System Definitions Review. Optoprime Conceptual Designs, LLC.

Transcription

1 YDC- NEXTGEN System Definitions Review Team 2 AJ Berger Colby Darlage Joshua Dias Ahmad Kamaruddin Pete Krupski Josh Mason Camrand Tucker Page of 4

2 TABLE OF CONTENTS MISSION STATEMENT... 4 INTRODUCTION... 4 USE CASE SCENARIOS... 5 Scenario #... 5 Scenario # Scenario #... 7 MAJOR DESIGN REQUIREMENTS... 7 AIRCRAFT CONCEPT SELECTION... 8 Process... 8 Outcome... 2 ADVANCED TECHNOLOGIES... Technology Readiness Levels... Carbon-Fiber Reinforced Plastic (CFRP)... 4 Ceramic Matrix Composites (CMCs)... 4 Glass-Reinforced Fiber Metal Laminate (GLARE)... 5 Central Reinforced Aluminum (CentrAl)... 5 Fuel... 6 Power Plant... 6 Wave Rotor Combustion... 7 Solar Energy... 8 Preliminary Engine Design... FUSELAGE & CABIN LAYOUT... CONSTRAINT ANALYSIS... 2 Major Performance Constraints... 2 Basic Assumptions Constraint Diagram Based on Assumptions SIZING STUDY Design Mission Sizing Method Trade Studies Carpet Plots SUMMARY OF CONCEPT... Overall Concept... Requirements Compliance... Next Steps... 2 APPENDIX A House of Quality... REFERENCES... 4 Page 2 of 4

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4 MISSION STATEMENT To satisfy our customers through the design of an advanced mid-range aircraft capable of relieving congestion at major hubs throughout the world. The aircraft will: Operate from lesser-equipped airports throughout the world, Maintain a high cruise speed while limiting negative impact on the environment, Satisfy customer needs without sacrificing safety. INTRODUCTION Having defined the System Requirements, Optoprime Conceptual Designs specified the System Definitions for an aircraft that will reduce network congestion within the United States as well as develop growing travel industries worldwide. The voice of the customer was heard through contact with airline management as well as market forecasts from major commercial airframers, including Airbus, Boeing and Bombardier Aerospace. The team used a House-of-Quality (HoQ) to determine the importance of the requirements, which would, in turn, guide design. Page 4 of 4

5 USE CASE SCENARIOS In these use case scenarios, we are describing possible missions our aircraft will be able to perform. These city pairs and the distances are representative of mission profiles that could take place between other cities as well. The routes are not limited to these areas of operations, but could possibly include city pairs throughout the world. It is important to note that all of these use cases have a built in fuel reserve and that in Scenario that reserve is specifically used and pointed out. Scenario # The world economy is booming and it is not practical to limit one s economic interests by country borders. Chinese investors, who are a part of the fastest growing economy in the world, have expanded their operations into technologically growing India. They are working closely with their Indian investment and this requires them to take frequent and regular trips from Hong Kong to Madras. From Hong Kong International, they depart using only part of the international runway utilizing simultaneous operations on one runway. They then set a direct course for Madras. The long haul ends when the aircraft lands on the 6 ft runway at Madras. There is however one problem, the passengers seem reluctant to leave. They don t want to trade the comfortable plane for a cramped limo. Figure : Scenario Mission Profile Page 5 of 4

6 Scenario #2 The school year Down Under is just over and it is time for vacation or holiday. Many students in Sydney desire to travel across the continent to Perth so they can forget about the rigors of the past school year. They board our aircraft in Sydney and fly the 8 NM to Perth. In Perth the aircraft is reconfigured and loaded with cargo instead of passengers. The aircraft then travels NM to the middle of the outback to the small mining town of Coober Pedy where cargo is transferred. Coober Pedy has a short runway and limited facilities but that is what our aircraft is designed for. Unfortunately, Coober Pedy does not have any fuel available, but that does not pose a problem. Our aircraft does not need to refuel to travel the NM back to Sydney where it is converted back to a passenger aircraft. Our aircraft will have the quick change cargo feature which will allow it to maximize revenue by allowing passenger use during peak times and cargo operations during the down times. Passenger seats could be removed and stored at multiple locations or carried onboard in a cargo container and then refitted when the cargo haul was finished. Figure 2: Scenario 2 Mission Profile Page 6 of 4

7 Scenario # Chicago O Hare International Airport is overly congested so Gary/Chicago International is used to relieve some of that congestion. Our aircraft departs on the ft runway at Gary/Chicago traveling 8 NM to Boulder. Denver International like O Hare is overly congested so the airport at Boulder is used to relieve some of the congestion at Denver. Shortly before landing at Boulder the cross winds pick up and the plane is forced into a holding pattern to wait for the weather to clear. Snow begins to fall and the aircraft needs to find a new airport to wait for the snowstorm to clear. The flight is diverted over the mountains to Durango 2 NM away. After the storm clears the plane takes off and flies back to Boulder and lands on the 4 ft runway without needing to take on fuel at Durango. Figure : Scenario Mission Profile MAJOR DESIGN REQUIREMENTS Concept selection is a difficult process and it is necessary to have a good tool for selecting a single concept for a long list of concepts. From the HoQ, we knew what we generally were looking for in a concept. However, the team generated a list of priorities to aid in the concept evaluation and subsequent selection phase. Pugh s Method is a -step design tool created for the purpose of selecting one concept from a list of concepts by eliminating designs, creating new designs, and/or creating hybrid designs. This process is simplified here, for clarity. The first step in Pugh s Method is to come up with a list of requirements. To start out, we had group members present ideas for requirements and we wrote them on a dry-erase board. Many of these requirements came from our group s House of Quality. We then wrote each requirement on a small piece of paper and laid them on a table. We then began the process of grouping the pieces of paper in categories. Page 7 of 4

8 For this process group members did not speak; anyone was allowed to move the papers around the table even if they had already been moved by someone else. If we noticed that a requirement was being moved back and forth between two groups, the team made a copy of that requirement and allowed it to coexist. After the general movement had ceased, the team discussed these categories. Some were eliminated, instead using a single requirement title. We continued this process until everyone was satisfied with the results and we had our list of requirements to use for the concept selection. Table, below, lists the resulting requirements from this process. Table : Design Requirements for Pugh s Method Design Requirements Passenger Climate/Comfort Short Runway Range Gate Time Easy Maintenance Low Noise Energy Efficiency Limited Terminal Service Obstacle Clearance Crew Cost AIRCRAFT CONCEPT SELECTION Process The second step in Pugh s Method is to come up with a list of concepts. The way we accomplished this was by having each group member come up with one design. We tried to focus on having each concept have only one (or a small number) distinguishing characteristic because we felt that this would better facilitate the process of coming up with a hybrid design. This section will go through the list of concepts the group came up with and will explain what the distinguishing characteristic(s) of each design are. Refer Figure 4 for sketches of each concept. The first concept was the Electric which is a low wing aircraft with electric powered propeller motors mounted on top of the wing. The second concept was the Blended Wing which has a blended wing body with engines mounted inside the fuselage. The Joined Wing concept is characterized by a tandem wing joined at the tips with pusher open rotor engines on the rear wing. The Dual Boom has a short forward mounted fuselage with a high wing and dual tail booms. The Biplane has two identical wings mounted above and below the fuselage with the engines mounted in between the wings. The Solar-Powered concept has solar cells on top of a high wing with 2 electric tractortype motors. The Dual Fuselage concept is characterized by two identical fuselages mounted side-by-side and joined by a high wing. Now that we had our list of requirements and our list of concepts defined we were able to create a matrix. Page 8 of 4

9 Electric Blended Wing Joined Wing Dual Boom Biplane Solar Powered Dual Fuselage Figure 4: Sketches of Considered Concepts For Pugh s Method we listed the requirements in a column down the left side of the matrix and the list of concepts in a row across the top (For simplification purposes we gave each concept a nickname by which to remember it). To begin the process of Pugh s Method we needed to select a datum to which we would compare all of the other designs. For the first time through Pugh s Method we used a standard tube & wing design to be our datum. We then went through line-by-line and rated each concept as better (+), worse (-), or the same (S) as the datum for each requirement. After we completed populating the matrix we summed the number of + s, - s, and S s for each column. We then went through each design and changed aspects of the concept where needed, or better defined an aspect of the design which raised questions in the first iteration. See Figure 5 for the spreadsheet of the first iteration. Page of 4

10 ITERATION # Design Criteria Standard Solar Electric Biplane Dual Fuse Blended Joined Dual Boom PAX Climate/Comfort + + S + S S Short Runway + S Range + S Gate Time + S S + S Easy Maintenance S S Low Noise S + Energy Efficiency Limited Terminal Service + S S S S Obstacle Clearance + S Crew Cost S S S S S S S Figure 5: Iteration # of Pugh's Method We then looked to see which design had the most + s for the least - s and said this concept had the best result from the first iteration and would be selected as the datum in the second iteration. We then repeated this process using the new datum. See Figure 6 and Figure 7 for the remaining iterations. We continued to iterate until we felt that we had exhausted the process and we had found a direction to work toward. Page of 4

11 ITERATION #2 Design Criteria Standard Solar Electric Biplane Dual Fuse Blended Joined Dual Boom PAX Climate/Comfort S S S + S Short Runway + S + + S + Range Gate Time Easy Maintenance + S S S Low Noise + + S Energy Efficiency Limited Terminal Service S S S S S S Obstacle Clearance S S Crew Cost S S S S S S + 2 S Figure 6: Iteration #2 of Pugh's Method ITERATION # Design Criteria SolarElectricBiplane Dual Fuse Blended Joined Dual Boom Swing Wing PAX Climate/Comfort Short Runway S + Range + Gate Time + + S S S S Easy Maintenance S Low Noise Energy Efficiency + + S + Limited Terminal Service + S S S Obstacle Clearance S Crew Cost S S S S S S S Figure 7: Iteration # of Pugh's Method Since Pugh s Method is not a quantitative tool concerned with mathematically eliminating concepts and is instead concerned with creating group discussion and hybrid designs; we decided that the two most feasible concepts are the Joined Wing and the Dual Fuselage. Although concepts such as the Solar-Powered and the Blended Page of 4

12 Wing seemed to do well in the Pugh s Method, the team determined that a solarpowered transport aircraft is not feasible, even given the extended development time frame. However, the blended wing design, although feasible, lacked the static stability required for the safety of commercial passengers in flight in the unlikely event of a control system failure. Outcome We decided as a group that we should do further research into the joined wing and the dual fuselage design since they were the most feasible given our mission. After doing some research, we determined that a dual fuselage design is beneficial for very large aircraft because one large fuselage is replaced by two slightly smaller fuselages. If our passenger aircraft was replaced by two even smaller fuselages, passengers might have to crawl to get to their seats. Also, we found that a dual fuselage design typically has longer takeoff and landing distances than a similar-sized single fuselage aircraft (due to higher profile drag). This makes this design even less desirable for our mission. A joined wing design has many aspects which make it desirable for our mission which will be discussed later in this document. This is the concept Optoprime has elected to move forward with. Page 2 of 4

13 ADVANCED TECHNOLOGIES Technology Readiness Levels In our concept selection process we investigated several advanced technologies. These technologies are at different stages of development and thus the readiness of them is at different stages. Since the readiness of a fuel is measured differently than that of an engine, it is beneficial to have a standard way of evaluating the readiness of all advanced technologies. One method, the method we employed, is the Technology Readiness Level created by NASA. The NASA Technology Readiness Levels are a quantitative method of evaluating qualitative characteristics but it is not limited to them. The readiness level of a technology is rated on a scale from to with whole numbers in between, as shown in Figure 8. The description of each level can be found below: Figure 8: NASA Technology Readiness Levels TRL - Basic principles observed and reported TRL 2- Technology concept and/or application formulated TRL - Analytical and experimental critical function and/or characteristic proof-ofconcept TRL 4- Component and/or breadboard validation in laboratory environment TRL 5- Component and/or breadboard validation in relevant environment TRL 6- System/subsystem model or prototype demonstration in a relevant environment (ground or space) TRL 7- System prototype demonstration in a space environment TRL 8- Actual system completed and flight qualified through test and demonstration (ground of space) TRL - Actual system flight proven through successful mission operations Page of 4

14 The TRLs are simply a scale to represent how ready a technology is for implementation into an operable system. Throughout this paper when a new technology is introduced it will be accompanied by a TRL rating. With representing widespread use and is not ready. Carbon-Fiber Reinforced Plastic (CFRP) The plastic used for CFRP is most often epoxy, polyester or nylon. CFRP is a very strong, light and expensive composite. It is much more durable compare to aluminum, has very high resistance to cracking and using CFRP on the aircraft body will allow higher cabin pressure during flight. Furthermore, CFRP is relatively insensitive to flaws. Fatigue testing of CFRP structures demonstrated its high resistance to cracking and that fractures generally do not propagate. Consequently, this would reduce the maintenance cost and time due to the result of a reduced risk corrosion and fatigue of CFRP compared to metals. CFRP has lower density, greater strength and stiffness than aluminum, thus a smaller lighter structure can carry the same load. CFRP has already been used on Boeing 787 and Airbus A8 aircraft, mostly for the fuselage and wings. Therefore, CFRP is on level for the Technology Readiness Level (TRL) by NASA. CFRP also can be used for the tail surfaces and doors. Based on the research conducted by Boeing, 8% composite structural weight can result in a 4% reduction in empty weight, % reduction in wing area, and % fuel saving for the same mission profile comparing to the aircraft built in metal structure. The concept is, with the reduction in weight, it will reduce the wing area as well since a smaller wing area can generate the same lift that is needed to fly the aircraft. This also can be said for the fuel saving, since with the weight reduction, the fuel consumption also can be reduced because heavier aircraft will to run engines at a higher level to remain aloft. Ceramic Matrix Composites (CMCs) The idea of developing Ceramics Matrix Composites (CMCs) is to design lightweight, high temperature composite materials for engine applications in order to reduce weight and fuel consumption while increasing the fuel efficiency. CMCs are currently developed by the Advanced High Temperature Engine Materials Program (HITEMP) conducted by NASA. As NASA has recently conducted tests, CMCs are on level 5 for the Technology Readiness Level (TRL). The CMCs that can be produced are strong and tough but only for short times at very high temperature. CMCs are made of bone (hydroxyapatite reinforced with collagen), cermet (ceramic and metal) and concrete. Ceramic matrix composites are built primarily for toughness, not for strength. CMCs have lower density, better oxidation resistance and potential to operate at significantly higher temperature than super alloys. CMCs can withstand temperatures as high as 65 C. CMCs can be used in hot engine shrouds and components such as combustors, Page 4 of 4

15 turbines and exhaust nozzles. With higher operating temperature, CMCs will produce greater combustion efficiency and consequently, reduce the fuel consumption. Moreover, with the low density of CMCs, the use of CMCs in the hot section of the engine can result in a 5% reduction in engine weight as compared with current technology. That also means, for an aircraft with four engines, 4% of the aircraft weight can be reduced thus contributing to lower initial costs and operating costs. Glass-Reinforced Fiber Metal Laminate (GLARE) Glass-Reinforced Fiber Metal Laminate (GLARE) is composed of several very thin layers of metal, usually aluminum, interspersed with layers of glass-fiber, bonded together with a matrix such as epoxy. GLARE is lighter with lower specific weight, has better corrosion, impact and fire resistance than conventional aluminum alloys used in aviation. These materials also can be repaired using conventional aluminum repair techniques. GLARE can be used for the fuselage, stabilizer and leading edges. GLARE is on level for the NASA TRL, as it is already been used on Airbus A8 for their upper fuselage and stabilizer s leading edge. GLARE can be tailored during design and manufactured such that the number, type and alignment of layers can suit the local stresses and shapes throughout the aircraft. This allows the production of double-curved sections (lofts), complex integrated panels or very large sheets. A structure properly designed for GLARE will be significantly lighter and less complex than an equivalent metal structure. In turn, this will require less inspection and maintenance and have much longer fatigue life, cheaper cost. Central Reinforced Aluminum (CentrAl) Central Reinforced Aluminum (CentrAl) concept is a central layer of fiber metal laminates (VML), sandwiched between one or more thick layers of high-quality aluminum. This yields a robust construction material which is not only extremely strong, but is also not susceptible to fatigue. In contrast to CFRP constructions, this application still enables simple repairs to be performed, same as aluminum constructions which can lower the manufacturing and maintenance costs. The new high-quality CentrAl constructions are stronger than the CFRP constructions recently applied in the wings of the Boeing 787. Using CentrAl as the wing constructions, weight could be reduced by a further 2% compared to CFRP wing constructions. With the reduction in weight and higher strength than CFRP, using CentrAl as the main composite material for an aircraft will lead to reduction in maintenance costs, fuel consumption and make the aircraft the truly energy-efficient green aircraft. CentrAl is on level 4 for the TRL by NASA as it is still been researched by the Faculty of Aerospace Engineering of Delft University of Technology in Netherlands with collaboration by GTM Advanced Structures company,which specializes in aircraft materials. Page 5 of 4

16 Fuel The YDC-NG will run on synthetic hydrocarbon-based jet fuel. The Fischer-Tropsch process extracts the hydrogen, carbon, and oxygen needed to produce hydrocarbon fuels from other sources. The extracted materials are then combined to produce liquid hydrocarbon fuels. To address environmental issues we will use a new process called Green Freedom which has been developed by Los Alamos Laboratories. 2 This process uses the fact that the only emission from nuclear power plants is steam. Green Freedom then extracts water from the steam and carbon dioxide from the air. The water and carbon dioxide are then dissociated into their individual elements, which are then used as the building blocks to make liquid hydrocarbon fuels using the Fischer-Tropsch process. This process can produce fuels with characteristics that are very similar to those of existing jet fuels, such as Jet A, Jet A- and Jet B. The synthesized fuel will have a weight and heating value similar to that of Jet-A fuel. Theoretically, we may be able to formulate a new type of jet fuel with even higher heating values than existing fossil fuels. The fuel, synthesized by the combination of the Fischer-Tropsch process and the Green Freedom process, is a carbon-neutral fuel. Burning the fuel does produce carbon dioxide as a byproduct. The carbon dioxide that is produced due to burning equals the amount of carbon dioxide that was used to synthesize the fuel. Since the carbon used by the Fischer-Tropsch process to make the fuel was extracted from the air using the Green Freedom process, no extra carbon is put into the atmosphere when the fuel is burned. Also the power used to make the fuel comes from the nuclear power-plant; therefore the fuel is virtually free of harmful emissions. Power Plant The engine used to power the aircraft is an Ultra High Bypass Propfan engine shown below in Figure. The reason for this choice of engine is because of its high efficiency and thrust capabilities. Compared to traditional turbofan engines currently used to power aircraft, the Ultra High Bypass Propfan engine can achieve thirty to forty percent lower specific fuel consumption than a traditional turbofan. The Propfan engine suffers negligible losses in power with this increased efficiency. It can still achieve speeds comparable to turbofans. Variable pitch external rotor blades can increase the efficiency of the engines. This allows for the maximum flow velocity. Page 6 of 4

17 Figure : Ultra High Bypass Propfan 4 One consideration is the large volume of noise produced by the open rotor of the engine. The use of composites in the construction of the aircraft fuselage will help to dampen the noise transmitted into the cabin. The location of the engines will also help to minimize the amount of noise transmitted to the cabin. The open rotor of the engine is positioned towards the aft portion of the fuselage. This will allow a large portion of the noise to be transmitted into the atmosphere behind the aircraft instead of into the cabin. The combination of composite materials and engine location will reduce the level of cabin noise from the open rotors of the engine. Other than critical system redundancy, the FAA does not specifically regulate against rotor burst. However, with safety in mind, the engines are located aft of the fuselage. If a rotor burst does occur, the offending blade will be ejected behind the aircraft. This eliminates the chance that a blade could be thrown into the fuselage. The engine will have a counter-rotating open rotor configuration. The efficiency of the engine is increased by an additional six to ten percent when compared to a single rotor configuration. The dual rotor configuration also increases the reverse thrust capability up to sixty percent of that of the takeoff thrust. A Wave Rotor Combustion System will replace the traditional combustion chamber. Wave Rotor Combustion A trade study on the wave rotor constant volume combustion (CVC) concept revealed several advantages that can be incorporated by our team into the propulsion concept. In a conventional gas turbine engine combustor, the combustion takes place at essentially constant pressure while the temperature rises. However, according to Won and Waters, the CVC creates a pressure rise in addition to the temperature rise, which results in the potential for greater work extraction by the turbine or higher pressure available at the exhaust nozzle to produce thrust. 5 Shown below in Figure is an exploded view of the CVC concept. Page 7 of 4

18 Figure : CVC Exploded View To obtain constant volume combustion, a rotor with many individual chambers is used. The rotor revolves while the two endplates remain stationary. In the figure above, air flows from left to right and the rotor is shown in blue. Air and fuel fill a chamber and is compressed by the upstream flow as it passes by an opening in the inlet endplate. As the rotor revolves, the chamber passes the opening and is then a sealed container. The airfuel mixture is then ignited, and the combustion increases the pressure and temperature in the chamber. As the rotor then rotates past an opening in the exit endplate, the combusted gas exhausts out of the chamber. A large advantage of combusting at constant volume is decreased SFC. In a simulation run by Won and Waters where a CVC was theoretically used aboard a 5 passenger Embraer CRJ-45 powered by twin Rolls-Royce AE7 high bypass turbofan engines for an 8 nm mission, a % benefit in mission fuel burn is seen which in turn drives a % reduction in DOC. This not only would benefit the customer, but also benefit the environment by reducing the NO x and CO 2 emissions from the lower fuel burn. Another advantage of using CVC is its impact on the compression system. By allowing a pressure rise in the combustor, less compression can be required out of the compressor. This could mean less complexity and fewer stages, which means a lower weight and less maintenance. Solar Energy The power that is diverted from the engine to supply the cabin and cockpit with electricity decreases the power supplied to propel the aircraft forward. To eliminate the need to divert power from the engine, the cabin and cockpit energy requirements will be powered largely by solar power. This will eliminate the need to carry extra fuel to help power the avionics and environmental control systems of the aircraft. The sun supplies an unlimited supply of power with no harmful emissions. Since no extra fuel is needed to power the Page 8 of 4

19 electrical needs of the aircraft, the operating cost of the aircraft will be decreased compared to an aircraft without solar power. This will also allow the aircraft to fly a longer distance on the same amount of fuel. Preliminary Engine Design The preliminary propulsion design consists of an Ultra High Bypass Propfan engine to increase the fuel efficiency of the engine. This engine will have a counter-rotating blade configuration to help further increase the efficiency of the engine. The combustion chamber will be the Wave Rotor combustion system. The fuel that the engine will use will be a fuel synthesized by the Fischer-Tropsch type process. The materials needed to make the fuel by this process will be obtained by the Green Freedom process to make the fuel carbon neutral. To further increase the efficiency of the engine, solar power will provide much of the power needed for the electrical system of the aircraft. FUSELAGE & CABIN LAYOUT The initial concept was for our aircraft to carry passengers in a single class. Thus, the goal became maximizing passenger comfort, while still allowing for enough cargo room and keeping a low cross-sectional area. The seating arrangement that best met these goals was a 2-aisle-2 arrangement. The size of the seats and seat pitch is comparable to an economy-plus or business class. The actual seat size and dimensions are located in Table 2. The larger seats should provide the passenger with more comfort and raise their confidence in our aircraft. We believe that larger seats will bring more passengers and more interest to our aircraft. Table 2: Seat Dimensions 6 First Class Optoprime Economy Seat width in Arm rest in Height off ground in Seat Pitch 8 4 in Overall Height in Two seats in The fuselage itself will be built in sections. This would allow the airline to order aircraft with more seating by inserting more passenger sections and making the over-all fuselage longer. The initial length of the passenger section would be ft. The overall length would be 22 ft. The passenger layout can be seen below in Figure. The overall dimensions are in Table : Fuselage Dimension. Page of 4

20 Emergency Exits Option for Cargo Door Lavatories Figure : Fuselage Layout Outer Diameter 6 in Inner Diameter 6.5 in Length 22 ft Passenger Volume (approx.) 6 ft^ Cargo Volume (approx.) 875 ft^ Volume per PAX (approx.) 6 ft^ Table : Fuselage Dimensions The design allows for the possibility of a large cargo door aft of the canard to avoid possible loading accidents. Also this allows for a more convenient emergency exit. The fuselage cross-section is still small to avoid pressure drag. The section diameter of the passenger cabin is 5.25ft. The height of the passenger cabin from the floor to the ceiling is 7.25 ft. The overhead bins are 5 ft from the floor. The aisle width is 2.5ft to allow for easy passage. Large overhead bins will allow for more carry-on storage. Figure 2 shows the section. Page 2 of 4

21 Figure 2: Section View The fuselage is also designed for a large cargo area under the passenger cabin. To allow for more room, the shape of the fuselage is not completely cylindrical. The resulting section can be seen in Figure. Figure : Section Cross-Section View The layout also features lavatories and a galley. The industry standard for an aircraft of this size is 2 lavatories and no galley. The layout also includes the possibility of an aft door that could contain stairs if the aircraft where to land at an airport/runway without a jet-way. CONSTRAINT ANALYSIS Major Performance Constraints The first major performance constraint that our team decided should be considered was takeoff. From the thrust-to-weight to wing loading relation for takeoff shown in Equation Page 2 of 4

22 (), the factors that most affect the takeoff constraint were determined to be fuel weight, takeoff altitude, takeoff distance, and C L,max. 2 2 W (.) β T S SL = W αgρc S L,max TO () The second major performance constraint that we decided to consider was a.5g constant rate turn. To give a more descriptive constraint diagram, we considered both a constant rate turn at top of climb (TOC) as well as at cruise. Both constraints are dependent on altitude, excess power, mach number, C D,, L/D max, and fuel fraction as can be seen in the thrust-to-weight to wing loading relation for constant rate turns in Equation (2). 2 TSL β q C D, nβ W dh dv = + W W + + α β πare q S V dt g dt S (2) The next major performance constraint that we decided to consider was landing. As can be seen in Equation (), the major factors that affect this constraint are C L,max, altitude, the landing surface coefficient of friction, landing distance, and fuel fraction. W S SLρCL,maxgμ = () 2 (.5) β Another major performance constraint we considered was second segment climb. This was important to consider because our concept has 2 engines and historically two-engine aircraft are sized by this constraint. As can be seen in the thrust-to-weight to wing loading relation for second segment climb in Equation (4), another important parameter to this scenario is L/D for the second segment climb portion of the flight. We approximated this in between a clean configuration with landing gear and all flaps fully collapsed and a dirty configuration with landing gear down and flaps extended. T W SL N = CGR + N ( L D) SS (4) Basic Assumptions Discussed in the following paragraphs are the basic assumptions that our team has made in creating our constraint diagram for our concept. These assumptions are also summarized at the end of this section in Table 4. Page 22 of 4

23 To give a good approximate thrust-to-weight to wing loading relation for our takeoff constraint of ft, all of the fuel was assumed to be on board at liftoff, despite the fact that fuel is burned as the airplane accelerates down the runway. Additionally, for a conservative estimate and to ensure the use case scenario involving Denver is feasible, a takeoff altitude of 5 ft was used. In the less-dense air, it will take longer to take off, so designing to perform a high altitude takeoff is crucial. We also estimated a C L,max value of 5 from preliminary estimates from current aircraft capabilities and technology predictions. Given that the C- can achieve a C L,max of.8, according to Roskam, and given our current configuration, we believe that a C L,max of 5 is feasible. To approximate the thrust-to-weight to wing loading relation for the constant rate turn at the top of climb, we used an altitude of, ft to correspond with our cruise-climb capability. Since it is early in the flight, we also chose a fuel fraction of.8 to represent that 8% of the fuel was still on board the airplane. We also chose to use an excess power of ft/min to correspond to flying at Mach.78 an altitude lower than our service ceiling of 5, feet. Finally, to determine a value for L/D max we used the value for the parasite drag coefficient, C D,, from Raymer of.5 for a clean jet. This, along with our aspect ratio and Oswald efficiency factor assumptions of 8 and.8 respectively, resulted in an L/D max value of 8.. Assumptions for the thrust-to-weight to wing loading relation for the constant rate turn at the cruise conditions were made at the service ceiling altitude of 5, ft. To go along with this, because it is the service ceiling, we assumed an excess power of ft/min. We also assumed a fuel fraction of. to represent the condition early in the flight. Regarding the other parameters affecting a.5g constant rate turn, we used the same values as in the top of climb condition. The next assumptions we made dealt with our landing constraint of landing within ft. We chose to evaluate this constraint using a dry concrete surface which has a coefficient of friction of. according to Raymer Table 7.. Similar to the takeoff constraint, we also chose to use a conservative altitude of 5 ft and C L,max of 5. Additionally, we noticed that in the constraint diagram the landing constraint is a vertical line independent of the thrust-to-weight ratio which shifts to the right as the fuel fraction decreases. Because of this, we chose a relatively high fuel fraction of.5 to show an extremely conservative landing scenario. To give a good approximate thrust-to-weight to wing loading relation for second segment climb, we chose to use a higher parasite drag coefficient than cruise because at this stage the landing gear will be assumed to be retracted but the flaps will still be in a takeoff configuration. As a result, we chose C D, for the second segment climb constraint to be.2 which corresponds to an L/D for second segment climb of 5., given our aspect ratio and Oswald efficiency factor. Finally, we used a climb gradient value of.24 based on Roskam s estimation for a twin-engine aircraft. Page 2 of 4

24 Constraint Table 4: Constraint Diagram Assumption Values Constant Constant Takeoff Rate Turn Rate Turn (TOC) (cruise) Landing Second Segment Climb Fuel Fraction, β Altitude, h [ft] 5, 5, 5 - C L,max Excess Power, dh/dt [ft/min] Mach Number C D, L/D - 8. (max) 8. (max) - 5. Landing Surface, μ (dry concrete) Climb Gradient, CGR Aspect Ratio, AR Oswald Efficiency Factor, e Distance (Takeoff/Landing), ft Constraint Diagram Based on Assumptions With the assumptions from the above discussion, we constructed a constraint diagram in order to determine preliminary values for thrust-to-weight and wingloading. Shown below in Figure 4 is the constraint diagram for our concept resulting from our preliminary assumptions Constraint Diagram Constant Rate Turn (TOC) Constant Rate Turn (cruise) Takeoff (β=) Landing (β=.5) Second Segment Industry Avg (.5,.425) Design Point (,.) T /W Wingloading [psf] Figure 4: Constraint Diagram Page 24 of 4

25 It can be seen in our constraint diagram that our concept is sized by our top of climb constant rate turn constraint and our takeoff constraint. While twin engine aircraft are historically sized by the second segment climb, it is not the case for us. We believe that this is due to the high L/D in second segment climb that we can achieve with our advanced technology compared to current aircraft. As a result of the constraint diagram, we chose a thrust-to-weight ratio of. and wingloading of psf. Using a database with data from Jane s All The World s Aircraft 7 and from various manufacturer s websites, we computed an industry average thrust-to-weight and wingloading which is shown on the constraint diagram. This database included aircraft such as Embraer 7-5, Canadair CRJ7 and CRJ, the Boeing 77 series as well as the Airbus A2 series. This average is shown on the constraint diagram and as a sanity check is relatively similar to our selected design point. SIZING STUDY Design Mission The design mission for YDC-NG sizing purposes consists of the following mission schedule:. Start (engine start, warm-up, taxi to active runway and departure) 2. Climb (second-segment climb to cruise altitude of FL, ft MSL). Cruise (85 NM cruise at FL) 4. Descent (descent from cruise and approach to final) 5. Full Stop (full-stop with taxi to terminal and shutdown) This mission has the advantages of being generally representative of one use case scenario as well as minimizing the complexity of sizing. Sizing Method Especially considering some of the unusual concepts generated such as solar-powered aircraft, the team needed a combination of sizing methods to evaluate different criteria during the concept selection phase. For this reason, the team generated Matlab scripts to understand the characteristics of each design. Some points to note include that in some designs, fuel was not the retarding factor. For instance, the solar-powered and electric concepts were limited by the total vehicle electrical capacitance as well as discharge and recharge rates. After narrowing down the concepts, the team elected to use the FLight OPtimization System (FLOPS) to complete the sizing analysis of the YDC-NG. Factors that led us to this sizing method included using standard and existing jet fuels. Research into Green Freedom technology allayed the team s reservations of burning hydrocarbon-based fossil fuels, such as Jet A, Jet A- and Jet B, beyond the year 28. Additionally, FLOPS is capable of giving us great amounts of detail in our studies. We initially assumed that our aircraft would be comparable in size to existing aircraft in the same category and class. These similar aircraft include: Page 25 of 4

26 Boeing 77-5/6 * Airbus A8/A/A2 * Bombardier CRJ-7/75 * Bombardier CRJ-/ER/LR * Bombardier C/CER/CLR * Embraer 7/7AR/7LR * Embraer /AR/LR * Embraer 5/5AR/5LR * Compiling data for similar vehicles allows designers understand the size of their airplane design. Additionally, these values can serve as a starting point to move forward from. Table 5, presents this summarized data for the YDC-NG, below. Table 5: Average Data for Same Category & Class Average Values Parameter Average Value Unit We 6,65 lb W 7,288 lb Empty Weight Fraction.567 nondim Wingspan 2.5 ft Wing Area 54.8 ft 2 Engine Weight,8.2 lb Baseline Thrust 8,2 lb TSFC.748 lb/hr/lb T/W.4 nondim W/S.5 psf Aspect Ratio 8.8 nondim After verifying these numbers and selecting a concept, Optoprime used these data to begin the detailed sizing process. In the FLOPS model, the team made assumptions as well as design decisions. Design Decisions included in the sizing model include: single-class passengers, 4 in seat pitch, 2 x 2 configuration, 2 NM Design Range, Extensive composites use and Cruising altitude at FL (, ft MSL) Key Assumptions included: M =.78, Aspect Ratio (AR) = 6, Thrust-to-Weight (T/W) ratio =., Wingloading (W/S) =, Specific Fuel Consumption (SFC) min =., Specific Fuel Consumption (SFC) max =.5, * These aircraft models are registered trademarks of their respective copyright holders. Page 26 of 4

27 Sweep angle of the forward wing quarter chord (Λ.25c, fore ) = 5 and Jet A fuel These assumptions have been made for the execution of the FLOPS sizing routine. While these values are in proximity to their final values, they may still vary before the configuration is finalized. We expect to be able to reduce the SFCs. No assumptions are made regarding noise level. We anticipate the positive global environmental impact of truly carbon-neutral flight to vastly outweigh the negative local impact of noise. However, the team is looking into technologies to mitigate this issue, such as feathering the blades on descent and approach to final. This would be especially possible through the future adoption of ADS-B continuous descent approaches or some other similar platform. Key items from the sizing output included: Takeoff Gross Weight (TOGW) = 86,78 lb, Lift-to-Drag (L/D) ratio takeoff = 7, (L/D) top-of-climb = 6, (L/D) cruise = 5 and Mission Fuel Fraction =.24 Trade Studies The team did not apply detailed sizing methods to our alternate concepts. The team chose the selected concept through better engineering judgment. These discussions can be found in the Pugh s Method section earlier in this document. Carpet Plots To understand the region of T/W & W/S that we are operating within, our team used graphically plotted T/W & W/S and analyzed the resulting change in the Takeoff Gross Weight of the YDC-NG. This carpet plot is shown in Figure 5, below. Page 27 of 4

28 Figure 5: Carpet Plot Trade Study of T/W and W/S From this graph, it is clear that Gross Weight is minimized by maximizing wingloading and T/W. Figure 6 makes this even more evident, below. Ideally, the YDC-NG would be designed in the greenest part of the surface, which corresponds to the same high wingloadings and low T/Ws as in the above figure. These plots were constructed in Phoenix Integration s ModelCenter software, which does not allow us to impose further constraints. However Figure 7, reprinted below from the constraint analysis, demonstrates that other constraints do, indeed exist. Therefore, our design point is as low as we can get without violating constraints. Page 28 of 4

29 T/W & W/S Trade Study Gross Weight [lb] T/W W/S [psf] Figure 6: Gross Weight Surface Plot of T/W & W/S Constraint Diagram Constant Rate Turn (TOC) Constant Rate Turn (cruise) Takeoff (β=) Landing (β=.5) Second Segment Industry Avg (.5,.425) Design Point (,.) T /W Wingloading [psf] Figure 7: YDC-NG Constraint Diagram Page 2 of 4

30 SUMMARY OF CONCEPT Overall Concept The concept that we have chosen to move forward with is a joined wing, single fuselage, aft mounted engine design. The aircraft is made of almost entirely composite materials resulting in a reduced weight with increased strength; composites also increase the feasibility of a non-circular cross section. A joined wing has advantages for both weight and drag. The aft mounted engines decrease cabin noise and increase safety. These features and more are pointed out in Figure 8. Figure 8: Walk-around Chart. Joined Wing Drag Reduction Structural Weight Savings 2. Aft-Mounted Engines Rotor Path behind PAX compartment Fuselage Noise Reduction Unducted fan Wave Rotor combustion Page of 4

31 . Trailing Edges Direct Lift/ Side Force Capability 4. Canard Possibly required for stability/ control Necessity to be determined 5. Accessibility Capability for Large Cargo Door Canard mounted high for jet way access 6. Composite Structure Weight savings Corrosion Resistance Increased Fuel Capacity Non-circular fuselage cross section Requirements Compliance The YDC-NG is likely to meet almost all of the requirements as set forth in the House of Quality, which is given in the appendix, as well as the requirements generated in Pugh s Method. The current compliance matrix is given below, in Table 6. Table 6: Compliance Matrix Compliance Target Threshold Current TOGW (lb) < 7, lb < 75, lb 5, lb Number of PAX Runway Length (ft) < 25 ft < ft 45 ft Range (NM) SFC (lb/lb*hr) > 25 NM <.5 > 2 NM 2 NM lbm lbm lbm lbf lbf lbf ( hr ) <.6 ( hr ).555 ( hr ) Thrust Available (lbf) 4, lbf 25, lbf 5,52 lbf Page of 4

32 Design Requirements Passenger Climate/Comfort Short Runway Range Gate Time Easy Maintenance Low Noise Energy Efficiency Limited Terminal Service Obstacle Clearance Crew Cost The YDC-NG is the epitome of passenger climate and comfort through its generous cabin layout. Optoprime Conceptual Designs continues to work hard to ensure that the airplane will takeoff in a short distance. The airplane has drag-reducing and weightminimizing features which help maximize range. The many exits and spacious cargo hold will enable fast turn-arounds. As this aircraft is capable of flying carbon-neutral flights, the team traded energy efficiency with solar panels and synthetic fuel for noise. However, opportunities at noise reduction do remain under consideration. The aircraft will have rugged landing gear which allows for remote service. My running on jet fuel, the aircraft will not need any specialized equipment. The joined wing configuration has the effect of reducing induced drag. Therefore, climb rates will be more than adequate. Optoprime is still analyzing remote operator opportunities to reduce crew costs. Next Steps Next steps will include detailed investigations into static stability to ensure that in case of primary system failure, the YDC-NG remains a safe aircraft to fly. Additionally, we will undertake tail sizing, engine and emissions studies. Having looked at sensitivities, we will finalize and present our aircraft conceptual design, including: aerodynamic information, performance charts, structural analysis, mass and balance, stability and control, manufacturability, production theory and other pertinent information. Finally, we will present risks, including plausibility checks for this futuristic design. Page 2 of 4

33 APPENDIX A House of Quality Relative Importance Engineering Characteristics Number of Flight Crew Negative Correlation Positive Correlation Customer Attributes Comfort (reclining seats, leg room, seat width, ceiling height) Climate (temp, air quality, lighting) Loads Quickly & Easily Green Aircraft Can Reach Appropriate Airports Gate Time (fuel, PAX luggage, cabin, overhead bins) Fuel Cost Maintenance Cost Crew Cost Low Noise Low Separation Time Short Runway Limited Terminal Service (jetway, ground crew available) Low Operating Cost / PAX Flies Fast Clears Obstacles on App/Dep Pilots on board Low Pilot Work Load Easy Maintenance Easy Weight and Balance Absolute Importance Relative Importance (%) Volume/PAX Max Altitude Runway Length Endurance Max Fuselage Height Above Ground Gross Weight Wing Height SFC Number of PAX Thrust Available Number of Engines Max Speed Best Cruise Speed Range Measurement Units (English) - # ft^/pax ft ft hr ft lb ft lb/lb*h r # lbf # kts Measurement Units (SI) - # m^/pax m m hr m kg m N # lbf # kts Target Values (English) - < kts kts nm nm Target Values (SI) - Threshhold Values (English) - Threshhold Values (SI) - Page of 4

34 REFERENCES Mankins, J.C., Technolgy Readiness Levels, NASA White Paper, Apr Matin et al., A Concept for Producing Carbon-Neutral Synthetic Fuels and Chemicals, Los Alamos National Laboratory, 27. Ciszek, Technology Review, Design Methodology, State-of-the-Art Deigns and Future Outlook, University of Virginia Department of Mechanical and Aerospace Engineering, Won, H. et al., Constant Volume Combustor implementation on a 5 passenger Commercial Regional Transport Mission Simulation, AIAA, 2. 6 Roskam, Jan Airplane Design Parts I-VIII DARCorporation, Lawrence KA Jane s, All The World s Aircraft, Cole, W. et al., Breaking the Mold, Boeing Frontiers, January 25. Raymer, D, Aircraft Design: A Conceptual Approach, 4e, AIAA 26. WOLKOVITCH, J., The Joined Wing: An Overview Journal of Aircraft vol.2 no. Now That s a Reliable Engine July 7, Boeing Current Market Outlook 27 The Airplane that Never Sleeps July 5, DC- Commercial Transport 5 AviationWeek & Space Technology, Aerospace Sourcebook, Jan 27 6 AviationWeek & Space Technology, Aerospace Sourcebook, Jan 28 7 Bureau of Transportation Statistics, 8 Bureau of Labor Statistics, Onishi, R., Flying Ocean Giant: A Multi-Fuselage Concept for Ultra-Large Flying Boat AIAA, AIAA , 24. Page 4 of 4