tractor Lycoming O-360-C1G High Y Fixed

ACA Super Decathlon 1950 1305 180 169 155 141 57 11.53846154 10.83333 0.330769 1.021242 1 Tractor Lycoming AEIO-360-H1B High Y Fixed

ACA Xtreme Decathlon 1950 1330 210 164 161 149 58 11.8902439 9.285714 0.317949 1.085905 1 Tractor Lycoming AEIO-390-A1B6

High Y Fixed

Aerotrek A220-A240 1235 655 100 122.53 143 120 40 10.07916429 12.35 0.469636 0.934514 1 Tractor Rotax 912 ULS High Y FixedAero-Works Aerolite 103** 600 275 40 124 70 63 35 4.838709677 15 0.541667 0.685824 1 Tractor Rotax 447 High Y FixedAirdrome Dream Classic* 393 223 40 86 76 68 31 4.569767442 9.825 0.43257 0.774796 1 Tractor Rotax 447 High Y Fixed

AMD Zodiac CH601 XL 1320 690 100 132 161 130 44 10 13.2 0.477273 0.911609 1 Tractor Rotax 912 S Low N Fixed Ameri-Cana Eureka 460 230 25 104 63 55 27 4.423076923 18.4 0.5 0.621778 1 Tractor Hirth F-33 Low Y Fixed

Avid Champion 594 254 28 114.5 65 63 26 5.187772926 21.21429 0.572391 0.625342 1 Tractor Rotax 277 High Y FixedB&F Fk14 Polaris 1042 626 98.5 101.2 160 151 40 10.29644269 10.57868 0.399232 0.991026 1 Tractor Rotax 912 ULS Low N Fixed

Beaujon Enduro** 436 230 16 102 65 55 26 4.274509804 27.25 0.472477 0.539312 1 Pusher Briggs & Stratton 401417

High Y Fixed

Birdman Chinook WT-11** 625 250 28 140 60 50 24 4.464285714 22.32143 0.6 0.584804 1 Pusher Rotax 277 High Y FixedBOT SC07 Speed Cruiser 1042 642 100 108 144 134 40 9.648148148 10.42 0.383877 0.974673 1 Tractor Rotax 912 ULS High N Fixed

Cessna 162 Skycatcher 1320 843 100 120 136 125 50.6 11 13.2 0.361364 0.941036 1 Tractor Continental O-200-D Flat 4

High Y Fixed

CGS Hawk (Classic)** 600 310 40 135 80 75 35 4.444444444 15 0.483333 0.666667 1 Tractor Rotax 447 High Y FixedChotia Gypsy** 500 250 26 144 55 35 22 3.472222222 19.23077 0.5 0.565202 1 Tractor Rotax 277 High Y Fixed

CZAW Mermaid*** 1235 838 120 124 130 115 40 9.959677419 10.29167 0.321457 0.98913 1 Pusher Jabiru 3300 Middle N RetractableCZAW PS-28 Cruiser 1320 855.4 100 132.4 137 107 36 9.96978852 13.2 0.35197 0.91069 1 Tractor Rotax 912 ULS2 Low N Fixed

Denney Kitfox Classic IV 1200 650 80 123 117 110 37 9.756097561 15 0.458333 0.86642 1 Tractor Rotax 912 High Y FixedDynAero Pick-Up 1040 682 100 89.3 162 152 42 11.64613662 10.4 0.344231 1.038443 1 Tractor Rotax 912 ULS Low N Fixed

DynAero ULC 1040 616 80 87.5 198 155 50 11.88571429 13 0.407692 0.970571 1 Tractor Rotax 912 UL Low N FixedEuropa XS 1370 780 99 102 180 155 60 13.43137255 13.83838 0.430657 0.990098 1 Tractor Rotax 912 ULS Low N Fixed

Europa XS Motorglider* 1370 883 80 141.14 140 126 52 9.706674224 17.125 0.355474 0.827586 1 Tractor Rotax 912 UL Middle N RetractableFantasy Air Allegro 2007 1321 606 80 123 130 109 40 10.7398374 16.5125 0.541257 0.86642 1 Tractor Rotax 912 UL High Y Fixed

Fisher Avenger 600 280 50 121 95 80 28 4.958677686 12 0.533333 0.744838 1 Tractor Rotax 503 Low Y FixedFisher FP-202 Koala 830 400 64 140 80 75 32 5.928571429 12.96875 0.518072 0.770343 1 Tractor Rotax 582 High Y Fixed

Fisher FP-606 Sky Baby 500 250 25 116 65 60 26 4.310344828 20 0.5 0.599553 1 Tractor Hirth F-33 High Y FixedFreebird II 970 385 50 132 85 70 32 7.348484848 19.4 0.603093 0.723545 1 Pusher Rotax 503 High Y Fixed

Grob G 109B Vigilant T1* 2000 1364 95 204.5 130 - 38.3 9.7799511 21.05263 0.318 0.77448 1 Tractor Grob 2500E1 Low N FixedIkarus C42 1041 583 100 134.5 120 108 39 7.739776952 10.41 0.439962 0.905926 1 Tractor Rotax 912 ULS High Y Fixed

InterPlane Skyboy S 1000 490 64 163.6 90 68 40 6.112469438 15.625 0.51 0.731362 1 Pusher Rotax 582 High Y FixedISON Airbike** 560 257 40 118 80 63 30 4.745762712 14 0.541071 0.697257 1 Tractor Rotax 447 High Y Fixed

Jabiru J120 1102 627 80 85 150 115 59 12.96470588 13.775 0.431034 0.979995 1 Tractor Jabiru 2200 High Y FixedJabiru J160 1188 704 80 86.5 150 115 61 13.73410405 14.85 0.407407 0.974297 1 Tractor Jabiru 2200 High Y FixedJabiru J170 1323 705 80 102.9 140 115 45 12.85714286 16.5375 0.46712 0.919514 1 Tractor Jabiru 2200 High Y FixedJabiru J230 1320 816 120 103 150 138 52 12.81553398 11 0.381818 1.05224 1 Tractor Jabiru 3300 High Y Fixed

Lockwood Super Drifter** 1000 495 80 160 80 75 34 6.25 12.5 0.505 0.793701 1 Tractor Rotax 912 UL High Y FixedParadise P-1 LSA 1323 750 100 136 137 115 36 9.727941176 13.23 0.433107 0.902583 1 Tractor Rotax 912 ULS High Y Fixed

Pipistrel Sinus 912 1200 626 80 132 138 126.5 39.1 9.090909091 15 0.478333 0.846263 1 Tractor Rotax 912 UL2 High N FixedPipistrel Alpha Trainer 1212 615 80 96.6 138 124.2 49.5 12.54658385 15.15 0.492574 0.939084 1 Tractor Rotax 912 UL High N Fixed

Pipistrel Virus UL 1039.5 625 80 118.4 149.1 140 44.7 8.779560811 12.99375 0.398749 0.877498 1 Tractor Rotax 912 UL High N FixedPipistrel SW 80 1039.5 631.4 80 102.3 164 153 49 10.16129032 12.99375 0.392593 0.921308 1 Tractor Rotax 912 UL High N FixedPipistrel SW 100 1039.5 635.8 100 102.3 179 170 49 10.16129032 10.395 0.38836 0.992449 1 Tractor Rotax 912 ULS High N Fixed

Pipistrel Apis Bee* 709.5 484 28 131.75 136.6 89.4 36.7 5.385199241 25.33929 0.317829 0.596764 1 Tractor HIRTH F33 BS Middle N Fixed

Pipistrel Taurus M* 990 627 53 132.7 139.75 101.2 44.1 7.460437076 18.67925 0.366667 0.736436 1 Tractor High Performance Engine

Middle N Fixed

Quad City Challenger II 960 460 50 177 96 85 28 5.423728814 19.2 0.520833 0.656144 1 Pusher Rotax 503 High Y FixedRainbow Cheetah XLS 1235 547 80 142.6 110 84 35 8.66058906 15.4375 0.557085 0.824752 1 Tractor Rotax 912 UL High Y FixedRans S-19 Venterra 1320 820 100 126.9 140 136 50 10.40189125 13.2 0.378788 0.923661 1 Tractor Rotax 912 ULS Low N Fixed

Remos GX 1320 705 100 118 130 123 44 11.18644068 13.2 0.465909 0.946323 1 Tractor Rotax 912 ULS High Y FixedSkyleader 100 695 463 50 94.94 120 87 34 7.320412892 13.9 0.333813 0.807558 1 Tractor Rotax 503 Low N FixedSkyleader 200 992 657 80 127.5 150 137 44 7.780392157 12.4 0.337702 0.856104 1 Tractor Rotax 912 UL Low N Fixed

Sling 2 LSA 1320 815 100 127.5 140 123 51 10.35294118 13.2 0.382576 0.92221 1 Tractor Rotax 912 ULS Low N FixedSpectrum Beaver SS** 650 340 40 138 85 67 30 4.710144928 16.25 0.476923 0.6618 1 Pusher Rotax 447 High Y Fixed

Sport Hornet LRS 1320 680 100 155 138 115 40 8.516129032 13.2 0.484848 0.864084 1 Pusher Rotax 912 High Y FixedStorm Rally 990 640 100 125 140 - 43 7.92 9.9 0.353535 0.928318 1 Tractor Rotax 912 High Y Fixed

Team Mini-Max Hi-MAX 560 328 40 112.5 80 70 31 4.977777778 14 0.414286 0.708439 1 Tractor Rotax 447 High Y FixedUFM-13 Lambada 1042 628 101 131 124 93 40.25 7.954198473 10.31683 0.397313 0.916959 1 Tractor Rotax 912 ULS Middle N Fixed

Van's RV-12 1320 740 100 127 135 131 47 10.39370079 13.2 0.439394 0.923419 1 Tractor Rotax 912 ULS Low N FixedWhisper Motorglider* 1705 1166 75 160 130 110 44 10.65625 22.73333 0.316129 0.776808 1 Tractor Limbach L2000 Low N Fixed

Zenith STOL CH 701 1100 580 80 122 85 80 28 9.016393443 13.75 0.472727 0.86878 1 Tractor Rotax 912 High Y Fixed* Gliders, ** Ultralights. *** Amphibius

Propulsion System Aircraft ClassPerformanceGeneral Parameters Loftin Parameters

SENIOR DESIGN:




C. Disciplines Relevant Figures

1. Aerodynamics

Figure 66. Lift Characteristics Comparison for S1223 and FX 63-137 Airfoils [4]

Figure 67. Drag Characteristics Comparison for S1223 and FX 63-137 Airfoils [4]

Figure 68. Aircraft Model Examined Under Computational Flow Simulations [4]

SENIOR DESIGN:




2. Propulsion and Energy

Figure 69. Propeller Design and Outputs [84]

Figure 70. Tradeoff Between Efficiency and Power [84]

Diameter,m 4.25NumberofBlades 4

RPM 800.00

Thrust,N 511.25Power,kW 59.34Efficiency 0.75OverallEfficiency 0.68

LocationofEnginesInlinewithVerticalTail

PropellerDesignAndOutputs

SENIOR DESIGN:




D. Mars Aircraft Concept Information Tables

Related WorksV. Clarke, A. Kerem, and R. Lewis, "A Mars Airplane…Oh Really,?" AIAA Aerospace Sciences Meeting, New Orleans, LA, 1979.

A. Colozza, "Comparison of Mars Aircraft Propulsion Systems," NASA CR-2003-212350.A. Colozza, "APEX 3D Propeller Test Preliminary Design," NASA CR-2002-211866.

b) Contains detailed propulsion systems diagrams

c) Calculation of power requirements for each system and

component.

Synthesis

b) General design trades and measures of merit when designing a

Mars aircraft

a) Description of HAPP design method and its adaptation

toaccount for Mars environmental conditions.

ApplicabilityPropulsion and Energy

a) Comparison of different energy systems

Structures and Weightsa) Detailed component breakdown

for each designb) The appendix contains equations

used for such weight estimations

Summary: A feasibility study for a Mars perpetual endurance aircraft concept. It considers two energy production systems: radioisotope/heat engines and solar cells. The design is driven by the goal of achieving the maximum endurance while keeping the minimum wingspan possible. The study concludes that it is indeed possible to have such aircraft, although the required wingspan for solar powered aircraft is above 100 m and around 40 m when using radioisotope powered aircraft. The study also analyzes different solar cells, assessing the effect of solar cell efficiency on the final design. Similarly, it compares using Pu 238 and Cm 244 heat sources. The study contains detailed descriptions of wing parameters (wing loading, aspect ratio, and others), weight breakdowns (for propulsion system, airframe, etc.) and basic aerodynamics/performance parameters (CL, CD, L/D, Endurance parameter, and speed). Finally, the study concludes with recommendations for future Mars aircraft design projects. 1) A definitive flight plan should be designated in order to model the environment in which the aircraft is expected to fly. 2) To increase the collected solar energy, movable or variable geometry wings should be considered. 3) The radiator system for both, radioisotope and solar powered aircraft, must be fully described in order to obtain accurate figures of merit for the system. 4) A deployment scheme for the aircraft should be devised. This scheme depends mostly on the mission type.

Mars Aircraft Concept OverviewAuthor(s): Colozza, Anthony Organization: NASA Publication Date: April 1990

Reference: A. Colozza, "Preliminary Design of a Long-Endurance Mars Aircraft," NASA CR 185243, 1990

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Aerodynamics

b) Lift and Drag profiles for multiple Reynolds numbers

Given the variety of scientific payloads, the mission scenario calls for 12 equal airplanes to be deployed into Martian atmosphere. A single spacecraft would carry up to four airplanes, each one of them folded within an aeroshell for atmospheric entry. Then, the airplanes would unfold and deploy while the aeroshell is still descending into the planet. Using up to 6 wing breaks, a 21-m wingspan - 20 m2 wing area could be folded to fit a 3.8 m diameter aeroshell.

ApplicabilitySynthesis Structures and Weights

a) Description of the aerodynamic challenges associated when flying in Mars

b) Analysis of loads during deployment, landing, and takeoff maneuvers

b) Calculation of power requirements for each system and component.

a) Large payload + VTOL: the design can be used as baseline for this project

a) Description of folding mechanism to fit within aeroshell contraints

b) Description of scientific missions and payloadsc) Mission architecture design.

Similar WorksA. Colozza, "Preliminary Design of a Long-Endurance Mars Aircraft," NASA CR-1990-185243.M. Guynn, et all, "Evolution of a Mars Airplane Concept for the ARES Mars Scout Mission," AIAA Unmanned Unlimited Systems, Technologies, and Operations, San Diego, CA, 2003.

Mars Aircraft Concept Overview

The airplane employs six hydrazine rockets to takeoff and land, whereas a propeller is used to cruise. Such rockets are a variation of the engine used on NASA’s Mini-Sniffer. The paper contains a detailed description of the landing and takeoff maneuvers. For the cruiser variation, electric propulsion was discarded due to the low energy density (100-150 Wh/lb) of the batteries available at the time. However, if the energy density was to be doubled, the range of the vehicle would be 10-30% higher than the hydrazine powered cruiser.

Finally, the paper describes the avionics equipment necessary for such missions. It also analyzes the use of the spacecraft as a comsat for the airplanes. Then, it presents a weight breakdown for each configuration (Lander and Cruiser), as well as a comparison between the hydrazine engine and an electric engine. All variations were designed for a maximum gross weight of 300 kg.

Propulsion and Energya) Comparison of different energy systems: hydrazine and electric

Author(s): Clarke, Kerem, and Lewis Organization: AIAA Publication Date: Jan 1979Reference: V. Clarke, A. Kerem, and R. Lewis, "A Mars Airplane…Oh Really,?" AIAA Aerospace Sciences Meeting, New Orleans, LA, 1979.Summary: The paper is a summary of a study conducted by Developmental Sciences Inc under NASA contract report 157942. Such report is the first and most complete study on the mission design, scientific use, and conceptual design of a Mars Airplane. The idea is that an airplane could be used as a versatile tool to conduct multiple missions such as delivery of science packages, collecting samples, and as an aerial survey platform. The authors designed to versions of the aircraft: a cruiser and one with landing/takeoff capability. Its main advantage is that it can reach places that other platforms (e.g. Rovers) cannot, as well as it provides a better resolution than orbiters. In addition, it has great mobility, as it cruises at speeds between 60-100 m/s, with a maximum range of 10000 km, and flight time between 17–31 hours, and a payload capability of 40 to 100 kg. In terms of design, it is similar to a conventional air glider, with an inverted (aft) v-tail.

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b)Folding and aeroshell fitting

Geometrya) Wing geometry tradeoffs: wing sweep vs. wing position, wing area vs. aspect ratio

Stability and Controla) Tail arrangement (flying wing, canard, or tail aft) and shapeb) c.g. location and stabilityc) Control surfaces layout

Structures and Weightsa) Detailed deployment and folding mechanismb) Preliminary weight estimationc) Fuselage shape and integration

Similar WorksSmith, et all, "Mars Airplane Airfoil Design With Application to ARES," AIAA Unmanned Unlimited Systems, Technologies, and Operations, San Diego, CA, 2003.J. Lede, R. Parks, M. Croom, "High Altitude Drop Testing in Mars Relevant Conditions for the ARES Mars Scout Mission," AIAA Unmanned Unlimited Systems, Technologies, and Operations, San Diego, CA, 2003.S. Kenney, M. Croom, "Simulating the ARES Aircraft in the Mars Environment," AIAA, 2004.

The design philosophy behind ARES is to meet the science requirements; hence, the risk, complexity, and cost of the system should be minimal. Within this context, another design requirement is to successfully transmit the data back to Earth, which was to be accomplished using a spacecraft executing a fly-by maneuver while the aircraft was flying. The aircraft was to crash land once its mission was completed, and no recovery system was conceived.

Synthesis Propulsion and Energy Aerodynamicsa) Analysis of different fuel sources and engines types

b) Airfoil selection and validation

a) Desing philosophy and mission selection

a) Wing twist and airfoil distribution study

b) Methodology for Concept Explorationc) Configuration selection process c) System integration c) Design refinements

b) Rocket vs Propeller propulsion tradeoff

This paper, as the name indicates, shows the design process behind the ARES, starting from the mission selection, the consideration of propulsion systems, and folding/deployment techniques. The latter is one of the design drivers for ARES, as it was intended to be deployed during atmospheric entry from an aeroshell.

The paper goes over the ARES design for each discipline: aircraft configuration, aeroshell packing, propulsion, tail arrangement, aircraft geometry, wing aerodynamics, structures, stability and control. The final design has a wing span of 6.25 m, an aspect ratio of 5.58, a CL range from 0.52-0.71, a L/D of 14-14.4, a T/W~0.1, a range of 500 km, and an endurance of 60 min.

Applicability

Mars Aircraft Concept OverviewAuthor(s): Guynn, Croom, Smith, Parks, Gelhausen Organization: NASA Publication Date: Sept 2003

Reference: M. Guynn, et all, "Evolution of a Mars Airplane Concept for the ARES Mars Scout Mission," AIAA Unmanned Unlimited Systems, Technologies, and Operations, San Diego, CA, 2003.Summary: The Aerial Regional-scale Environmental Survey of Mars (ARES) is an aircraft concept intended to conduct scientific experiments of Mars. This concept has already been tested on Earth and it was considered to be included in one of the latest NASA’s missions to Mars. The ARES missions the most complete and advanced Mars aerial platform designed, and so it is one of the best references for this project. The final design, as shown on the right, is the result of multiple design iterations and data obtained from flight tests.

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Similar WorksJ. Jacob, S. Smith, "Design Limitations of Deployable Wings for Small Low Altitude UAVs," AIAA Aerospace Sciences Meeting, Orlando, FL, 2009.J. Jacob, S. Smith, "Design of HALE Aircraft Using Inflatable Wing," AIAA Aerospace Sciences Meeting and Exhibit, Reno, NV, 2008.

J. Murray, T. Frackowiak, J. Mello, B. Norton, "Ground and Flight Evaluation of a Small-Scale Inflatable-Winged Aircraft," NASA TM-2002-210721.

Summary: This project studies the feasibility of using inflatable wings for planetary exploration, with emphasis on Mars. It summarizes the current state of technology of inflatable wings, including low-altitude flight testing on Earth and computational fluid dynamic simulations of inflatable wing shapes. The main advantage that inflatable wings propose is their ease to be folded and packet into an aeroshell (minimal packed-volume-to-weight ratio), which facilitates space travel and deployment.

J. Jacob, S. Scarborough, "A High-Altitude Test of Inflatable Wings for Low-Density Flight Applications," AIAA, 2006.

a) Detailed computational and wind tunnel analysis of wing aerodynamics for multiple angles of attack and at different Re. b) Deployment techniques and

system integrationc) Internal wing arrangement

a) Pros and Cons of each deployable mechanism

a) Stowed volume considerations and aeroshell fittingb) Preliminary load estimation

techniques for inflatable wings

Three types of wing surfaces were tested: “smooth”, “bumpy”, and a combination of the two. The preliminary analysis of the surfaces was performed using UNCLE, a software developed at the University of Kentucky. Then, the results were confirmed with wind-tunnel tests. It was concluded that at low Reynolds numbers (<100,000), the “bumpy” profile had a better aerodynamic performance, as its surface reduced flow separation. However, for moderate (~250,000) and high (~500,000) Reynolds number, the smooth profile performs better than the bumpy one. As a reference, the paper “Design Limitations of Deployable Wings for Small Low Altitude UAVs” contains an extensive review of all possible deployable wing mechanisms: folding, inflating, nesting, extending, morphing, and hybrid. Moreover, as the name indicates, the paper analyzes the aerodynamic and structural design requirements for each one of the previous options. Other papers have also been published as an effort to increase the knowledge on inflatable wings (see “related works”), and they should all be considered into the discussion.

ApplicabilityStructures and Weights Aerodynamics Synthesis


Author(s): Jacob, Reasor, LeBeau, Smith

Organization: University of Kentucky, Oklahoma State University

Publication Date: Jan 2007

Reference: D. Reasor, S. Smith, R. Lebeau, J. Jacob, "Flight Testing and Simulation of a Mars Aircraft Design Using Inflatable Wigns," AIAA 2007.

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Author(s): Datta, Roget, Griffiths, Pugliese, Sitaraman, Bao, Liu, Gamard

Organization: University of Maryland

Summary: The Martian autonomous rotary wing vehicle (MARV) was designed at the University of Maryland as a response to the NASA/Sikorsky American Helicopter Society student design competition in the year 2000. The study concludes that such vehicle is feasible to be flight in Martian conditions. The paper describes its design, deployment, and mission profile.

Synthesis Aerodynamicsa) Design requirements and considerations for vertical lift devicesb) Comparison of MARV design performance to common rotorcraftc) Deployment techniques for rotorcraft devices

b) Current airfoils were not suitable, a new one had to be

The paper presents a detailed aerodynamic design for the blades of the rotorcraft (planform, airfoil, and structure), as well as the power plant selection. The structure of the MARV is fully explained with technical drawings for the blades, the control system, and the landing gear. The only apparent downside of using rotorcraft devices is on the high power requirements and the limited payload capabilities. On the MARV, ~20% of the gross takeoff weight is used for the power supply system. On RHOVER (see reference below), power train is ~40%, with a payload of roughly 15%.

Similar WorksJ. Fuentes, "Martian RHOVER Feasibility Study," AIAA SciTech, Kissimmee, FL, 2015.A. Datta, et all, "The Martian Autonomous Rotary-Wing Vehicle (MARV)," University of Maryland, College Park, MD, 2000.


Reference: Datta, et all, "Design of a Martian Autonomous Rotary-Wing Vehicle," Journal of Aircraft, Vol. 40, May-June 2003.

The use of vertical lift devices has been conceived as an alternative for planetary exploration due to its maneuverability and precision, thus combining the benefits of a rover and an air vehicle. Thus, multiple teams have decided on rotorcraft (or one of its alternatives) for their projects (see similar works below). The mission requirements for this project were: maximum takeoff mass of 50 kg, autonomous deployment, ~25 km range, ~30 min of controlled flight, ~1 min hover, and optional restart capability. The study concludes that the most appropriate rotorcraft configurations for Mars are: single rotor-tail rotor, quadrotor, and coaxial. The later was chosen because of the large rotor radius and a long tail rotor arm.

L. Young, G. Pisanich, C. Ippolito, "Aerial Explorers," 43rd AIAA Aerospace Sciences Meeting and Exhibit, Reno, NV, 2005.R. Zubrin, "Mars Gashopper," Pioneer Astronautics, Lakewood, CO, 2000

a) Detailed rotor aerodynamic design: airfoil, planform, and structure

ApplicabilityPropulsion and Energy

a) Power requirement calculations for different mission b) Power plant options (electric,

nuclear, and fuel cells)c) Detailed propulsion system

design

Publication Date: June 2003

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A. Colozza, "Comparison of Mars Aircraft Propulsion Systems," NASA CR-2003-212350.

Related WorksM. Guynn, et all, "Evolution of a Mars Airplane Concept for the ARES Mars Scout Mission," AIAA Unmanned Unlimited Systems, Technologies, and Operations, San Diego, CA, 2003.

V. Clarke, A. Kerem, and R. Lewis, "A Mars Airplane…Oh Really,?" AIAA Aerospace Sciences Meeting, New Orleans, LA, 1979.

It was concluded that due to the low atmospheric density and low Re number environment, the resulting wing loadings should be around 1 lb/ft2 (~ 40 N/m2). For the propulsion system, the specific power is on the order of 2x10-2 kW/N. Finally, it was not possible to meet the landing/takeoff requirement of 1000 m, and so some sort of vertical takeoff/landing system has to be implemented. Of the eight resulting designs, two employed canards, two had joined wings, and four used the conventional tail-aft configuration. The gross weights ranged from 4600-7500 N, with wing spans of 37.5-72.0 m, cruise speeds between 60-109 m/s, aspect ratios of 8-25, and wing loadings from 19.1 to 100 N/m2. As for the propulsion system, they had power loadings from 71-222 N/kW, and propeller diameters between 3.0-8.7 m.

a) Analysis of Martian Atmosphere and how it affects conventional aerodynamics

a) Analysis of propeller-driven aircraft and its feasibility for a Mars aircraft

a) Use (and adaptation) of Loftin sizing methodology for multiple cases. Consideration of a manned aircraftb) Identification of critical design points and design trades

b) Consideration of VTOL propulsion systems

ApplicabilityAerodynamics Propulsion Synthesis

c) Interaction of aerodynamics with other disciplines

c) Detailed energy and power calculations

c) Mission analysis (deployment, takeoff, cruise, and landing)

b) Airfoil selection and detailed data for various configurations

Mars Aircraft Concept OverviewAuthor(s): K.R. Sivier and M.F. Lembeck

Organization: University of Illinois at Urbana-Champaign Publication Date: Sept. 1988

Reference: K.R. Sivier, M.F. Lembeck, "The Marsplane Revisited," AIAA-88-4412, Atlanta, GA, 1988Summary: This paper summarizes the senior design project for the spring 1988 UIUC class. The goal to conceptually design a Mars Airplane capable of carrying two persons (plus support systems) and enduring eight hours of flight without refueling. The students used Loftin’s sizing methodology (adapted to Mars gravity and atmosphere) to start off the design process. The critical design areas were identified to be: 1) low atmospheric density, 2) light and efficient propulsion systems, 3) takeoff and landing.

[Figure: Vehicle Illustration]

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Related WorksK.R. Sivier, M.F. Lembeck, "The Marsplane Revisited," AIAA-88-4412, Atlanta, GA, 1988A. Colozza, "Preliminary Design of a Long-Endurance Mars Aircraft," NASA CR 185243, 1990

Finally, the takeoff runway was determined to be around 20,000 ft, with 6,100 ft more required to meet the 50-ft ground clearance requirement. For this project, such length is not an issue, as it assumes the existence of base camps in Mars and clear runway areas.

A propeller was designed and optimized to operate within Mars atmosphere. The resulting design had ten 5.9 ft diameter blades. At 900 RPM, each propeller provided 54 lb at cruise speed, and therefore 3 were required to meet the thrust requirements. The structural analysis concluded that the design could withstand loads of +/- 2g for different maneuvers. The static margin was determined to be around 6%, with acceptable static longitudinal, lateral, and dynamic stability. A performance analysis determined the stall/maximum speeds, climb rates, turn, service ceiling, airport performance, and V-n diagram.

b) Static and Dynamic stability analysis

a) Analysis of required control surface deflection to provide stability

a) Creation of V-n diagram based on multiple factors

ApplicabilityStructures Performance Stability and Control

a) The aircraft configuration was chosen due to structural considerations and a load analysis

Mars Aircraft Concept OverviewAuthor(s): B.R. Morrissette and J.D. DeLaurier

Organization: University of Toronto

Publication Date: March 1999

Reference: B.R. Morrissette, J.D. DeLaurier, "The Zephyr - Manned Martian Aircraft," Canadian Aeronautics and Space Journal, Vol. 45, p. 25-31, 1999Summary: The "Zephyr" is an aircraft designed to, 1) provide quick and versatile manned transportation within Mars, 2) conduct scientific experiments, and 3) participate in search-and-rescue operations. The resulting aircraft is capable of carrying two people (including life support systems) and has a range of 200 nm. The design is a biplane with the two fuselages joined by a horizontal stabilator as shown in the picture. The design was based on powered sailplanes and hydrogen-oxygen fuel cell specifications. The sizing process concluded that the aircraft should weight about 2000 Mars pounds, with a wing area of 2000 ft2. Such structure was both large and light, and therefore an important focus was placed on the structural design. This was the main reason why the biplane was chosen at the expense of the drag.

[Figure: Vehicle Illustration]

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E. Individual Discipline Analysis Diagrams

Figure 71. Aerodynamics IDA

Figure 72. Structures and Weights IDA

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Figure 73. Propulsion and Energy IDA

Figure 74. Performance IDA

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Figure 75. Stability and Control IDA

F. MATLAB Scripts

1. Loftin’s Parametric Sizing Method % Updated: 07/25/16 clear all; close all; % Atmospheric Data g=3.71; rho=0.015; % Mars Gravity [m/s^2] and density at surface [kg/m^3] rho_earth=1.225; % Earth's atmos density [kg/m^3] sigma=rho/rho_earth; % Density Ratio a=250; % Speed of Sound at Mars [m/s] %% Assumed Aircraft Parameters: AR=26; % Wing Fixed Aspect Ratio eps=0.7; % Oswald's Efficiency Factor CLmax=1.8; % Maximum Lift Coefficient CD0=0.0230; % Eq. 6.7b (pg.333), historical for Class I

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K=1/(pi*AR*eps); %(pg.346) prop_eff=0.85; % Propeller Efficiency L2D=24; %% Performance Requirements: % Speed Requirements: Vstall_req=70; % Stall Speed [m/s] Vmax_req=150; % Maximum Speed [m/s] Vstall_mph=Vstall_req/0.44704; % Conversion from m/s to mph Vmax_mph=Vmax_req/0.44704; % Conversion from m/s to mph V=(Vstall_mph-100):1:(Vmax_mph+10); % Creation of speed range for plots % Trajectory Prediction: Range=400000; % Range Requirement [m] Mpayload=250; % Payload Mass [kg] Mfrac_VTOL=0.23; % Avg. Fuel Fraction Required for VTOL (round trip) (Ryan's Code) %% Power Loading Calculations K1=170; Ip=(Vmax_mph+48.38)/201.25; % From Light Aircraft Database Wing_loading=sigma.*CLmax.*(1./19.778.^2).*V.^2; % Range of Wing Loadings [lb/ft^2] Power_loading=Wing_loading./(sigma.*Ip.^3); % (pg.328) W2S_stall_req=(1/g)*0.5*rho*CLmax*Vstall_req^2; % Wing loading in kg/m^2 W2S_stall_req_ips=W2S_stall_req*2.2*0.3048^2; % Wing loading in lb/ft^2 %% Climb Analysis CLc=sqrt(3*CD0*pi*AR*eps); % Lift at maximum value of CL^(3/2)/CD, eq. 6.20 (pg.347) CLm=sqrt(CD0*pi*AR*eps); % Lift at L/D)max, eq. 6.44 (pg.370) Max_climb_parameter=(1.345*(AR*eps)^.75)/(CD0^.25); % Eq. 6.21 (pg.347) K1_climb=10; % Aerodynamic Analysis CDi_cruise=0.1.*CD0./0.9; % (pg.336) CD_cruise=CD0/0.9; % (pg.333) CD_climb=4*CD0; % (pg.348) CDi_climb=(3/4)*CD_climb; % (pg.349) Lift2Drag_max=0.5*sqrt(pi*AR*eps/CD0); % Climb Power2Weight=(7.784.*Wing_loading./(CLmax.*sqrt(sigma))+1736.8.*sqrt(Wing_loading)./(Max_climb_parameter.*sqrt(sigma)))./(33000.*prop_eff); % Eq. 6.27 (pg.355) P2W_design2=(7.784*W2S_stall_req_ips/(CLmax*sqrt(sigma))+1736.8*sqrt(W2S_stall_req_ips)/(Max_climb_parameter*sqrt(sigma)))/(33000*prop_eff); % Eq. 6.27 (pg.355)

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%% Plots Wing_loading_si=Wing_loading.*4.8927; % Range of Wing loading [kg/m^2] Power_loading_si=Power_loading.*0.6093; % Power loading curve [kg/kW] Power2Weight_si=Power2Weight.*1.6412; % Climbing curve [kg/kW] W2P_design=(1/P2W_design2)/(2.2*0.746); % Design Power loading [kg/kW] figure(2) plot(Wing_loading_si,Power_loading_si,'r*'); hold on; plot(Wing_loading_si,1./Power2Weight_si,'k.'); hold on; plot([W2S_stall_req W2S_stall_req],[0 40],'LineWidth',2); hold on; legend('Max. Speed Requirement','Climb Req.','Stall Speed Req.','location','best') title('Matching Chart') xlabel('Wing Loading (W/S), kg/m^2') ylabel('Power Loading (W/P), kg/kW') %% Weight and Size Convergence %W2P_design=12; U=0.025*(1/P2W_design2)+0.125; % Useful Load Fraction, figure 6.24 (pg.366) % Battery Powered Cruise Flight E=300; % Assumed Energy Density, Wh/kg Mfrac_battery=Range/(E*3600*prop_eff*L2D/g); % (M. Hepperle, "Electric Flight - Potential and Limitations," Eq. 9) Mfract_prop=Mfrac_battery+Mfrac_VTOL; % Summation of VTOL fuel weight and Battery Weight % Mass, Size, and Power Determination Mgross_kg=Mpayload/(U-Mfract_prop) % Gross Mass Estimate [kg] Mempty_kg=Mgross_kg*(1-U) % Empty Mass [kg] S_sqm=Mgross_kg/W2S_stall_req % Wing Area [m^2] Power_kW=Mgross_kg*P2W_design2 % Power Requirement [kW]

1. Vertical Takeoff/Landing Analysis [53]

a) Function for Vertical Lift function [t_tot,fuel_tot,Solution_A,Solution_B] = Vertical_Lift_Complete(m0_1,Isp,mdot_max,h1,h2) %% Code for optimization of vertical lift profile on Mars %by Ryan Manns %Note that the velocity is assumed to be 30 m/s in ascent and 60 m/s in descent in %order to neglect drag V1=30; %m/s V2=60; %m/s %Note that this program only calculates possible solutions. It is extremely %important to evaluate the mass flow rates of the propellent to ensure they %are not exceeding maximum possible nozzle performance

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%Number of computation points %Warning: Do not attempt more than 500 N=100; %Givens g0_E=9.81; %m/s^2 g0_M=0.38*g0_E; %m/s^2 c=Isp*g0_E; %m/s %Inputs: %m0=Initial aircraft mass (kg) %Isp=Fuel specific impulse (s) %h1=Height of ascent (m) %Note, 1000 m recommended %h2=Height of descent (m) %Note, 2000 m recommended %mdot_max=maximum fuel mass flow rate possible (kg/s) %Note that this parameter is limited by the nozzle(s) %Outputs: %t_tot=Total time of ascent and descent (s) %fuel_tot=total fuel consumed for ascent and descent (kg) %Note that the following notation is used: %1=Acceleration to V1 ascent phase %2=Constant speed V1 ascent to h1 %3=Constant speed descent at V2 from h2 %4=Decceleration phase to ground %a=ascent phase (1+2) %b=descent phase (3+4) %Solution_A=Solution vector of ascent phase parameters shown below %[Solution_A=[fuel_a_ans,... % t_a_ans,... % h_a_ans,... % t_1_ans,... % t_2_ans,... % fuel_1_ans,... % fuel_2_ans,... % h_1_ans,... % h_2_ans,... % mdot_1_ans,... % mdot_2_start_ans,... % mdot_2_end_ans]; %Solution_B=Solution vector of descent phase parameters shown below %[Solution_B=[fuel_b_ans,... % t_b_ans,... % h_b_ans,... % t_3_ans,...

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% t_4_ans,... % fuel_3_ans,... % fuel_4_ans,... % h_3_ans,... % h_4_ans,... % mdot_4_ans,... % mdot_3_start_ans,... % mdot_3_end_ans]; %% Determine most efficient ascent %Acceleration phase, vary n1 to reach delta_V required n_1=linspace(1,1.25,N); mf_1=m0_1./n_1; %kg %Find mdot required for each n mdot_1=(-g0_M.*(m0_1-mf_1))./(V1-c.*log(n_1)); %kg/s %Find time of each burn t_1=(m0_1-mf_1)./mdot_1; %s %Find fuel used in each burn fuel_1=m0_1-mf_1; %kg %Find height traveled for each case h_1=(c./mdot_1).*(mf_1.*log(1./n_1)+m0_1-mf_1)-0.5.*(((m0_1-mf_1)./mdot_1).^2)*g0_M; %m %Find remaining heights for constant speed ascent h_2=h1-h_1; %m %Find time of ascent, knowing speed of ascent and height left t_2=h_2/V1; %s %Solve for mass fraction n_2, then solve for final mass and fuel used n_2=exp((t_2*g0_M)/(Isp*g0_E)); m0_2=mf_1; %kg mf_2=m0_2./n_2; fuel_2=m0_2-mf_2; %kg %Also determine maximum mass flow rates of each condition to check %validity mdot_2_start=(m0_2*g0_M)/(Isp*g0_E); %kg/s mdot_2_end=(mf_2*g0_M)/(Isp*g0_E); %kg/s %Total time, height, and fuel used for ascent t_a=t_1+t_2; %s h_a=h_1+h_2; %m fuel_a=fuel_1+fuel_2; %kg %Eliminate impossible solutions based on impossible mass flow rates X_a=zeros(N,1);

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for i=1:N if mdot_1(i) <= 0 X_a(i)=NaN; elseif mdot_1(i) > mdot_max X_a(i)=NaN; elseif mdot_2_start(i) > mdot_max X_a(i)=NaN; elseif mdot_2_end(i) > mdot_max X_a(i)=NaN; else X_a(i)=1; end end t_a=t_a.*X_a'; %s h_a=h_a.*X_a'; %m fuel_a=fuel_a.*X_a'; %kg %Isolate most efficient ascent [fuel_a_ans,Indices_1]=min(fuel_a); t_a_ans=t_a(Indices_1); h_a_ans=h_a(Indices_1); t_1_ans=t_1(Indices_1); t_2_ans=t_2(Indices_1); fuel_1_ans=fuel_1(Indices_1); fuel_2_ans=fuel_2(Indices_1); h_1_ans=h_1(Indices_1); h_2_ans=h_2(Indices_1); mdot_1_ans=mdot_1(Indices_1); mdot_2_start_ans=mdot_2_start(Indices_1); mdot_2_end_ans=mdot_2_end(Indices_1); %% Determine most efficient descent m0_3=mf_2(Indices_1); %Assume constant speed descent from h2 %Iterate time of constant speed descent for best approximation %Find maximum time possible t_desc_max=h2/V2; %s %Create time space of possible flight times t_3=linspace(0,t_desc_max,N); %s %Determine possible flight heights h_3=V2.*t_3; %m %Determine constant speed descents of these possibilities n_3=exp((t_3*g0_M)/(Isp*g0_E)); mf_3=m0_3./n_3; %kg fuel_3=m0_3-mf_3; %kg %Determine maximum mass flow rates of each condition to check %validity mdot_3_start=(m0_3*g0_M)/(Isp*g0_E); %kg/s

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mdot_3_end=(mf_3*g0_M)/(Isp*g0_E); %kg/s %Check possibilites of final decceleration phase m0_4=mf_3; %Note that the final height has been set h_4=h2-h_3; %m %Find mdot that will provide the appropriate deltaV and height Func=@Decceleration; Options=optimoptions('fsolve','Display','Off','TolFun',1e-9,'TolX',1e-9); mdot_4_guess=mdot_max; n_4_guess=n_1(Indices_1); Guess=[mdot_4_guess,n_4_guess]; Solution=zeros(N,2); for i=1:N Solution(i,:)=fsolve(Func,Guess,Options,h_4(1,i),m0_4(1,i),V2,c,g0_M); end mdot_4=Solution(:,1); n_4=Solution(:,2); %Find mf_4 mf_4=m0_4./n_4'; %kg %Find t_4 t_4=(m0_4-mf_4)./mdot_4'; %s %Find fuel used fuel_4=m0_4-mf_4; %kg %Add togethor all descent solutions t_b=t_3+t_4; %s h_b=h_3+h_4; %m fuel_b=fuel_3+fuel_4; %kg %Eliminate impossible solutions based on impossible mass flow rates X_b=zeros(N,1); for i=1:N if mdot_4(i) <= 0 X_b(i)=NaN; elseif mdot_4(i) > mdot_max X_b(i)=NaN; elseif mdot_3_start > mdot_max X_b(i)=NaN; elseif mdot_3_end(i) > mdot_max X_b(i)=NaN; else X_b(i)=1; end end t_b=t_b.*X_b'; %s

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h_b=h_b.*X_b'; %m fuel_b=fuel_b.*X_b'; %kg %Determine most efficient [fuel_b_ans,Indices_2]=min(fuel_b); t_b_ans=t_b(Indices_2); h_b_ans=h_b(Indices_2); t_3_ans=t_3(Indices_2); t_4_ans=t_4(Indices_2); fuel_3_ans=fuel_3(Indices_2); fuel_4_ans=fuel_4(Indices_2); h_3_ans=h_3(Indices_2); h_4_ans=h_4(Indices_2); mdot_4_ans=mdot_4(Indices_2); mdot_3_start_ans=mdot_3_start; mdot_3_end_ans=mdot_3_end(Indices_2); %Add togethor totals and solution vectors fuel_tot=fuel_a_ans+fuel_b_ans; %kg t_tot=t_a_ans+t_b_ans; %s Solution_A=[fuel_a_ans,... t_a_ans,... h_a_ans,... t_1_ans,... t_2_ans,... fuel_1_ans,... fuel_2_ans,... h_1_ans,... h_2_ans,... mdot_1_ans,... mdot_2_start_ans,... mdot_2_end_ans]; Solution_B=[fuel_b_ans,... t_b_ans,... h_b_ans,... t_3_ans,... t_4_ans,... fuel_3_ans,... fuel_4_ans,... h_3_ans,... h_4_ans,... mdot_4_ans,... mdot_3_start_ans,... mdot_3_end_ans]; function [Sol] = Decceleration (Guess,h_4,m0_4,V2,c,g0_M) Sol(1)=(c)*log(Guess(2))-(g0_M/Guess(1))*(m0_4)*(1-(1/Guess(2)))-V2; Sol(2)=(c/Guess(1))*(m0_4)*((1/Guess(2))*log(1/Guess(2))+1-(1/Guess(2)))-0.5.*(((m0_4)*(1-(1/Guess(2)))./Guess(1)).^2)*g0_M-h_4; %m end end

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b) Script for Fuel Fraction Mass for VTOL Maneuver %Vertical Lift Profile %Ascend and descend on destination, then repeat to return %By Ryan Manns clc clear close all %Set Aircraft weight range N=10; Mass=linspace(100,2000,N); %kg %Set parameters Isp=295; %s h1=1000; %m h2=2000; %m mdot_max=10; %kg/s %Iterate profile fuel_tot_1=zeros(N,1); for i=1:N [~,fuel_tot_1(i),~,~]=Vertical_Lift_Complete(Mass(i),Isp,mdot_max,h1,h2); end %Find new mass value Mass_2=Mass-fuel_tot_1'; %Repeat return trip for new mass fuel_tot_2=zeros(N,1); for i=1:N [~,fuel_tot_2(i),~,~]=Vertical_Lift_Complete(Mass_2(i),Isp,mdot_max,h1,h2); end %Find total fuel used fuel=fuel_tot_1+fuel_tot_2; %kg %Find remaining weight Mass_Remain=Mass-fuel'; %kg %Find mass fraction for fuel Mass_frac=Mass_Remain./Mass; %Plot figure plot(Mass,fuel) xlabel('Aircraft Initial Mass (kg)') ylabel('Fuel Used (kg)') figure plot(Mass,Mass,Mass,Mass_Remain)

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legend('Mass of Aircraft','Mass After Profile') xlabel('Aircraft Initial Mass (kg)') ylabel('Mass (kg)') figure plot(Mass,Mass_frac) xlabel('Aircraft Initial Mass (kg)') ylabel('Fuel Mass Fraction') figure plot(Mass,fuel'./Mass) xlabel('Aircraft Initial Mass (kg)') ylabel('Fuel/Mass')

2. Thrust and Power Requirements % updated: 08/01/16 % Power Required Code close all;clc;clear all; % Input Parameters h=0:100:5000; Vcruise=input('Enter flight speed in m/s \n'); % Atmosphere if h>= 7000 t=-23.4-0.00222.*h; %Celcius p=0.699.*exp(-0.00009.*h); %K-Pa else t=-31-0.000998.*h; %Celcius p=.699.*exp(-0.00009.*h); %K-Pa end rho=p./(.1921.*(t+273.1)); %kg/cu m q=0.5.*rho.*Vcruise.^2; % Aircraft Parameters, from table (updated, 07/26) Mass=1500; S=110; g=3.71; Weight=Mass*g; W2S=Weight/S; CLmax=1.8; AR=17.6; eps=0.7; CD0=0.025; K=1/(pi*eps*AR); L2D=24; L2D_min_power=0.25*sqrt(3/(CD0*K)); L2D_min_drag=0.5*1/sqrt(CD0*K); % Speed and Climb

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Vstall=sqrt(2.*W2S./(rho.*CLmax)); % Stall Speed Determination, m/s Vstall_mph=Vstall.*2.237; % Conversion of Vstall to mph ROC=0.02.*Vstall_mph.^2; % Rate of Climb, ft/min ROC_si=ROC.*0.3048./60; % Conversion of ROC to m/s if Vstall>Vcruise warning('Cruise speed is less than stall speed') end % Thrust Required Treq=CD0.*q.*S+(K.*Weight.^2)./(q.*S); % Treq including induced drag, N Treq1=Weight./L2D; % First order estimate of Thrust Required, N Treq2=CD0.*q.*S; % Base Thrust Required, N Preq=Vcruise.*Treq; Pava=(ROC_si.*Weight+Preq); % Necessary Power, kW Tava=Pava./Vcruise; % Thrust Available (engine must provide it), N figure(1) plot(h,Preq/1000); hold on; plot(h,Pava/1000,'r'); xlabel('Height, m') ylabel('Power, kW') title('Power Available and Power Required') legend('Min. Power Required','Power Available','location','best') figure(2) plot(h,Treq); hold on; plot(h,Tava,'r'); xlabel('Height, m') ylabel('Thrust, N') title('Thrust Available and Thrust Required') legend('Min. Thrust Required','Thrust Available','location','best')

3. Cost Estimation Function % Cost Estimation Functions % Updated: 08/01/16 % Based on Jan Roskam, Vol 1 and Vol 8 g=9.81; % Earth Gravity Mto=1623; % Aircraft Mass [kg] Wto=Mto*g; % Aircraft Weight [N] Wto_lb=Mto*2.2; % Aircraft Weight [lb] Wampr=10^(0.1936+0.8645*log10(Wto_lb)); % Vol 8, Eq. 3.5 (pg.67 PDF) CEF_2016=6.27; CEF_1973=1.14; % Cost Escalation Factor Nprot=1; % Number of prototypes to be built Cprot=1115.4*10^3*(Wampr^0.35)*(Nprot^0.99)*(CEF_2016/CEF_1973); % Vol 8, Eq. 3.20 (pg.67 PDF)

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XI. Acknowledgments The author, as the Chief Engineer of Aurora Concepts, would like to thank Dr. Casey Handmer for

his willingness and attention to our inquiries.

Similarly, thanks to Guillaume De Forton from OXIS Energy Limited for providing us with information about their Ultra Light Batteries.

It is important to acknowledge the work of the course instructor, Dr. Bernd Chudoba, and the teaching assistants, Thomas McCall, James Haley, and Loveneesh Rana. Their guidance and feedback was greatly appreciated at every stage of this project.

Finally, and more importantly, the author would like to acknowledge the work performed by the members of Aurora Concepts: Ismael Sanabria, Nic Dwyer, Juan Lopez, Yasir Rauf, Ryan Manns, Justin Kenna, and Ian Maynard. They were the backbone of this project and the main contributors to the work presented in this report.

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salas nunez, luis, mars aerial platform, chief engineer, final report (final)

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