Executive Summary (Gersh)



The Daedalus One

System Definition Review

Daedalus Aviation-Team 1

AAE 451-Spring 2008

March 26, 2008

Design Team

James Bearman

AJ Brinker

Dean Bryson

Brian Gershkoff

Kuo Guo

Joseph Henrich

Aaron Smith

Executive Summary

Daedalus Aviation is a conceptual design firm that is creating the aircraft of the future. This report is a definition of the aircraft concept, including aerodynamic configuration and fuselage layout, a detailed investigation into advanced technologies, and an outline of the engineering constraint analysis and aircraft sizing predictions.

To do this, Daedalus Aviation has generated several aircraft concepts, modified and improved them, finally arriving at a working configuration that is being used for the analysis: currently, a high-aspect ratio rear wing with a lifting canard in the front, propelled by two over-wing mounted engines. Additionally, the basic layout of the fuselage that could seat one hundred eight people, and have ample space for lavatories and galley spaces went through a careful design process.

Deadalus One, the aircraft under development, is scheduled to be released for sale in the late 2030s and early 2040s. The intervening twenty to thirty years will see marked advances in aviation technology, and Daedalus Aviation will capitalize on technologies currently in development that will be ready for mainstream use by 2040. These technologies include everything from artificial intelligence aircraft control to vectored thrust; from composite materials to blown flaps.

Currently, the Daedalus One is within acceptable range for all engineering requirements as calculated by our specialized aircraft sizing program. According to the computer models used by Daedalus Aviation, the Daedalus One

• Has a range of 1800 nautical miles

• Carries 108 passengers

• Can takeoff and land in 2500 feet

• Cruises at .75 M

Table of Contents

Executive Summary 2

1. Business Case 6

1.1. Mission Statement 6

1.2. Use-Case 6

1.3. Design Requirements 8

2. Concept Selection 8

2.1. Overview 8

2.2. Ideation 9

2.2.1. “Conventional” 9

2.2.2. “Tandem Wing” 10

2.2.3. “Bird of Prey” 11

2.2.4. “Air Whale” 12

2.2.5. “Turtle with Jetpacks” 13

2.2.6. “Canard Ray” 14

2.2.7. “Inverted Ray” 15

2.2.8. “Bat Plane” 16

2.2.9. “Hammerhead” 17

2.2.10. “Conventional ++” 18

2.2.11. “Sharkbat” 19

3. Current Configuration 20

3.1. Description 20

3.2 Fuselage Layout 20

3.2.1. Design Justification 20

3.2.2. Seating Parameters 21

3.2.3. Galleys, Lavatories, and Cockpit Parameters 21

3.2.4. Emergency Exits 21

3.2.5. Overall Dimensions 21

3.2.6. Fuselage Diagram 22

4. Potential Technologies 22

4.1 Weight/Cost Saving 22

4.1.1 Composites 22

4.1.2. AI/UAV 22

4.2. Propulsion Types 23

4.2.1. Pulse Detonation Engines 23

4.2.2. Unducted Fans 23

4.2.3. Geared Turbofans 24

4.3. Propulsion Enhancing Devices 24

4.3.1. Vectored Thrust 24

4.3.2. Magnetic Bearings 25

4.4. High-Lift Devices 25

4.4.1. Blown Flaps 25

4.4.2. Circulation Control 27

4.4.3. Co-Flow Jet Flow Control 28

4.4. Technology Readiness Level Summary 29

5. Constraint Analysis and Sizing 30

5.1. Performance Constraints 30

5.1.1. Takeoff Roll 32

5.1.2. Landing Roll 32

5.1.3. Takeoff Field Length—One Engine Inoperative 33

5.1.4. Second Segment Climb Rate—One Engine Inoperative 34

5.1.5. Climb Rate at Top of Climb 34

5.1.6. Climb Rate at Service Ceiling 35

5.1.7. 2G Maneuver 35

5.1.8. Steady, Level Cruise 36

5.2. Constraint Analysis Results 36

5.3 Sizing 38

5.3.1. Sizing Software Selection 38

5.3.2 RDS Sizing Method 38

5.3.3. Sizing Inputs and Results 39

6. Conclusion 43

6.1 Walk Around Diagram 43

6.2 System Compliance Matrix 43

6.3. Next Steps 44

References 45

1. Business Case

1.1. Mission Statement

The goal of Daedalus Aviation is to provide a versatile aircraft with a medium range and medium capacity to meet the needs of a commercial aircraft market that is still expanding in 2058. Daedalus Aviation will incorporate the latest technology into our aircraft so as to improve efficiency and reliability. This is particularly important when an environmentally friendly aircraft is demanded by not only the general public, but also by the airlines as well. The technologies that will most improve the versatility of the aircraft are the enabling technologies that allow the aircraft to have Extremely Short Take-Off and Landing (ESTOL) capability.

1.2. Use-Case

The following mission profile and use case is designed to demonstrate the maximum range operation of the Daedalus One to its fullest extent. Because maximum range accompanied with a missed approach is the most fuel strenuous scenario of the Daedalus One, it was used for sizing of the aircraft.

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Figure 1: Maximum range mission profile.

The mission profile is a simple long range mission as shown in Figure 1. The mission starts with a startup and taxi out to the active runway. The aircraft then takes off and climbs to an altitude of 10,000 ft at best rate of climb to clear the airspace quickly. The climb is then switched to a cruise climb which typically only sacrifices little vertical speed to gain ground speed and give the pilots better visibility. Once the assigned altitude has been reached, typically around 25,000 ft to 30,000 ft this altitude can either be held for the duration of the cruise or if a higher altitude is desired, a step climb can be used to clear weather, reach maximum cruising altitude, or to avoid turbulence. After the cruise portion, the profile goes into a descent and hold. The hold is included to allow time for busy airports where ATC cannot vector the aircraft in right away and is told to hold away from the airport.

Due to the fact that the plane will be used in instrument metrological conditions, it must follow the instrument flight rules. By these rules, the plane must hold enough fuel to proceed to the alternate airport listed as filed in the flight plan and have 45 minutes of fuel in reserve. To comply with this regulation we have given the plane the capability to climb out from a missed approach, proceed to an alternate airport that could be up to 200 nmi away and land.

A use-case that could be fulfilled by the Daedalus One using the above mission profile is from South Bend, Indiana, to Burbank, California. This would be representative of a chartered route for a football team with staff traveling to a rival college. This type of route maximizes convenience for the passenger by using small, nearby airports, in this case South Bend Regional near Notre Dame University and Burbank Bob Hope near University of Southern California as outline in Table 1.

|Airport |Location |Range (nmi) |Runway Length (ft) |

|South Bend Regional (SBN) |South Bend, IN |1580 |4300 |

|Bob Hope Airport (BUR) |Burbank, CA | |5800 |

Table 1: Ranges and runway length for use-case 2.

1.3. Design Requirements

The primary engineering requirements developed by Daedalus Aviation can be seen in Table 2.

|Engineering Requirement |Condition |Target |Threshold |

|Takeoff Distance |≤ |2,500 ft |3,500 ft |

|Landing Distance |≤ |2,500 ft |3,500 ft |

|Takeoff Weight |≤ |80,000 lb |100,000 lb |

|Range |≥ |1800 nm |1500 nm |

|Maximum Cruise Speed |≥ |0.85 M |0.75 M |

|Maximum Passenger Capacity |≥ |110 |90 |

|Payload Capacity |≥ |28,300 lb |23,300 lb |

|Fuel Burn Per Seat Mile |≤ |0.10 lbs/(pax-nm) |0.12 lbs/(pax-nm) |

Table 2: Engineering Requirement for Daedalus One

The takeoff and landing distance requirements are set so as to provide the ESTOL capability for the aircraft, allowing the Daedalus One to operate at nearly any airport in the world. Our takeoff weight requirement is restricted for the purposes of ramp weight restrictions at airports. The takeoff weight and fuel burn are both minimized to reduce operating and acquisition cost. Range, speed, and capacity, are tailored to fit the niche outlined in the Daedalus Aviation mission statement (section 1.1).

2. Concept Selection

2.1. Overview

The major goal in this phase of the design process is to generate an aircraft concept configuration. To do this, Daedalus Aviation used Pugh’s Method to evaluate ideas and narrow the field to a final set of concepts and then finally choose one. Pugh’s Method is effectively a five step process: generate criterion, generate concepts, choose a datum, evaluate, and improve. All designs are evaluated qualitatively relative to the datum in the categories stated by the selected criterion. At this point, some designs are scrapped, others combined, and some improved upon. One of the “new” designs is selected as the datum for the next round of evaluation, and the next iteration of evaluation and improvement begins. These designs are evaluated and improved upon over and over again until a final design choice is selected.

2.2. Ideation

Daedalus Aviation went through two rounds of ideation and improvement, enhancing the positive aspects of each aircraft, combining the best ideas, and eliminating the negative aspects. The following section is an index of all concepts generated by Daedalus Aviation.

2.2.1. “Conventional”

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Figure 2: Conventional

Most aircraft today have the same basic configuration: a tube-shaped fuselage with a low main wing, tricycle landing gear, a conventional tail, and two turbofans located under the wings near the fuselage. Recognizing that aircraft with these configurations are obviously successful, Daedalus Aviation determined that it would be unwise to not fully explore this concept during our ideation. This configuration functioned as the datum for the first iteration of Pugh’s Method, as it is the most obvious basis of comparison for any unconventional aircraft.

The Conventional design is a popular, proven concept. The aircraft gives good visibility to passengers and crew, good jetway access, easy access to engine maintenance areas, convenient locations for the landing gear, and an easily pressurized fuselage. The air flowing into the engines is very aerodynamically clean, and the aircraft is naturally stable. The design has no major problems associated with it, as the baseline, this aircraft doesn’t do anything new, different, or better.

2.2.2. “Tandem Wing”

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Figure 3: Tandem Wing

The Tandem Wing concept aircraft is shown in Figure 3. This design features the common circular fuselage cross-section associated with most all aircraft in production today. The major design aspect of this aircraft is that it makes use of tandem wings. The wings in the front of the plane are moved farther forward than in the conventional layout and are similar in size to the wings at the rear of the plane. This design makes use of two turbofan engines which are mounted to the fuselage at the rear of the aircraft, underneath the rear wing.

The Tandem Wing theoretically will reduce the induced drag by 50%. Induced drag is a function of the square of the lift being produced which means that if the weight of the aircraft is evenly distributed to two wings, each wing would have only one-fourth of the induced drag of a single wing [1]. The main disadvantage of this design is that the second wing will be flying in the downwash of the first wing and in order to compensate for this the second wing would need to be mounted at a higher angle of incidence, this would cause the lift of the back wing to have a component in the rearward direction creating a drag term. Effectively, this more than negates the theoretical induced drag advantage of the tandem wing configuration.

2.2.3. “Bird of Prey”

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Figure 4: Bird of Prey

This concept is derived from the design of the Concorde, given the nickname of “Bird of Prey” due to the shape similarity with the namesake ship from Star Trek. It is a long cylindrical fuselage with a rear delta wing. However in a departure from the Concorde, a more blended wing body was to occur in the wing, maximizing passenger space. The propulsion system is placed at the aft of the aircraft to allow for possible thrust vectoring. This location also allows for lower noise from the engines inside the cabin.

The primary benefits of this design were from the placement of the engine system as well as passenger space. This particular engine placement would allow for better takeoff performance. However the design would encounter some problems shared with the Concorde. The primary flaw is the delta wing. The Concorde had a delta wing because it was ideal for supersonic flight. Supersonic flight is not a design requirement of this aircraft which makes the delta wing a hindrance to the aerodynamic efficiency.

2.2.4. “Air Whale”

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Figure 5: Air Whale

The Air Whale, shown in Figure 5 is a very unique concept in that it combines the designs of both a blimp as well as a traditional aircraft. The fuselage is embedded in a large cavity which is intended to be filled with a lighter than air, stable gas such as helium. Attached to the large fuselage design are conventional aircraft wings and a conventional tail. There are four turboprop engines, two under each wing.

The main advantage of this design is that the use of a large space filled with lighter than air gas would provide exceptionally short take-off distances. However, due to the large size of the aircraft there are many disadvantages. Even though the engines are mounted underneath the wings, the body of the aircraft is so large the wings would still be very high off the ground, making the engines difficult maintain. The aircraft also has a very large wetted area which would result in a tremendous amount of parasite drag. With the large body and parasite drag, achieving acceptable cruise speeds would be quite difficult.

2.2.5. “Turtle with Jetpacks”

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Figure 6: Turtle with Jetpacks

Typically, the Achilles Heel of designing an aircraft is moving the center of gravity far enough forward such that the design is statically stable. With so much weight needing to be in the tail for control systems, stabilizers, and many times auxiliary power units, it is easy to understand how this becomes a challenge. Additionally, ESTOL capability is a primary concern for Daedalus Aviation. The design which affectionately became known as the “Turtle with Jetpacks” attempts to solve both of these problems simultaneously. A control canard is located in the front of the aircraft, along with the engines, mounted to the canard tips. Furthermore, if the canard functions as a stabilator, when the aircraft pitches up or down, the thrust is automatically vectored in the proper direction. The main lifting wing is in the rear with large winglets which could be used as additional vertical stabilizers.

The main advantage of this aircraft is that it has a huge static margin due to the engines being so far forward. Pitch control with the canard and auto-vectoring thrust is very good, which should reduce the takeoff roll significantly. Landing gear placement is also very easy, and there is plenty of space for fuel in the aircraft as well.

The biggest problem is that this aircraft is a structurally very difficult to develop. The canard area must be designed such that it can not only support the engines, but also move the entire surface, and support huge bending moments from the engines being so far outboard. The air coming to the main lifting wing is also “dirty” having either flowed over the canard or through the engine. There is also concern that this disturbed air will interact in some way with the yawing control surfaces. Despite the magnitude of advantages this aircraft had, the sheer number of disadvantages this configuration possessed made it undesirable for further consideration.

2.2.6. “Canard Ray”

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Figure 7: Canard Ray

The aircraft concept Canard Ray is shown in Figure 7. The Canard Ray features a conventional forward fuselage section, but, in the aft section, the fuselage is blended with the large wings. The wing has a large chord and thickness at the root to accommodate blending into the fuselage, and tapers down to a more typical scale along the span. The taper is most pronounced at the root where the wing section quickly scales down, and at the tip where the wing tips are raked back slightly. The Canard Ray has a conventional vertical stabilizer and low-mounted canards. Turbofan type engines (either geared or direct-drive) are mounted above the wing at the leading edge. These are not shown in Figure 7, but were specified in the concept description.

Major advantages of the Canard Ray are its predicted reduction in gross takeoff weight, reduced fuel burn, and increased lift-to-drag ratio. These predictions are based on significant improvements obtained by the true blended-wing-body [1]. Because of the of the large wing root, there is ample space for fuel, and potentially an expanded passenger cabin. The high mounted engines allow for upper-surface blown flaps, though the large wing area reduces the required wing loading and has the potential to eliminate the need for high lift devices all together. Eliminating high lift devices would reduce aerodynamic noise. The Canard Ray would have good compatibility with small airports because the aircraft is low to the ground and could use an air stair instead of a jetway, and the reduction in empty weight and gross takeoff weight would be more likely to meet ramp weight restrictions. The canard is also beneficial because it produces positive lift during takeoff and landing.

A major disadvantage of the Canard Ray is passenger visibility. If the passenger cabin was extended into the wing root, it may not be feasible to put windows in that part of the cabin. Also, views in the forward cabin could be blocked by the canard. The canard can also make it difficult for the aircraft to pull up to a jetway. Drag could also potentially be increased over the conventional design. The large wing area lends itself to a lower aspect ratio, which would lead to an increase in induced drag. The Canard Ray could also potentially have a much larger wetted area compared to a conventional aircraft, increasing parasite drag. Maintenance on the Canard Ray would also be made more difficult because of the above wing engines.

2.2.7. “Inverted Ray”

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Figure 8: Inverted Ray

The Inverted Ray concept, shown in Figure 8, is very similar to the Canard Ray. The difference is that the canards and vertical tail of the Canard Ray are replaced with a top-mounted, inverted V-tail. In general, the advantages and disadvantages of the Canard Ray (except for those relating directly to the canards and vertical tail) are expected to hold true to the Inverted Ray.

There are a few advantages to having the inverted V-tail instead of the canard. One advantage is that eliminating the canard makes the aircraft much easier to pull up to a jetway and keeps the stabilizer away from ground operations, which could cause structural damage. The V-tail also eliminates a control surface by going from a rudder and two elevators to two ruddervators. The use of an inverted V also provides a proverse roll-yaw coupling [1]. An empty weight improvement might also be seen as the two tail surfaces would brace each other, and less material would be needed to produce adequate stiffness.

The inverted V-tail also has some disadvantages. By combining the elevator and rudder controls, complexity is added to the controls and actuation. The joint region would also be expected to increase interference drag. Finally, during takeoff and landing, the tail would produce down force instead of lift to cause the aircraft to pitch up. This down force would not be desirable in an ESTOL situation.

2.2.8. “Bat Plane”

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Figure 9: Bat Plane

This concept is a modification of the two “Ray” (Section 2.2.6 and 2.2.7) designs. It was intended to provide improvements to the designs for the second iteration of Pugh’s method. The primary modifications made for this design was the selection of a pulse detonation engine for the propulsion system, which was then embedded inside the wing. The other change was T-tail as opposed to the inverted V tail or canards and a standard vertical tail.

The embedding of the engines improved the drag properties which arose in the “Ray” designs. This greatly reduced the parasite drag while not greatly affecting the induced drag. Upon further analysis it was found that the T-tail made performance worse for the aircraft. The tail would not have allowed a great enough moment arm to provide proper control for the aircraft. Aspects from this design, however, were used in later concepts.

2.2.9. “Hammerhead”

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Figure 10: Hammerhead

The Hammerhead design configuration in Figure 10 has a control canard mounted on the lower front of the plane. The main wing is a high mounted wing on the rear of the aircraft with three turbofan engines between a twin tail mounted on the top. The fuselage is rather large in that it is two floors high.

An advantages of the plane is minimal noise for the passengers because of the rear mounted engines. A major disadvantage for the aircraft is the mounting position of the canard. The canard is fairly large and because of the position, bringing a jetway up to the plane will prove to be complicated. Along with canard placement, the placement of the engines above the rear wing will increase the maintenance costs for servicing due to the fact that they are harder to reach than wings underneath the plane.

2.2.10. “Conventional ++”

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Figure 11: Conventional ++

The Conventional ++ design concept shown in Figure 11 closely resembles aircraft that are currently in production. The fuselage is the standard cigar shape with wings that are swept back and a Conventional tail. Contrary to most current configurations, the engines are mounted above the wings.

This design concept was intended to be made mostly out of composites in order to reduce weight. The placement of the engines on top of the wings allows for upper surface blowing of the flaps which will greatly increase the lift coefficient for takeoff and landing. Because the engines are mounted on top of the wing, this configuration will increase the maintenance cost because they will not be as accessible compared to under wing mounted engines.

2.2.11. “Sharkbat”

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Figure 12: Sharkbat

The “Sharkbat” was one of the final concepts generated, using ideas from many of our previous concepts, and became the basis for our current configuration. A lifting canard is located far forward, with the main wing in the far aft. In this concept, the propulsive system is pulse-detonation engines imbedded into the aircraft near the wing root. The winglets on the main wing could be used as additional vertical stabilizers if additional yaw control is required.

3. Current Configuration

3.1. Description

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Figure 13: Four View of Concept

The current configuration of the Deadalus One is based on the Sharkbat design. It features two geared turbofans for main propulsion, located above the main wing. The wing is located far aft and high on the fuselage, while the canard is far forward and low-mounted. The wing has anhedral, and curves up at the ends in winglets or additional vertical stabilizers, allowing for a possible tri-tail if the control surface sizing requires it. The canard has a slight dihedral and sweep. The landing gear (not shown) are in a tricycle configuration with the main gear mounted in the aft fuselage.

3.2 Fuselage Layout

3.2.1. Design Justification

The initial cabin design was chosen with keeping the passengers in mind. By using personal experience and some industry standards, a cabin layout was designed to optimize a single aisle layout while tailoring to external aerodynamic effects.

3.2.2. Seating Parameters

With passenger comfort as one of the top priorities, the design errs to the larger side of coach seats. A seat width of 20 in. and a seat pitch of 32 in. should provide passengers with plenty of room to stretch out. The armrests are a typical 2 in. width. To provide ample room for the galley cart, the aisle has a width of 24 in. to prevent passengers from being bumped into when it comes time to serve drinks.

To accommodate the flight attendants, there will be one fold down seat attached to each galley wall near the emergency exits so they are in optimum position in case of an emergency.

3.2.3. Galleys, Lavatories, and Cockpit Parameters

To make the passengers feel like they are at home, the lavatories have been given a decent area of 20 square feet. The galleys have also been given ample room to provide plenty of preparation area and storage. The rear galley is currently sized at 7 ft. by 5 ft. and the front galley is 4 ft. by 4 ft. The cockpit will also err to the larger side by having a depth of 8 ft.

3.2.4. Emergency Exits

The Federal Aviation Regulations (FARs) state that the plane must be completely evacuated of passengers in 90 seconds or less. It also states that no passenger may be more than 60 feet from any exit. In order to accommodate these requirements, the aircraft currently has four Class II exits, two in the front of the aircraft and two in the rear. Included on the aircraft but not depicted on the drawing, will be two Class III exits with blow up slides since there is no wing to exit out onto at row nine.

3.2.5. Overall Dimensions

Currently the cabin length including the cockpit is 76 ft. This helps in providing a slenderness ratio of 9 for the overall aircraft. The width of the aircraft is 13 ft which includes the 2 in. fuselage thickness on either side of the aircraft.

6 Fuselage Diagram

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Figure 14: Daedalus One Cabin Layout

4. Potential Technologies

4.1 Weight/Cost Saving

4.1.1 Composites

Composite materials are a great benefit to the aviation industry. Not only are they much lighter than metals, but also much stronger. Composites utilize a fiber material to reinforce the structure, which is held together by a matrix material. Through this construction, the composite’s strength is greater than that of any of its component parts.

Weight is the enemy of any aircraft. The structure alone accounts for a great deal of the weight on today’s airplanes. By using all-composite construction, this weight can be reduced significantly. Current approximations of these weights savings range from 20% to 30% [2]. This is very beneficial in terms of fuel burn, take-off distance, and a number of other aspects. Because of this, many airplanes that are currently in their design phase have a significant portion of their construction made of composite materials.

However, there are a few issues with composites. Building large composite structures can be difficult with today’s technology. Since molds of very large parts such as the fuselage or wings do not exist, they need to be created by gluing together smaller parts. This can cause a weakness in the structure at the point where they connect. Also, the repair of composites is not an easy or cheap process. In general, composites are more expensive than traditional materials. The technology is still fairly new, but very promising. Daedalus Aviation is confident that by 2058, these issues will have been resolved, and an all-composite construction of the aircraft will be possible.

4.1.2. AI/UAV

Artificial Intelligence (AI) and Unmanned Aerial Vehicle (UAV) technology technologies are other innovations that will make the Daedalus One safer and more efficient. In these types of aircraft, the pilot is replaced by either a computer or a remote operator. Current aircrafts using these technologies are typically smaller, but there is no reason why the technology couldn’t be eventually scaled up for operation on a commercial jet such the Daedalus One over the next few decades. Instead of two skilled pilots, only one pilot would be needed to make sure the computer keeps operating, and be able to take control in case of emergency. Eventually, even human air traffic controllers could be replaced by an AI-controlled system. The computer would never get tired, sick, or need a break, and would never make mistakes. Also, they would not require the years of training humans need to be effective at the job.

At first people may be afraid of surrendering control of an aircraft to a machine. Also, there would still be a human in the cockpit that knows how to fly, so even if something happens to the computer, the airplane would not completely lose control. Once people realize how much safer air travel will be when computers do the flying, their fears will be alleviated, and the technology will be commonplace.

4.2. Propulsion Types

4.2.1. Pulse Detonation Engines

Pulse detonation engines (PDEs) are being considered by Daedalus Aviation for their increased fuel savings and their lack of complexity. Pulse detonation engines are an up and coming technology currently being flight tested in small aircraft. By 2038 when the aircraft is to go into production, PDE technology will be mature and fully ready to enter the market, if it hasn’t already.

The PDE would be used as the main propulsion device and could be integrated into the wing structure and would not have to stick out into the free stream flow, reducing the ram drag caused by traditional turbofans. PDEs are also predicted to have a 10% fuel savings over current engines [3]. These fuel savings could be used to increase the range of the aircraft or lower the fuel required for a flight, thereby reducing the takeoff gross weight to help meet our extremely short takeoff distance. Due to their design, PDEs also have the capability to burn multiple types of fuel thereby decreasing the reliance on oil based fuels, thus opening the gates for many alternative fuel choices.

PDEs also have the capability to be used as a replacement for a combustor in current high bypass turbofans. This will not have any effect on the drag of the aircraft but will have a significant impact on fuel efficiency. The current obstacle facing PDEs is getting the frequency of detonation high enough to make it a continuous flow and making efficient and reliable starting procedures.

4.2.2. Unducted Fans

Unducted fans are a type of engine technology with the goal of increasing fuel economy. They look very similar to a typical turbofan engine, with the exception that the blades are on the outside of the engine, giving it the appearance of a very large turboprop. This essentially makes it a very high bypass ratio turbofan, with efficiency similar to a turboprop. Due to the fuel savings from this type of engine, the range is also significantly increased.

One problem with using these types of engines is the noise factor. Early tests yielded high cabin noise. However, more modern tests have met noise criteria, but could still use improvements. Also, since there is no cowling, the risk of a blade out exists. This happens when the propeller blades become separated from the rest of the engine during operation. Obviously, this would be very hazardous for anyone unlucky enough to be inside the cabin or near the engine.

4.2.3. Geared Turbofans

Geared turbofans are also being considered for the aircraft because of their 12% savings in fuel, 40% reduction in maintenance cost, and 70% lower emissions [4]. This is a huge improvement over current technology and provides much needed reductions in fuel and emissions to be competitive in 2058. Geared turbofans currently have a technology readiness level of six because Pratt and Whitney is currently producing the engine for Airbus and is in the process of ground testing. The technology will only improve by 2038 and will be mainstream. This will give it a technology readiness level of nine by then.

These savings are made possible by having the high by-pass fan geared to rotate at a slower speed than the core. This allows the blade size to increase and increase the by-pass ratio. Because of this, the number of blades inside the compressor and turbine can be reduced thereby reducing maintenance costs. These improvements inherently also reduce fuel burn and emissions by reducing the amount of air traveling through the core. Drag will however increase on the aircraft due to the larger nacelles required by the engines.

4.3. Propulsion Enhancing Devices

4.3.1. Vectored Thrust

In order to achieve the short takeoff distance goal of 2500 feet, different types of propulsion enhancement devices are being considered. Among these is the vectored thrust technology. This is already a tried and true enhancement with such examples as the AV-8B Harrier II and the F-22 Raptor. Both aircraft are capable of doing short takeoffs using multidirectional nozzles providing lift and improved maneuverability. Since vectored thrust is currently being used, it has a technology readiness level of nine.

The Daedalus One could use this technology to enhance the lifting capabilities of the wings at lower speeds. By being able to direct some of the thrust force downwards, this will decrease the weight required to be lifted by the wings, thereby decreasing either the wings size themselves, decreasing takeoff speed, or decreasing required takeoff coefficient of lift. Vectored thrust could only be used on the pulse detonation engines or the geared turbofans, not the unducted fan.

4.3.2. Magnetic Bearings

With the focus of today’s engines on reducing emissions and increasing efficiency, having little to no internal friction on the main rotor would have a huge impact on the two mentioned items. With magnetic bearings this would be feasible. Currently in the research and development phase with only a few practical tests done, it is highly likely that by 2038 magnetic bearings will be in use with the rate in advancement in technology.

This technology could have a significant impact on the aircraft’s performance. Not only being limited to improving the engine, the magnetic bearings could also be used on the landing gear to reduce takeoff ground roll by reducing the friction on the main hubs of the wheels. Magnetic bearings could also reduce emissions and increase efficiency because less power would be lost due to friction and would allow the engine to run at a cooler temperature. The hurdle now facing magnetic bearings is the heat generated by the electromagnets while in operation. This heat is significant enough to cause structural damage with materials available today which may not be a problem in thirty years or may require a cooling system to operate.

4.4. High-Lift Devices

4.4.1. Blown Flaps

This system is also known as the Boundary Layer Control System (BLCS). This system taps into the compressor section of the engine and bleeds air through the wing to be vented through the flaps. This increases the amount of lift on the wing. This system was first employed on the Lockheed F-104 Starfighter with great success. The system helped create a large amount of lift for the Starfighter’s small wing surface. A disadvantage to this system is its complexity. It is a highly difficult system to maintain and manufacture, which results in greater expenses for the manufacturer as well as the customer. Because blown flaps have already been produced and put into production on some airplanes they have a technology readiness level of nine.

There are three major types of blown flaps shown in Figure 15.

[pic]

Figure 15: Examples of Blown Flaps [5]

Internally blown flaps, which use compressor bleed air which is vented through the wing and then out over the flaps, are the device employed on the F-104 Starfighter. These types of flaps use compressor bleed air which is vented through the wing and then out over the flaps.

Externally blown flaps, like those used on the C-17 Globemaster use the direct exhaust of the engines to help increase lift, and allow the Globemaster to take off on runways as short as 3500 feet. Upper surface blowing is another form of externally blown flaps. This type of blowing uses the Coanda effect, which is the ability of a fluid such as air to follow a surface. Figure 16 shows that blown flaps can increase the lift coefficient up to a value of nine [5].

[pic]

Figure 16: Performance of Blown Flaps [5]

Overall, blown flaps are being considered for Daedalus One because of the drastic increase in lift for the takeoff portion of the mission. It is important to keep in mind; however, that having a high coefficient of lift results in higher amounts of induced drag. Therefore, the lowest coefficient of lift that gives Daedalus One the required performance factors is under research.

4.4.2. Circulation Control

Circulation control is another type of flow control designed to create high lift at low speeds, similar to blown flaps. Figure 17 shows the basic idea of a circulation control wing. A jet of air is shot through a slot at the trailing edge of the wing energizing the boundary layer. The airflow then follows what is known as a Coanda surface. The sheet of air stays attached to the surface due to a balance in pressure and centrifugal forces. The faster the air is blown through the slot the farther around the surface it travels. This high lift technology has been tested and proven effective in labs but has not been commercially put into effect. By 2058 however, enough will be known about circulation control and its benefits that it will be ready for production.

The intent of circulation control is to take the place of slats and flaps that most all planes use for takeoff and landing. Flaps and slats greatly increase the effective lift of an airplane at low speeds; however, while doing so they greatly increase the amount of drag the aircraft experiences. Circulation control is able to increase the lift coefficient while not adding any extra drag. Scaled wind tunnel tests have been conducted using circulation control with a Boeing 737. These tests have shown an increase in the lift coefficient of 85%, a 65% reduction in landing ground roll, a 60% reduction in takeoff ground roll, a 35% reduction of power on approach speed, and a 30% reduction in lift-off speed [6].

[pic]

Figure 17: Schematic of Circulation Control [6]

4.4.3. Co-Flow Jet Flow Control

Another high lift device that is being considered by Daedalus Aviation is co-flow jet flow control. Co-flow jet flow control, like the other high lift devices, increasing the lift coefficient of the airplane allowing for shorter takeoff and landing distances. The device, seen in Figure 18, works by blowing a high pressure stream of air tangentially out of the leading edge of the wing over the suction surface. The same amount of air is then sucked in near the trailing edge of the wing. This technology has been proven effective in wind tunnel, and the year 2058, co-flow jet flow control should be in use.

Test results show outstanding lift improvements for this technology. The maximum coefficient of lift was increased from 1.57 to 5.04; this is an increase of 220% for the NACA0025 airfoil. Along with an increase in the lift coefficient there was also a great increase in the stall angle of attack from 19° to 44°, this is an increase of 153% for the stall angle of attack [7]. A high angle of attack can be uncomfortable for passengers therefore this technology would not be used to climb at high angle, rather it would be used as a safety feature, allowing the plane to safely operate at a higher angle of attack if needed.

[pic]

Figure 18: Schematic of Co-Flow Jet Flow [7]

4.4. Technology Readiness Level Summary

The current 2008 technology readiness levels for these technologies is shown below in Table 3.

|Type |Description |TRL |

|Weight/Cost Savings |Composites |9 |

| |UAV/AI Pilot |6 |

|Propulsion Type |Pulse Detonation |3 |

| |Geared Turbofans |6 |

|Propulsion Enhancement |Magnetic Bearings |3 |

| |Thrust Vectoring |7 |

|High Lift |Circulation Control |7 |

| |Blown Flaps |9 |

| |Co Flow Jet Control |4 |

Table 3: Summary of Technology Readiness Levels.

5. Constraint Analysis and Sizing

5.1. Performance Constraints

To better understand how the aircraft mission constraints are impacting the design space, a constraint analysis was performed. The power available and the power required to perform under certain flight conditions, such as climbing, accelerating, or steady-level flight were analyzed to produce functions relating thrust-to-weight ratio and wing loading. The general equation produced is

[pic]. (1)

This master constraint equation was used to plot each of the mission requirements, narrowing the feasible design region. The nomenclature for the constraint analysis is given in Table 4.

|Symbol |Description |Value [Dimensions] |

|[pic] |Thrust-to-Weight Ratio |[non-dimensional] |

|[pic] |Mission Segment Weight Ratio |[non-dimensional] |

|[pic] |Thrust Lapse Ratio |[non-dimensional] |

|[pic] |Parasite Drag Coefficient |0.015 [non-dimensional] |

|[pic] |Wing Loading |[psf] |

|e |Oswald Efficiency Factor |0.8 [non-dimensional] |

|AR |Aspect Ratio |10 [non-dimensional] |

|n |Load Factor |[non-dimensional] |

|q |Dynamic Pressure |[psf] |

|h |Altitude |[ft] |

|V |Flight Speed |[ft/s] |

|g |Acceleration Due to Gravity |32.2 [ft/s2] |

|CLmax |Maximum Lift Coefficient |4.0 [non-dimensional] |

|sTO |Takeoff Rolling Distance |2500 [ft] |

|sL |Landing Roll Distance |2500 [ft] |

|( |Rolling Friction Coefficient |0.5 [non-dimensional] |

|TOP |Takeoff Parameter |[psf] |

|L/D |Lift-to-Drag Ratio |[non-dimensional] |

|N |Number of Engines |[engines] |

|CGR |Climb Gradient |[%] |

|M |Mach Number |[non-dimensional] |

Table 4: Constraint analysis nomenclature.

5.1.1. Takeoff Roll

A critical constraint for the Daedalus One is the takeoff roll. Because operation from short runways is the primary mission, a takeoff roll constraint is prescribed. The target distance is 2500 ft, and the threshold is 3500 ft. For takeoff roll, the master constraint equation is not used, but rather,

[pic]. (2)

This assumes that thrust is much larger than the drag and rolling friction, and that the takeoff velocity is 10% above the stall speed. A standard atmosphere at 5000 ft with a temperature offset of +25° was used to simulate high-hot conditions. Variable parameters used in this analysis are summarized in Table 5.

|Parameter |Value |

|β |1 [non-dimensional] |

|( |0.82 [non-dimensional] |

|ρ |0.00195 [slug/ft3] |

|sTO |2500 [ft]—Target |

|sTO |3500 [ft]—Threshold |

Table 5: Summary of takeoff parameters.

5.1.2. Landing Roll

An approach similar to the takeoff roll was used for the landing roll. However, because engines are assumed to be at idle during landing, the constraint equation takes a slightly different form of

[pic]. (3)

This again assumes high-hot conditions, that landing occurs at 15% above stall, roll out on dry asphalt, and landing in an emergency shortly after takeoff with a nearly full fuel load. Variable parameters used in this analysis are summarized in Table 6.

|Parameter |Value |

|β |0.98 [non-dimensional] |

|ρ |0.00195 [slug/ft3] |

|sL |2500 [ft] |

Table 6: Summary of landing parameters.

5.1.3. Takeoff Field Length—One Engine Inoperative

The balanced field length (the distance to clear a 35 ft obstacle when one engine fails at the point where the pilot must decide to takeoff or abort) is estimated using an empirically based takeoff parameter [1]. The takeoff parameter is

[pic]. (4)

Based on a sampling of the field lengths of small airports, a target field length of 5000 ft was prescribed. The takeoff parameter required to achieve the 5000 ft field length given a certain number of engines was estimated from Raymer, and the results are summarized in Table 7.

|# Engines |Takeoff Parameter [psf] |

|2 |115 |

|3 |125 |

|4 |135 |

Table 7: Summary of takeoff parameter estimates.

Using the estimated takeoff parameter, the required thrust-to-weight ratio was found as

[pic]. (5)

The thrust lapse ratio and maximum coefficient of lift use were the same as those used for the takeoff roll analysis. Raymer also correlates the takeoff roll to the takeoff parameter, and this method agrees well with the takeoff analysis discussed previously.

5.1.4. Second Segment Climb Rate—One Engine Inoperative

A second requirement that must be met with one engine inoperative is the second segment climb after the aircraft has cleared a 35 ft obstacle. FARs require a certain climb gradient with the landing gear up and high-lift devices in takeoff configuration. The required climb gradients are summarized in Table 8.

|# Engines |Climb Gradient [%] |

|2 |3.0 |

|3 |2.7 |

|4 |2.4 |

Table 8: Summary of required second-segment climb gradients.

The required climb gradient was applied to the constraint analysis by

[pic]. (6)

The lift-to drag ratio was estimated as 11.5 using flow-controlled high lift devices by

[pic]. (7)

5.1.5. Climb Rate at Top of Climb

To ensure the ability to climb at a reasonable rate at the initial cruising altitude, a climb rate of 500 ft/min is prescribed at 36,000 ft. This constraint was applied using the master equation (1). The non-zero variable terms are summarized in Table 9.

|Parameter |Value |

|β |0.98 [non-dimensional] |

|( |0.3 [non-dimensional] |

|q |187.4 [psf] |

|M |0.75 [non-dimensional] |

|dh/dt |8.33 [ft/s] |

|n |1 [non-dimensional] |

Table 9: Summary of top of climb parameters.

5.1.6. Climb Rate at Service Ceiling

FARs require that at the aircraft service ceiling a climb rate of 100 ft/min is attainable. The Daedalus One will be capable of achieving a service ceiling of 41,000 ft, which is typical for civil transport jets. This constraint uses the master equation (1) and the non-zero variable terms are summarized in Table 10.

|Parameter |Value |

|β |0.98 [non-dimensional] |

|( |0.24 [non-dimensional] |

|q |147.6 [psf] |

|M |0.75 [non-dimensional] |

|dh/dt |1.67 [ft/s] |

|n |1 [non-dimensional] |

Table 10: Summary of service ceiling parameters.

5.1.7. 2G Maneuver

In the event that the Daedalus One would have to perform a sudden maneuver while cruising, a 2G maneuver at 36,000 ft is prescribed. This would simulate making a turn at the initial cruising altitude with a tighter turning radius than under normal operations. This constraint is modeled using the master equation (1) with an increased load factor. The non-zero variable terms are summarized in Table 11.

|Parameter |Value |

|β |0.98 [non-dimensional] |

|( |0.24 [non-dimensional] |

|q |187.4 [psf] |

|M |0.75 [non-dimensional] |

|n |1 [non-dimensional] |

Table 11: Summary of 2G maneuver parameters.

5.1.8. Steady, Level Cruise

Lastly, to ensure the aircraft is capable of maintaining steady, level flight at altitude, the master equation (1) is used to model the cruise condition. The Daedalus One is modeled to cruise at mach 0.75 between 36,000 ft and 41,000 ft. A summary of the non-zero variable terms is presented in Table 12. This initial analysis neglects wave drag, which could be significant in this transonic regime.

|Parameter |Value |

|β |0.98 [non-dimensional] |

|( |0.24-0.3 [non-dimensional] |

|q |147.6-187.4 [psf] |

|M |0.75 [non-dimensional] |

|n |1 [non-dimensional] |

Table 12: Summary of cruise parameters.

5.2. Constraint Analysis Results

The results of the constraint analysis are shown in Figure 19. The current configuration of the Daedalus One has two engines. The driving constraints for a two engine aircraft are takeoff field length and second segment climb, both with one engine inoperative. According to the constraint diagram, the optimal initial design point for a two engine aircraft has a thrust-to-weight ratio of 0.23 and a wing loading of 84 psf. These values are both low compared to 2008 aircraft of similar size. However, the low wing loading could be explained by the ESTOL capability and the need for a larger wing with lower loading. The low thrust-to-weight ratio is likely an artifact of the takeoff field analysis using the takeoff parameter. Because of the low wing loading for ESTOL, a smaller thrust-to-weight ratio can be used to achieve the same takeoff parameter. This could suggest that the takeoff parameter estimates need to be adjusted to account for ESTOL in future analyses.

[pic]

Figure 19: Constraint analysis results.

5.3 Sizing

5.3.1. Sizing Software Selection

The sizing software selected by Daedalus Aviation is RDS [8]. This selection was made due to a unique feature the software provides. RDS has a design module in which the user actually designs the aircraft. This module was found be extremely helpful due to the unique design of Daedalus One. The other sizing codes available do not contain this feature, a lack of which could potentially cause inaccuracies in sizing due to the programs nature to assume a conventional design.

After completion of the aircraft model, the geometric data is then used for the aerodynamic, weights and propulsion analysis. Each analysis is run using the geometric data and will use it to create its respective data. After these analyses are run, the data is compiled into one master input file. In addition to the input file, a mission profile must also be created in order to size to the particular mission. The user must also input data for this master input file. The inputs from the user include preliminary values for the weight, wing loading and thrust to weight ratio. The program will perform iterations around the estimated weight to find a more accurate weight for the design. RDS will change some input values in order to allow the aircraft to perform the given mission profile.

5.3.2 RDS Sizing Method

As required by the RDS software, a model of Daedalus One was created to approximate specifications. These dimensions were obtained from the cabin layout (Section 3.2) as well as from the historical database. A general layout of the design was created for this initial stage of sizing. The model is not a fully accurate design; this decision was made due to not all aspects of the design being finalized. Upon completion of the model the approximate locations of extra weights were added to the model. These weights include the cockpit crew, avionics, APU, fuel and payload. This data was then saved to the geometric data file. The geometric data was then transferred to the different analysis modules.

The first analysis is the aerodynamic analysis. RDS will estimate the parasite, induced drags, lift curve slope and maximum lift for both subsonic and supersonic speeds. These values are computed using a user defined matrix including values for max altitude, max velocity and cruise lift coefficient as well as geometric data from the model. The data obtained from this analysis is then sent to a master data file.

The second analysis module is the weights. This analysis finds the weights of specific components as well as the center of gravity for the aircraft. These values are obtained by the user entering the class of aircraft, transport in this case, as well as several other inputs including estimated gross takeoff and empty weights. This data is then sent to the master data file as well.

The final analysis module is the propulsion analysis. This module allows for input from an already designed manufactures engine as well as an option to use a “rubber” engine. The analysis will then calculate the thrust and specific fuel consumption performance of the engines while in use on the aircraft. Like the other two analysis modules the data will be sent to the master data file.

Upon completion of the three analysis modules a mission profile must be created. The mission profile will include the individual stages during a mission, which must include points such as take-off, cruise, landing etc. Specific data must be entered in each mission point such as the cruise speed and altitude. The user must then finalize the master input data file by including values such as the weight of the crew, payload, weight coefficient and wing loading.

When this input file is competed the aircraft may be sized. The results from the analysis will include the mission segment weight fraction, fuel segment burn, cruise speeds, ranges, and the fuel, empty and total gross weight of the aircraft.

5.3.3. Sizing Inputs and Results

The model of Daedalus One was made with specific inputs from a historical database as well as the cabin mockup. These inputs can be seen in Table 13. The historical database values were used due to the nature of most aircraft of Daedalus One’s class having those respective values.

|Fuselage Length |100 ft |

|# Passengers |102 |

|# Crew |4 |

|Aspect Ratio |10 |

|Wing Sweep |10° |

|Wing Loading |84 lb/ft2 |

|Thrust to Weight Ratio |0.23 |

|Max Gross Takeoff Weight |100,000 lb |

|Estimate | |

|Empty Weight Estimate |80,000 lb |

|Max CL |4.0 |

|Max Cruise Speed |0.75 M |

|Max Cruise Altitude |42,000 ft |

Table 13: Sizing Inputs

Some of these inputs, such as the fuselage length and aspect ratio were used in the model creation; others such as the max gross takeoff weight estimate were used in the analysis data files. The values used are preliminary values and do not reflect the final sizing results of Daedalus One. These values will likely change as iterations continue and final decisions are made. Technology savings was used for the crew number input data. Since Daedalus Aviation seeks to replace the pilot with an artificial intelligence, one member of the flight crew was subtracted from the total number. This subtraction does not affect the number of flight attendants, but rather the personnel in the cockpit.

Using these values the sizing analysis was run using the method described in Section 5.3.2. The results from this analysis can be found in Table 14.

|Max Gross Takeoff Weight |88,000 lb |

|Empty Weight |52,800 lb |

|Range |1,800 nmi |

|We/Wo |0.6 |

|CG Range |65-70 ft |

|(Aft of Nose) | |

Table 14: Sizing Outputs

Like the inputs, these results are not the final and will change during the many design iterations to follow this stage. These results neglect any technology which may add or subtract weight from the final design. These savings have been neglected due to the fact that the current analysis is only preliminary. Savings will be included after the final sizing analysis has been performed.

The results from this preliminary sizing were checked against the historical database values as a “sanity” check to insure that the RDS software was providing reasonable values for the aircrafts size. The following charts display the historical database and trend lines as well as where RDS currently places Daedalus One.

[pic]Figure 20: Database Plot of Range vs. Take-off Weight

[pic]Figure 21: Database Plot of Take-Off Distance vs. Take-off Weight

[pic] Figure 22: Database Plot of Take-Off Weight vs. Passenger Capacity

As can be seen in Figures 20, 21, and 22 the current sizing analysis places Daedalus One approximately on par with the historical database. More work must be done on the RDS model to allow for a more accurate sizing.

6. Conclusion

The Daedalus One is a 102-passenger commuter jet, with ESTOL capabilities. It is designed to be used at smaller airports throughout the country, in an effort to alleviate the traffic and congestion at larger airports. It has a range of 1800 nautical miles, which gives it the ability to cross a large portion of the continental United States, along with access to parts of South America. With its use of technology to increase performance and reduce weight, it will be a strong competitor in the medium-range, medium-capacity commuter jet market.

6.1 Walk Around Diagram

[pic]

Figure 23: Walk around diagram of Daedalus One current configuration.

6.2 System Compliance Matrix

As shown in Table 15, the current configuration of the Daedalus One meets or exceeds all of the design requirement targets or thresholds. The fuel burn currently seems exceedingly low. Further investigation will be performed to determine the accuracy of this value.

[pic]

Table 15: System compliance matrix.

6.3. Next Steps

Several more steps need to be taken in order to complete the final design of the Daedalus One. The final advanced technology devices that the aircraft will utilize, including high lift, propulsion and aerodynamic devices need to be selected. Improvements made by the advanced technologies then need to be integrated into the sizing analysis which will help with the creation of carpet plots. The carpet plot will then be used to choose the design point. Advanced aerodynamic studies also need to be completed on the Daedalus One along with final airfoil selections for the wings and tail. Propulsion and performance values of the aircraft need to be studied and calculated. Development of the aircraft structure needs to take place accounting for important load paths, wing-fuselage intersection, motor mounts, and landing gear integration. The center of gravity and neutral point need to be found in order for the static margin to be calculated. Finally, cost analysis including development and production costs needs to be researched along with studies on the environmental impact of the aircraft need to be conducted.

References

1 Raymer, Daniel P. (2006) Aircraft Design: A Conceptual Approach. 4ed. Reston, Virginia: American Institute of Aeronautics and Astronautics, Inc.

2. Sun, C.T. Personal Interview. 13 Jan. 2008.

3. AIN Online. “GE Convinced PDE Holds Game-Changing Potential” [Online] Available (February 21, 2008)

4. Pratt-Whitney. “Pratt & Whitney’s Geared Turbofan Demonstrator Engine Completes Phase I Ground Tests Ahead of Schedule” [Online] Available vgn-ext-templating/v/index.jsp?vgnextoid=2e35288d1c83c010VgnVCM1000000881000 aRCRD&prid=37df431e0b418110VgnVCM100000c45a529f____ (February 14, 2008)

5. W. H. Mason “Some High Lift Aerodynamics Part 2 Powered Lift Systems” Virginia Tech University 1998

6. Nichols, J. H., Englar, R. J. “Advanced Circulation Control Wing System for Navy STOL Aircraft” AIAA Paper 80-1825R

7. Zha, G., Carroll B. F., Paxton, C. D., Conley, C. A., Wells, Adam “High Performance Airfoil Using Co-Flow Jet Flow Control” AIAA Paper 2005-1260 January 2005

8. Raymer, Daniel P. (2006) RDS-Student Software for Aircraft Design, Sizing, and Performance. Reston, Virginia: American Institute of Aeronautics and Astronautics, Inc.

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Daedalus One

Daedalus One

Daedalus One

Taxi

Takeoff &

Climb

Step Cruise

For Best Range

Descend & Hold

Land & Taxi

Climb- Miss

Approach

Cruise

Descend & Hold

Land & Taxi

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