CONCPET DESIGN REQUIREMENTS REVIEW



CONCPET DESIGN REQUIREMENTS REVIEWChad Carmack Aaron Martin Ryan Mayer Jake Schaefer Abhi Murty Shane Mooney Ben Goldman Russell Hammer Donnie Goepper Phil Mazurek John Tegah Chris Simpson Spring 2010Table of contents TOC \o "1-3" \h \z \u EXECUTIVE SUMMARY PAGEREF _Toc260938804 \h 1MISSION STATEMENT PAGEREF _Toc260938805 \h 2AIRCRAFT DESIGN MISSION PAGEREF _Toc260938806 \h 2SELECTED BEST AIRCRAFT CONCEPT PAGEREF _Toc260938807 \h 3ADVANCED TECHNOLOGIES PAGEREF _Toc260938808 \h 7RESULTS OF AIRCRAFT SIZING AND CARPET PLOTS PAGEREF _Toc260938809 \h 9MAJOR DESIGN TRADEOFFS PAGEREF _Toc260938810 \h 15AIRCRAFT DESCRIPTION PAGEREF _Toc260938811 \h 16AERODYNAMIC DESIGN ANALYSIS PAGEREF _Toc260938812 \h 25PERFORMANCE PAGEREF _Toc260938813 \h 32STRUCTURES PAGEREF _Toc260938814 \h 35WEIGHTS AND BALANCE PAGEREF _Toc260938815 \h 45STABILITY AND WEIGHTS PAGEREF _Toc260938816 \h 46NOISE PAGEREF _Toc260938817 \h 51SUMMARY PAGEREF _Toc260938818 \h 63REFERENCES PAGEREF _Toc260938819 \h 66EXECUTIVE SUMMARYWhile the technology employed in business aviation is ever advancing, the ultimate goal of comfortably and quickly transporting corporate customers to their required locations remains steadfast. World-conscious airframe designers must not only be aware of their customer’s needs, but their environmental impact as well. Team two has designed a business jet aircraft capable of meeting our target customer’s needs while maintaining environmental responsibility. The main requirements of a long range, high cruise speed, low environmental noise and low emission aircraft were established in the System Requirements Review document. A number of concepts were generated to meet the requirements, and two designs were chosen in the Systems Definition Review document. Finally, one design was selected which best met the requirements, which was optimized for its design mission during this document, the Conceptual Design Review.In order to meet the customer needs and NASA’s Environmentally Responsible Aviation Project N+2 goals, it was determined that drastic changes would need to be made to modern business jet aircraft. In an effort to meet those goals, advanced technologies were analyzed on a lifting canard aircraft with a swept back wing incorporating two geared turbofan engines. A sizing code was developed, and historical and advanced concept data was researched. Aircraft weights and balance, aerodynamics, engine selection, and noise were analyzed in the design process. Major structural component locations were investigated, the noise produced by the aircraft was predicted, and a cost analysis has been conducted for our aircraft. The aircraft’s performance was also determined, and consideration was given to the major design tradeoffs used throughout the design process.Our aircraft is capable of transporting 16 passengers in a luxury cabin with a maximum still air range of 7,100 nm. Our aircraft will weigh 71,300 lbs, and requires 3,900 feet of ground roll. All of the customer needs were met or exceeded, but the NASA N+2 goals proved too difficult to meet with our design. Though not obtaining the NASA N+2 goals, our aircraft provided a 20db reduction in environmental noise, a 50% emissions reduction, a 25% fuel burn reduction, and a 33% takeoff field length reduction over modern aircraft of the same class. While unsuccessful in reaching the NASA N+2 goals, the customer needs were met; significant environmental improvements were achieved, and a great deal was learned about the aircraft design process.MISSION STATEMENTThis project’s main goal is the design of a cost effective, high speed, luxury aircraft capable of transporting customers to any destination as quickly as possible. The project’s secondary goal is to meet NASA’s N+2 criteria, reducing the environmental impact of the aircraft. Our proposed aircraft will be able to compete with other aircraft of the ultra long range category. AIRCRAFT DESIGN MISSION The design mission was developed and optimized with the city pair of Los Angeles and Hong Kong in mind, accounting for a 60 knot head wind. Eight mission legs between nine points constitute the developed design mission as illustrated in Figure 1 below.Figure SEQ Figure \* ARABIC 1: Design Mission Flight Plan.The first leg of the mission, from points 0 to 1, is taxi and takeoff to an altitude of 50 feet. Points 1 to 2 represent the climb portion of the mission where the aircraft climbs at best rate of climb to an altitude of 41,000 feet. From there the aircraft enters the cruise leg of the mission, between points 2 and 3, and begins cruising at a Mach number of 0.85 for 6350 nautical miles. Cruise is then followed directly by a no-range credit descent to land where the aircraft will attempt a landing from points 4 to 5. In the absence of a landing, the aircraft will then climb to an altitude of 5,000 feet at best rate climb, depicted between points 5 and 6, and cruise at an altitude of 5,000 feet to an alternate airport 200 nautical miles away. Upon arrival at the alternate airport, the aircraft will enter a holding pattern for 45 minutes, from points 7 to 8, and then begin a no-range credit descent to land. Finally, the aircraft lands at the alternate airport and completes the last mission leg at point 9.While the design mission is the intended use of this aircraft, several other operating missions can be made by this aircraft as well. One additional operating mission example would be flying from New York to Van Nuys. The distance between the two cities, which is 2,146 nautical miles, falls well within the design mission range of 6,350 miles. To compensate for the largely unused range, the aircraft can then be flown at its maximum Mach number of 0.9 with a maximum capacity of 16 passengers. This range tradeoff allows for tremendous flexibility in speed and capacity for shorter ranged flights.SELECTED BEST AIRCRAFT CONCEPTMotivated by our mission statement, the system requirements design stage produced a set of criteria by which individual concepts could be evaluated based upon the customer needs. Following the initial system requirements design stage, eight concepts were generated for further consideration. Pugh’s method was implemented as an objective procedure for selecting the best aircraft to narrow the eight concepts down to two. Early conceptual sketches of the aircraft are shown in Figure 2, concept one on the left, and concept two on the right. Some of the design criteria used in Pugh’s Method were the NASA N+2 goals, which are listed below in REF _Ref260078420 \h Table 1.Figure SEQ Figure \* ARABIC 2 Aircraft Concepts Selected by Pugh’s MethodTable SEQ Table \* ARABIC 1 NASA N+2 Goals Technology Benefits Relative To a Large Twin Aisle Reference ConfigurationCorners ofthe Trade SpaceN+2 GoalsNoise(cum. below Stage 4)-42 dBLTO NOx Emissions (below CAEP 6)-75%Performance:Aircraft Fuel Burn-50%Performance:Takeoff Field Length-50%Concept two possessed a potential reduction in induced drag at cruise due to the incorporation of a lifting canard, thereby addressing a reduction in aircraft fuel burn. Additionally, concept two addressed noise reduction by offering a main wing and twin vertical tails positioned to shield engine noise from the aircraft’s surroundings. NOx emission reduction was tied directly to aircraft fuel burn, thereby offering a potential edge to concept two due to its lower cruise drag. Neither of the two designs displayed a clear advantage in performance field length.After evaluating each of the two concepts through Pugh’s method, the potential to satisfy the NASA N+2 goals, and sizing studies, the canard configuration from Figure 2 was selected for the remainder of our design.Aircraft Walk-aroundSeveral important features provided the canard design with advantages over the other aircraft configurations evaluated in the concept selection process. Particular aspects from all or many of the concepts were also shared with the canard configuration’s final design in order to meet the specified criteria. A number of the unique and shared external characteristics present on the canard configuration aircraft are depicted in Figure 3 below. Lifting CanardsFuselage – aft Mounted EnginesNoise ShieldingVertical StabilizersNoise ShieldingLow WingCircular FuselageSpiroid Wing-TipsFigure SEQ Figure \* ARABIC 3 Canard Concept Walk-Around ChartShown in Figure 3, the canard design features lifting canards in an effort to obtain a reduction in induced drag during the cruise portion of the design mission. Also aimed at reducing induced drag at cruise, our design incorporates spiroid wing-tip devices as shown. The previous features were included to fulfill the aircraft fuel burn reduction metric of the NASA N+2 goals listed in Table 1. These features also act to reduce NOx emissions since a lower fuel burn will produce fewer emissions. Also present on the canard configuration are several noise reducing design features, which are shown in Figure 3. Engine noise was determined to be the primary contributor to environmental noise, and the canard configuration made several efforts to shield the environment from noise generated by the engine. The wing was placed low and aft on the aircraft, beneath both of the engines to shield the engine noise from the ground during the takeoff, climb, descent, and landing mission segments. As lateral noise propagation is also a significant concern for airports in populated areas, twin vertical stabilizers were incorporated on top of the main wings outboard of the engines, and in line with the engine exhaust.From the walk-around chart depicted in Figure 3, it is clear that several of the external design features directly impact the internal cabin of the aircraft. Both engines are mounted aft of the main cabin and on pylons connected to the fuselage. The aft mounting of the engines affords a quieter cabin interior, while the fuselage mounted pylons facilitate a reduction in cabin vibration and provides a minimum structural weight for engine mounting. The design of a circular fuselage cross section provides a relatively easy method of cabin pressurization, and allows a relatively large internal cabin volume for passenger comfort. A circular fuselage can be constructed from current industry practices, thereby lowering manufacturing developmental costs.The low, aft mounted main wing also allows for minimal main spar and cabin interaction. The lack of structural requirements inside of the main cabin helps maximize internal cabin volume, yielding additional passenger comfort. Furthermore, the aft mounted main wing facilitates a clear view of the ground from nearly every seat in the cabin, which provides each passenger with a visually pleasing exterior view.The benefits of the aircraft’s design features from Figure 3 provide insight to the reasons for the choice of concept. A more thorough investigation of the important design features and advanced technologies incorporated on the design follows in the next section of the report.ADVANCED TECHNOLOGIESOf all the advanced technologies that could be included, engines are perhaps the most important since the engines will have a significant impact on all four of the NASA N+2 goals. Our group considered two engines, an unducted turbofan and a geared turbofan. For reasons discussed later in the report, a geared turbofan was chosen over the unducted fan. A geared turbofan engine utilizes a gearbox between the fan and main power shaft to decouple the fan from the engine’s compressor. The technology allows for a very large fan to rotate at slower speeds while moving large amounts of air with little noise. The gearbox also allows for the compressor and turbine to spin at faster speeds to further increase efficiency. Pratt and Whitney claims that the PurePower 1000 series offers proven efficiency with no life limited parts, a 20 dB reduction in noise over current engines, a reduction in NOx emissions by 50% over the CAEP/6 margins, and a 15% reduction in fuel burn over current engines.3 The PurePower 1000 series offer a thrust range between 13,000 lbf and 24,000 lbf depending on engine selection. This thrust range is ideal for our aircraft and allows the team to continue using the rubber engine assumption in sizing.Another advanced technology included was the use of composite materials. Aircraft such as the Boeing 787 have achieved a significant empty weight reduction by utilizing composite materials in the majority of the airframe. A reduction in aircraft weight will allow for reduced fuel consumption and a reduction in take off length. Up to 50% of the structure of some aircraft has been constructed out of composite materials. There has been a lot of research on the use of composite materials within the past few decades. The technology has been proven to be reliable and beneficial in the aerospace industry and is currently being fielded in many production designs. This technology is likely to keep improving and delivering even more advantages in aircraft design by the 2020 production date. Research and past applications suggest a reduction of approximately 20% in structural weight. This value is used directly as a technology factor in the sizing process and is applied to the structural weight components of the fuselage, wings, horizontal and vertical tails, canards, pylons, and nacelles.Spiroid wingtips were the last advanced technology incorporated into the design. These wingtips have been shown to significantly reduce fuel burn in flight testing. When the wingtips first flew on a Gulfstream II, they yielded a 10% improvement in fuel burn. Figure 4 shows the general geometry of these devices, which estimates state can provide a 6-10 % reduction in cruise drag. This technology has been incorporated on more than 3,000 aircraft, including several business jets, as well as the Boeing 737 and 757 airliners. Spiroid wingtips also aid the FAA in increasing airspace capacity near airports. Spiroid wingtips have the potential to decrease wake intensity, which could substantially alter the requirements for separation distances between aircraft in airport traffic patterns. Figure SEQ Figure \* ARABIC 4: Spiroid Wingtip DevicesValues of Major Design ParametersWith the basic aircraft configuration, important design features, and advanced technologies chosen, the design was iterated through the developed sizing code until the aircraft’s major design parameters were optimized for the design mission. A number of the optimized parameters including thrust to weight ratio, aspect ratio, wing loading, and main wing and canard areas were selected and are shown in Table 2. While each of these parameters can be further broken down into dimensions more commonly provided for an aircraft, these dimensions are more easily shown in an aircraft three-view drawing provided in the Aircraft Description section of this report. Table SEQ Table \* ARABIC 2: Major Aircraft Design ParametersParameterValueThrust / Weight Ratio0.34Aspect Ratio12Wing Loading87 (lb/ft2)Wing Area796.4(ft2)Canard Area147.4 (ft2)RESULTS OF AIRCRAFT SIZING AND CARPET PLOTSIn order to numerically evaluate and size our designs, a sizing code was developed in MATLAB. The code follows the general procedures outlined in the flow chart shown in REF _Ref260670726 \h Figure 5. Figure SEQ Figure \* ARABIC 5: Sizing Code Flow ChartThe empty weight calculation is based on a component build-up method using equations from the Transport Aircraft category of Daniel Raymer’s textbook.8 Once the code was established; it was calibrated against the Beechcraft Starship for our canard design and against the Gulfstream G550 for our conventional design. The results of the calibration process yielded the calibration factors listed in REF _Ref260670752 \h Table 3. Table SEQ Table \* ARABIC 3: Calibration FactorsWeightConventionalCanardFuel Weight0.890.89Empty Weight1.160.96Gross Weight1.030.98 It is important to note that the fuel weight calibration factor is the same for both designs, which was intentional. In calibrating the canard model to the Beechcraft Starship, the fuel weight calibration factor was extremely small, as low as 0.50. This large discrepancy was attributed to the turboprop engine used on the Starship. Because of the fundamental differences between a turboprop and a geared turbofan or unducted fan, the fuel weight calibration factor found for the conventional design was used to replace that of the canard design. Finally, with the code completed and calibrated, the technology factors were then applied and each model was analyzed. A list of the technology factors used in the sizing code is provided in REF _Ref260670823 \h Table 4. The value of each factor comes from the applied advanced technologies discussed previously in this report.Table SEQ Table \* ARABIC 4: Technology FactorsApplicationTech ValueWstructure0.80Di (canard only)0.93SFC0.75The conventional and canard aircrafts were sized and optimized through the use of carpet plots by running the sizing code through many iterations by varying thrust to weight, wing loading, and aspect ratio. These three variables were the main driver for sizing the aircrafts. The most optimized case was a result of whichever combination of these variables gave the lowest gross weight. The goal of this optimization was to produce the lightest plane that could carry out the design mission. Several constraints limited the weight of the aircraft and the three design variables. The constraints for the aircraft were top of climb (>100 ft/m climb rate at cruise altitude), 2-g maneuver, takeoff ground roll (<4000 ft), landing ground roll (<3000 ft), and climb gradient for second segment climb (2.5%). Thrust to weight and wing loading were varied in order to find the minimum gross weight at a given aspect ratio. Five different aspect ratios were tested for each model, and a trend was then established between minimum gross weight and aspect ratio. Below in REF _Ref260239016 \h Figure 6 is the trend that was seen for the conventional aircraft. From this plot it can be seen that the minimum gross weight for the conventional design is 76,000 lbs at an aspect ratio of 10.Figure SEQ Figure \* ARABIC 6: Gross Weight vs. Aspect Ratio of Conventional AircraftFurther analysis on this design point can be seen by examining the carpet plot for an aspect ratio of 10, shown in REF _Ref260239074 \h Figure 7 below. If a point lies within an area shaded in grey, it means that it cannot meet one or more of the constraints. Thus the minimum gross weight is at a wing loading of 85 and a thrust to weight of .32 which yields a gross weight of 76,000 lbs. This point is limited by both top of climb and takeoff ground roll. Landing distance, second segment climb, and subsonic 2-g turn are not limiting constraints.Figure SEQ Figure \* ARABIC 7: Carpet plot for conventional aircraftThe same analysis was done for the canard aircraft. A similar trend was seen for the relationship between gross weight and aspect ratio. A plot of this trend is shown in REF _Ref260239123 \h Figure 8 below. The major difference from the conventional design is that the optimal aspect ratio is now slightly higher at 12 compared to 10. At an aspect ratio of 12, the aircraft weights only 71,000 lbs. This is nearly 5,000 lbs lighter than the conventional aircraft and is one of the reasons for choosing the canard configuration over the conventional.Figure SEQ Figure \* ARABIC 8: Conventional a/c sizing resultsThe carpet plot for the canard aircraft corresponding to the minimum gross weight is shown below in REF _Ref260239194 \h Figure 9. This shows that the minimum gross weight occurs at a wing loading of approximately 87 lb/ft2 and thrust to weight slightly below 0.34. This optimal point is constrained by top of climb and takeoff ground roll. The other three constraints were not a factor, which was the same result as the conventional aircraft. Figure SEQ Figure \* ARABIC 9 Carpet plot for conventional aircraftFrom this sizing optimization, the results show the canard aircraft is approximately 5,000 lbs lighter than the conventional design. The canard aircraft weights 71,300 lbs, of which 38,000 lbs is empty weight and 31,500 lbs is fuel. The takeoff ground roll is just below 4,000 ft and the landing ground roll is 2,200 ft.MAJOR DESIGN TRADEOFFSOne of the concepts selected proposed the use of a lifting canard as opposed to a conventional tail configuration. The lifting canard can theoretically provide the design with better cruise performance by reducing induced drag at cruise conditions. In a conventional tail configuration the horizontal stabilizer provides a down force in flight for aircraft stability, therefore the wings have to create additional lift to counter act the down force from the tail which in turn means the wings create additional induced drag. The lifting canard balances the pitch stability of the aircraft along with providing lift, therefore the main wings do not have to produce as much lift as a conventional design would require. The cons for a canard can outweigh the benefits. Specifically, the downwash from the canards could actually produce enough interference on the main wings to actually create more drag. In fact, the interference between the canard and main wing occurs in almost all flight conditions except for the specific design point at cruise where the canard configuration will benefit induced drag. Additionally, the FAA mandates that the canard must stall before the main wing to prevent loss of aircraft stability. Should the main wings stall before the canards, the lifting force provided by the canards will cause a violent pitch up causing the main wings to go into a deeper stall and possibly prevent the aircraft from recovering. The pros of having the canards stall before the main wing actually prevents the main wings from stalling unintentionally by preventing the aircraft from flying at high angles of attack. The cons of having the canard stall first means that main wings cannot ever reach maximum lift which requires the addition of high lift devices for slow flight and landing. The position of the main wings at the very rear of the aircraft actually provides a difficult design problem for installing high lift devices due to the large pitching moment created by high lift devices so far aft of the center of gravity and careful attention must be paid when designing the aircraft for landing configurations.The canard design was calculated to have a smaller empty weight than the conventional design, but also had a slightly larger fuel burn during the design mission. Therefore the canard design actually had a worse drag profile than the conventional aircraft which would suggest that the canard did not perform well in reducing the aircraft’s drag during cruise. The team decided to use the canard configuration despite this fact to try and meet the N+2 noise reduction goals. The canard position allows the engines to be placed in the rear of the aircraft above the main wings which provide a significant amount of noise shielding to the ground. Two vertical stabilizers are also mounted on the main wings of the aircraft to reduce engine noise. The two engines are essentially blocked by the main wings and vertical stabilizers which significantly reduced the noise signature of the aircraft.Additional tradeoffs were considered pertaining to the aircraft’s cabin. Space versus performance considerations for the cabin involved installing two lavatories, a galley, and larger sleeping quarters for a reserve pilot. The addition of such amenities forced the cabin size to be rather large, but a large selling point in business aircraft is comfort. The two lavatories were necessary for a full passenger load, especially for long distance flights. The galley was added for long distance flights including trans-pacific flights where meals would have to be served due to the long duration. The pilot resting quarters were added for a reserve pilot, which in the case of the trans-pacific flights a reserve pilot is mandated by FAA law due to limits on pilot-in-command time. The cabin also was designed to have a stand up center aisle to push the design into the “plush” category of aircraft comfort. AIRCRAFT DESCRIPTIONOnce the external design features of the aircraft were determined, the aircraft configuration was iterated through the sizing code, the major design parameters were optimized, and the aircraft’s outer mold line was set and modeled in CATIA V5, R19. The model incorporated the design features as previously mentioned, a low rear wing design with a lifting canard, circular fuselage, twin vertical outboard mounted tails, and aft-fuselage mounted engines on pylons. The fuselage seats sixteen passengers, two crew, two pilots, and an additional pilot in the pilot rest area for trans-oceanic missions at maximum capacity. The following three subsections offer a more detailed investigation of the aircraft’s outer mold line, the aircraft’s interior layout, and the included cabin amenities and features respectively.Dimensioned Three – View To ScaleFrom the aircraft sizing code and the major design parameters listed in Table 2, a thorough set of dimensions were calculated for the aircraft concept, which are shown in the scale three-view drawing of the aircraft presented in Figure 10, generated from the CATIA model. The three-view drawing was constructed at a scale of 1:300 with reference to the actual aircraft’s completed size when plotted on standard 8-1/2” x 11” paper. Figure SEQ Figure \* ARABIC 10: Dimensioned Scale Three View of Team 2 Final Design, Scale 1:300To provide additional ease of reference, a table of the major exterior aircraft dimensions was compiled and is presented below in Table 5. The values previously listed in the major design parameters compilation presented in Table 2 are omitted here for brevity.Table SEQ Table \* ARABIC 5: External Aircraft DimensionsDimensionValueMain WingsWingspan97’-9”Wing Sweep30?Root Chord13’-9”Tip Chord2’-8”Strake Sweep45?Strake Root Chord24’CanardsWingspan36’-5”Canard Sweep35?Mean Chord3’Vertical TailsSpan Each13’-5”Tail Sweep30?Mean Chord10’-1”Width Between30’-6”Fuselage Overall Length88’Main Cabin Length50’Main Cabin Outer Diameter8’-10”Sill Height6’-3”Internal Layout / Arrangement To ScaleWith the dimensions of the outer mold line and arrangement in place, the interior arrangement of the aircraft could be finalized. The crew area of the aircraft including the flight attendants reclining seats, as well as the pilot berthing area were located at the foremost part of the main cabin, forward of the main entrance door. Seating for sixteen passengers at maximum capacity was distributed about the main cabin, with a conference area, lavatory, and small cocktail galley present in the main cabin area. An emergency exit was also placed across from the conference area and over the leading edge of the main wing on the opposite side of the aircraft from the main entrance. At the rear end of the main cabin area, a divider wall separates the rear lavatory, preparation galley, and in cabin storage from the main seating area. As with the external mold line of the aircraft, the cabin interior was completely modeled in CATIA V5 R19, and a 1:150 scale three view drawing was generated and is depicted in Figure 11.Figure SEQ Figure \* ARABIC 11: Overall Interior Cabin Three-view DrawingIn order to provide more insight to the interior cabin dimensions, a detail drawing of the conference and small galley area was generated from the above plan, and more dimensions were added, as shown in Figure 12. Additionally, a cross sectional view of the interior and exterior cabin dimensions was constructed to provide a clearer view of the relevant seating dimensions and aisle height, and is depicted in Figure 13. Figure SEQ Figure \* ARABIC 12: Detail Drawing of Conference Area and Cocktail GalleyFigure SEQ Figure \* ARABIC 13: Fuselage Cross Section and Seating DimensionsSimilar to the table constructed for the exterior aircraft dimensions, a table of interior layout and seating dimensions was constructed for the main cabin, shown in REF _Ref260937755 Table 6.Table SEQ Table \* ARABIC 6: Interior Cabin and Seating DimensionsDimensionValueMajor Cabin DimensionsOuter Cabin Diameter106”Inner Cabin Diameter100?Aisle Height6’-3”Minimum Aisle Width27”Main Cabin Length50’Additional Cabin Length10’Seating DimensionsStandard Recliner Width27”Standard Recliner Depth33”Square Table Length / Width26”Flight Attendant Recliner Width20”Flight Attendant Recliner Depth33”Sofa DimensionsSofa Length7’-6”Sofa Depth35”Conference Area DimensionsConference Seat Width4’-4”Conference Seat Length8’-4”Conference Table Width2’-6”Conference Table Length4’-4”Cabin Layout / AmenitiesThe aircraft’s cabin was designed with the intent of providing a plush experience for each and every passenger on board. A number of amenities and features were incorporated to heighten the travel experience. The main cabin features a modern technology conference area, complete with four passenger conference seating around a common table, a conference computer desk and chair, and a retractable projector screen. More typical aircraft seating can be found in the six large fully reclining chairs, two pairs of which are arranged in a face to face seating and share common fold down tables. For more casual seating, two sofas are included in the main cabin area, capable of seating three passengers each with a maximum capacity, or one passenger each in complete luxury. An emergency exit is provided on the opposite side of the main cabin entrance at the rear of the cabin and over the leading edge of the main wings, near the conference area as shown in Figure 14.Figure SEQ Figure \* ARABIC 14: Emergency Exit Located at Rear of Main CabinAs the business traveler requires sustenance to maintain a high level of productivity, two galleys were included in the aircraft cabin layout. A large fully equipped galley is located at the rear of the aircraft designed for both flight attendant and passenger use; while a smaller more easily accessible cocktail and serving galley is located near the center of the aircraft, and near the conference area, shown in Figure 15.Figure SEQ Figure \* ARABIC 15: Cocktail and Serving Galley, as Viewed from Conference Area Based upon the maximum cabin capacity, it was ascertained that two full lavatories would be required to provide the comfort level desired. One lavatory was placed forward in the fuselage near the crew seating and main entrance door, and one placed after the rear main cabin divider easily accessible from the conference area.Cabin design and comfort was not only focused around the passenger, a comfortable crew rest area was incorporated in the main cabin, isolated from the passenger area. A curtained pilot berthing area was incorporated in the foreword most section of the main cabin, capable of providing a comfortable rest area for a third pilot in rotation on extended flights as shown in Figure 16. Figure SEQ Figure \* ARABIC 16: Pilot Berthing Area and Crew Rest SeatingAdditionally, two fully reclining flight attendant seats are located forward in the main cabin, also capable of being separated from the main cabin seating area by a dividing curtain. Storage for the cabin materials and crew baggage was provided by the reserved areas aft of the rear cabin divider adjacent to the main preparation galley. A rendered overhead view of the internal cabin is provided below in Figure 17.Figure SEQ Figure \* ARABIC 17: Overhead View of Internal Cabin LayoutAt full capacity, the aircraft cabin holds a maximum of sixteen passengers, two crew, and three pilots. Also at maximum capacity, each passenger is provided with a volume per passenger of 150 ft3. The volume per passenger metric equals or exceeds other plush business aircraft of the same class, but is typically specified in terms of volume per passenger per hour, from which varying comfort levels are classified. These comfort levels are defined in Figure 18.3 Extrapolating the data to longer durations and larger cabin volumes, it can be easily shown that the 150 cubic feet offered by the cabin design maintains the plush level of comfort for any trip at maximum capacity, and improves even further for shorter missions.Figure SEQ Figure \* ARABIC 18: Comfort Level of Cabin Volume / Passenger / Trip Duration3Given the large internal cabin volume, the plush comfort level for any duration and capacity flight, and ample amenities, it becomes fundamentally clear that our interior cabin was designed with every level of passenger and crew comfort and productivity in mind.AERODYNAMIC DESIGN ANALYSISAirfoil SelectionTo meet the desired lift on the aircraft, the airfoils for the main wing, canard, and the vertical tail were selected based on the needed coefficient of lift, or CL, for three different phases of flight. These phases were the takeoff CL, cruise CL, and landing CL. These three values are listed in REF _Ref260241531 \h Table 7.Table SEQ Table \* ARABIC 7: Desired Coefficient of liftCLTakeoff1.2Cruise0.46Landing2From the sizing code, the airfoil was set to have a 0.1 thickness to cord. This limited the airfoil selection some. The supercritical wing is advantageous for this aircraft as it provides lift and stall characteristics comparable to non-supercritical wings while minimizing wave drag as is inherent in the design of a supercritical wing. It was also ideal to pick a supercritical airfoil because it would lower the MDD for the drag prediction. To pick a supercritical airfoil, the RAE 2822 was used as a baseline model. This was done because the RAE airfoil had much more information available for it. Airfoils with the proper thickness to cord were compared using CL vs. angle of attack plots. The results of this comparison can be seen in REF _Ref260241595 \h Figure 19. This comparison was done using a relatively low Reynolds number of 100,000. It is important to note that the hump that each of the airfoils has is due to the fact that the information is being gathered at a relatively low Reynolds number. From this table, the NASA SC(2)-0610 airfoil was selected. This particular airfoil was selected because it can meet the desired cruise CL of 0.46 with having minimal twist in the wings. This particular airfoil has a significant pitching moment, but it was deemed to be manageable and could be worked around if the lifting canard could balance this out.Figure SEQ Figure \* ARABIC 19: Supercritical Airfoil comparison13With the airfoil selected, it can be analyzed at different Reynolds numbers. The geometry information for the NASA SC(2)-0610 is given in REF _Ref260241644 Figure 20. A visual of the airfoil can also be seen in REF _Ref260241644 \h Figure 20, and more information about the geometry can be seen in REF _Ref260939175 Table 8. Using this geometry, the airfoil had been further evaluated at different low Reynolds numbers. This comparison can be graphically seen in REF _Ref260241661 \h Figure 21. It is important to note that as the Reynolds number increases, the stall angle transitions higher along the curve. The drag polar for this airfoil can also be seen in Figure 22, where the trend is that as the Reynolds number increases, the coefficient of drag decreases. The canard used a slightly thinner supercritical airfoil, with a thickness to cord of 0.08. This thinner airfoil, NASA SC(2)-0608, does not provide as much lift or as great of a moment, but is more than sufficient to meet the requirements of the canard. This airfoil will also stall earlier depending on the angle that it is set at, which will help to keep the aircraft stable. Table SEQ Table \* ARABIC 8: NASA SC(2)-0610 geometryGeometryThickness:10.00%Camber:1.00%Trailing edge angle:2.9 degLower Flatness:12.50%Leading edge radius:1.70%Figure SEQ Figure \* ARABIC 20: NASA SC(2)-0610 airfoil.Figure SEQ Figure \* ARABIC 21: CL and Cm vs. Angle of Attack13Figure SEQ Figure \* ARABIC 22: Drag Polar13The selected airfoil meets the cruise CL, but does not meet the climb or the landing CL of 1.2 and 2.0. To obtain the required CL for these two phases of flight, a high lift device needs to be included on the aircraft. The landing CL was used for choosing the appropriate high lift device because it is the higher of the two coefficients of lift, requiring a CL of 2.0. To find the effects of adding a regular or slotted flap, historical data from Kenneth W. Goodson’s report was used to see how greatly the two types of flaps would affect a supercritical airfoil.14 The airfoil used for this data is different than the airfoil selected for our aircraft, but it provides a good baseline for comparison of the effects the flaps will have on the coefficient of lift. The results from this data are shown in REF _Ref260241872 \h Figure 23. From this plot it can be seen that the regular flap will not produce a high enough CL to reach 2.0. The single slotted flap, on the other hand, can reach the desired amount of lift with only a minimal angle of attack, assuming the flap is deflected at a sufficient angle.Figure SEQ Figure \* ARABIC 23: Effects of flaps on a supercritical airfoil14The next airfoil that needed to be selected was that of the vertical tail. For this, a laminar flow symmetric airfoil was ideal. This was picked because the vertical tail has a relatively small operating range, and does not need to produce lift at a zero-degree angle of attack. Laminar flow airfoils extend the transition point from laminar to turbulent flow on the airfoil. This can help to reduce drag. The desired thickness to chord was obtained from the sizing code and found to be 0.08. After comparing different laminar flow airfoils, the NACA 64(1)-008 was selected. This airfoil can be seen in REF _Ref260241946 \h Figure 24. The resulting Cm vs. angle of attack for this airfoil can be seen in REF _Ref260241970 \h Figure 25. This data is again taken at a very low Reynolds number. When a rudder is attached to the airfoil, it will generate the desired Cm needed for maneuverability.Figure SEQ Figure \* ARABIC 24: NACA 64(1)-008 Laminar flow airfoil.Figure SEQ Figure \* ARABIC 25: Cl and Cm vs. Angle of Attack for Vertical Tails13Drag PredictionThroughout our design we considered drag to be built up of four components, namely wave, induced, pressure and miscellaneous drag. We also considered the miscellaneous drag to be a steady five percent of the pressure drag. This meant that we immediately reduced the number of drag variables from four to three. Our earlier reports simply assumed the wave drag coefficient to be 20 “counts,” or CD,wave = 0.002. Our pressure drag approximations were much more in depth and took into consideration all of the major components of the aircraft including the fuselage, engines, nacelles, pylons, wings etc. The remaining, or miscellaneous, components were appropriately bundled into the miscellaneous drag coefficient approximation. With the aircraft’s wing airfoil chosen, the calculation of induced drag was possible. However, a number of assumptions needed to be made. It was assumed that at cruise, the combination of the air’s viscosity and the local length scales were not enough to damp out the downwash produced by the canard prior to its interaction with the main wing. With this assumption made, induced drag of the canard was calculated with an induced drag prediction code developed in AAE 334, while the main wing’s induced drag was found assuming the linear summation of the downwash angle of attack produced by both the canard and main wing. The induced drag prediction found an overall induced drag coefficient of approximately 0.0175. A technology factor was then applied of 0.93, which is a result of the advanced technologies included in the design, to yield a final induced drag coefficient of 0.01002. The last part of the drag component buildup to be completed was a better calculation of wave drag. The wave drag coefficient is approximated according to the following equation. CD,wave = 20*(M – Mcr)4Here, Mcr is the critical Mach number, the Mach number at which wave drag first occurs. Mcr is usually considered to occur at a Mach number 0.07 less than the drag divergence Mach number, MDD. However, the use of a supercritical airfoil pushes MDD forward such that Mcr occurs at a Mach number only 0.01 less than the MDD calculated for a non-supercritical wing. MDD was found to be approximately 0.83, yielding a Mcr of 0.82. All of the information needed to calculate the wave drag coefficient at cruise is now known. Overall, the aircraft was found to generate a drag coefficient of 0.02665 at cruise. This drag coefficient is less than that created by the smaller Cessna 172 Skyhawk, but slightly higher than Boeing’s newest premier airliner, the 787.PERFORMANCEV-n DiagramIn the V-n diagram for the aircraft, several factors were considered. First as seen in REF _Ref260242459 \h Figure 26, the aircraft was limited by the line outlined in red. On the positive load values, at low velocity (up to 790 ft/s) the aircraft is limited by Clmax. The equation used for this calculation is listed below. After this velocity, the aircraft is limited by the gust curve up to the cruise velocity and the first Δn value. This value and the other Δn values were found by taking the gust speeds from the FAA at their altitude and interpolating for 41,000 ft at each velocity. Finally the curve is limited by the dive velocity at the far end. This was chosen to be Mach 0.87 for this aircraft. This same process was done for the negative load factors.n= Clmax*12**V2W0/S Figure SEQ Figure \* ARABIC 26 Aircraft V-n DiagramAdditionally, a payload range diagram was generated for our final design. This simple plot shows how far the aircraft could fly in any loading scenario. The payload range diagram is shown in REF _Ref260674227 \h Figure 27, and is applicable to a cruise Mach of 0.85 at an altitude of 41,000 feet. The ranges listed are still air ranges.Figure SEQ Figure \* ARABIC 27: Payload Range DiagramPROPULSIONThe engine was modeled using a rubber engine approach. The MATLAB code that models the engine uses the tabular engine data for the CF-34 that was provided in class. The engine code scales this data based on the sea-level static thrust of the engine deck and that which is required by the aircraft. The SFC data is also scaled using a technology factor to account for new technologies such as the use of a geared turbofan. This technology factor also includes a small correction factor for installation losses. The sizing code uses the developed engine model to calculate the weight of the fuel required in each segment of the design mission.Once the aircraft is sized, the engine and drag models were used to generate thrust required and thrust available curves for a range of speeds at sea-level and cruise altitude. The thrust available curves are generated by simulating the engine model at full throttle at a given altitude and a range of velocities and multiplying the resulting thrust by the number of engines, REF _Ref260941456 \h Figure 28. The thrust required is determined by the drag on the aircraft. Figure 29 shows that there is a maximum operating airspeed, which corresponds to a maximum operating Mach number of about 0.89.Figure SEQ Figure \* ARABIC 28: Required and available thrust at sea-level for various airspeeds.Figure SEQ Figure \* ARABIC 29: Required and available thrust at cruise altitude for various airspeeds.STRUCTURESThe placement of major structural components during the preliminary design process was investigated in order to provide a starting point for future detailed design. Depicted in Figure 30 is a diagram of the major structural components that compose the outer mold line of the aircraft. Examining historical cutaways from similarly sized and loaded aircraft, two spars were deemed necessary for the preliminary design process. The upper and lower spars are connected by shear webbing as indicated on the diagram in green, and the wing and fuselage intersection is filleted to provide lower aerodynamic interference as shown in blue. Ribs are modeled with equal spacing to maintain the airfoil’s shape and transfer the distributed skin loads from to the spars. The vertical tails and canards share a similar construction to the main wings, though with one main spar in each to account for the lower force distribution and span on each component.Also shown in Figure 30 is a rough sketch of the major components of the fuselage structure. The fuselage is constructed from structural frames to maintain its cross-sectional shape, connected with longerons to resist bending and to support the aircraft’s skin. Reinforced bulkheads are integrated into the main frame at the fore and aft most portion of the cabin to provide a barrier for the pressurized ends of the main cabin. Sills were modeled above and below the fuselage cutouts for the doors and windows, and beams with several webbing cross members were placed on the sides of each door and window to tie the sills together and to connect them to their respective fuselage frames. The sills and beams are indicated in red and blue around each door and window in Figure 30.Figure SEQ Figure \* ARABIC 30 Overview of Major Structural MembersLoad PathsAfter modeling a sampling of the major structural components on the aircraft frame, the major aircraft loads were applied to the structural model in low fidelity to provide evidence of the structural members needed to carry the specified loads. Distributed lift and weight loads were applied to the airframe through the skin, and ribs, which was transferred from the ribs to the wing’s spars, and passed to the fuselage. The drag force on the wings and fuselage are carried by the aircrafts skin, passed to the ribs, spars, frames, and longerons, and is distributed throughout the entire aircraft framework along with the fuselage weight. Point loads from the thrust of both engines are carried through the pylon supports into the beams between the large fuselage frames surrounding the pylons, and is likewise distributed through the fuselage assembly through the frames and longerons. The major loads applied to the aircraft and a low fidelity model of the structural members that carry them are shown in Figure 31.Figure SEQ Figure \* ARABIC 31 Structural Highlights: Major Loads and Load PathsWing-Fuselage IntersectionA more detailed sketch was drawn for the wing-fuselage intersection, and is presented in Figure 32. The distributed weight, lift, and drag loads carried by the wings are transferred from the wing’s skin to the ribs, then distributed across the wings spars, and are carried from the spars to the fuselage. Surrounding the wings and the wing spars at the fuselage are reinforced frames that are connected with beams above and below the wing intersection, to form a wing box that accepts the wing cradle formed by the strake assembly. The wing box distributes the loads imparted by the wing cradle to the longerons, thereby distributing the wing forces across the entire fuselage frame assembly. Also visible in Figure 32 is the fillet surrounding the root of the wing at the fuselage intersection, which distributes both the forces carried by the wing skin and smoothes the aerodynamic transition point between the wing and fuselage. Figure SEQ Figure \* ARABIC 32 Wing-Fuselage Structural IntersectionEngine Pylons / MountsSimilar to the wing–fuselage intersection above, a more detailed sketch was also compiled for the engine pylons and pylon mounts, shown in Figure 33. Carrying the entire thrust loads of the aircraft, the pylons and pylon mounts distribute a focused large amount of force over a small cross sectional area. Additional frame reinforcement is expected in the frames immediately fore and aft the pylon areas. The main engine weight is supported by a high strength rod spanning the center of the pylon shroud, and bending and other loads are carried by the pylon shrouds constructed from ribs and a single spar in a similar manner to the canard and vertical tails. Loads transferred from the pylons and main pylon support rod encounter a large beam at the fuselage, spanning several reinforced structural fuselage frames, and the aft cabin pressure bulkhead. The loads carried by the reinforced frames and bulkhead are then distributed throughout the fuselage frame assembly by longerons and additional frames.Figure SEQ Figure \* ARABIC 33: Pylon / Engine Mounting Structural FrameworkLanding Gear IntegrationSimilarly to the engine-pylons and wing–fuselage intersection, a more detailed sketch was constructed for the landing gear placement in the main wings shown in Figure 34. Investigation of historical landing gear structure for similar aircraft indicated the landing gear assembly to be supported by beams attached to reinforced ribs outboard on the main wings. In order to provide adequate wing thickness to accommodate retractable landing gear, the landing gear must be placed near the twin vertical tails, to allow a large enough bay into which the landing gear could retract. The loads from the main gear on landing are transferred through the landing gear assembly, the reinforced ribs, and into the wing spars to be passed to the fuselage in the same manner as the lift, weight, and drag forces on the wing.Figure SEQ Figure \* ARABIC 34: Landing Gear Structural Framework DetailMaterial SelectionIn order to provide a reduction in empty weight fraction, research was conducted on the incorporation of a combination of both advanced and traditional materials on the airframe. The materials selected to be used on the aircraft were chosen based on five main factors.Strength to Weight RatioFatigue ResistanceMaintenanceCostIce-phobic CharacteristicsIn order to meet the chosen requirements, the following materials were considered.Advanced CompositesAdvanced composite materials were taken into consideration because of their high strength to weight ratio and tailoring capabilities. Use of advanced composites on commercial aircraft has been limited due to material and manufacturing costs, as well as poorly developed inspection and repair practices. With few exceptions, commercial airliners have incorporated a maximum structural weight of 15% composites until the last decade. More recent undertakings such as the Boeing 787 have achieved a composite makeup of up to 50% structural weight, and will help promote the development of more established inspection and maintenance practices, resulting in potentially easier aircraft certification. Our design will incorporate a quasi-isotropic [45°/0°/-45°/90°]s symmetric layup of AS4-3501-6 carbon-epoxy unidirectional pre-preg on the wing skin and control surface panels.15 Quasi-isotropic carbon-epoxy layups possess a specific strength of three times to five times that of the aluminum alloys typically used in wing skins panels.15 Due to the addition of cooling, ice-phobic coatings, and electrical grounding systems required in higher composite fraction aircraft based upon historical research, a more modest 50% reduction in empty weight fraction for the relevant components is expected.While there are significant potential benefits from the use of advanced composite materials, a number of difficulties have historically limited their use on aircraft. Current theoretical failure prediction incorporates a variety of inaccurate assumptions, but this adversity is becoming largely offset by the growing historical database of composite aircraft. Additionally, fatigue and fracture detection on composite materials lacks a wide array of non-destructive techniques, leading to historically difficult certification of primarily composite aircraft. Recent improvements in ultrasonic scanning, x-ray, and acoustic emission techniques are projected to increase with the advent of the Boeing 787, and it is expected that the newly developed field will become sufficient to ease the certification of primarily composite airframes by the 2020 manufacturing date.The development of new open-and-closed molding techniques, as well as injection molded resin transfer methods for cylindrical fuselage sections, have reduced manufacturing costs for composite aircraft construction. Despite the manufacturing cost reduction, composite manufacturing costs are still higher than more conventional manufactured aluminum aircraft. A reduction in the manufacturing cost of composite materials on aircraft is expected resulting from the combination of a lower parts count afforded by composite panel layup, as well as mass production advances developed through the increased use of composites on the Boeing 787.Advanced Aluminum AlloysWhile a great deal of research and discussion has been conducted to the use of composites in the aerospace industry, aluminum alloys have continued to advance as well. Recent developments in Aluminum-Lithium alloys have been targeted directly at the aerospace industry due to the high strength to weight ratio, exceptional fatigue performance, and cryogenic toughness properties. Aluminum-lithium alloys have existed since the 1950’s, and were developed by adding lithium to aluminum-copper, aluminum-magnesium, and aluminum-copper-magnesium alloys. Aluminum-lithium alloys promise superior crack propagation resistance when compared to the more traditionally used 2000 and 7000 aluminum alloys used in airframe construction. Alcoa’s 2090 series of aluminum-lithium alloys have been used on military aircraft and the NASA space shuttle to provide significant weight reductions, thereby allowing heavier payload capacities.16 The increased strength of aluminum-lithium alloys has promoted the use of less material for the same strength and safety margin compared even to advanced composites, and are employed up to 20% by structural weight on commercial aircraft such as the Airbus A350.Though more fatigue and crack resistant than traditional aluminum alloys, a number of disadvantages exist for aluminum-lithium alloys in aerospace applications. Due to the increased strength and crack resistance properties, aluminum-lithium alloys contain a reduction in ductility and retain a degree of anisotropic properties. Specifically, some lithium-aluminum alloys exhibit reduced fracture toughness in the transverse direction from rolling, and promote accelerated fatigue crack extension for structurally small micro-cracks.16 Despite the listed negative aspects of aluminum-lithium alloys, their use on aircraft promotes a 1.5 to 3 times increase in expected life, and can be manufactured, inspected, and repaired using conventional aircraft metal techniques. Due to their superior strength to weight ratio, aluminum-lithium alloys were afforded the same technology factor as advanced composite materials for the relative components.AluminumAluminum has been traditionally used in the aviation industry for up to 80% by weight of typical modern aircraft. Forged and machined aluminum components are relatively easier and cheaper to manufacture than their more modern material alternatives, and have decades of reliable data regarding their strength, fatigue, and fracture properties. Alloys typically used in commercial aircraft include 7075 for high stress applications due to its high ultimate strength, relatively low weight, ease of machining, anodization properties, and smooth finish. Traditional uses of aluminum plate, coils, and sheets will be employed on our aircraft in the absence of more technologically advanced materials due to their historically proven performance in harsh environments, reduced manufacturing and developmental costs, and established testing and repair techniques. It is also expected that the incorporation of a large percentage of traditional aluminum alloys will provide easier aircraft certification, and more reliable performance and airframe life data. As aluminum is the most traditionally used material on modern aircraft, no technology factor is incorporated for the aluminum aircraft components. Steel AlloysHeat-treated steel alloys were chosen for highly loaded components such as the aircraft’s landing gear. Containing the most stringent performance requirements, the landing gear is subject to severe loading conditions in a variety of environments. Today’s commercial and military aircraft use 300M, HP9-4-30, and newer AF-140 or AerMet 100 steel alloys. Additionally, the use of Ferrium S53 high strength stainless steel alloys have been proposed due to strength to weight ratio improvements and corrosion resistance over more traditional steel alloys.Latrobe Specialty Steel has recently licensed the production of Ferrium S53 as a high strength, high toughness, and corrosion resistant alloy suitable for aircraft landing gear. Ferrium S53 has also been labeled an environmentally friendly landing gear by the U.S. Air Force, having met the U.S. Air Forces requirement of a landing gear material not requiring a toxic cadmium coating.17 Latrobe’s S53 steel is expected to expand into the commercial aviation market in the foreseeable future, with other non-landing gear aerospace applications such as jet engine bearings on the near horizon.17 While a significant empty weight fraction reduction from the use of Ferrium S53 is not expected on our aircraft application, it is an attractive material for the landing gear components due to its superior toughness, corrosion resistance, and environmental friendly manufacturing techniques through the lack of toxic coatings. Additionally, currently manufactured Ferrium S53 landing gear provides a manufacturing cost reduction to titanium alloys, and is expected to decrease further in price as their use becomes more widespread.Material Choice BenefitsThe use of exotic materials on aircraft promises a significant reduction in empty weight fraction, however one must take into account the costs associated with each material, including developmental, manufacturing, and maintenance costs, ease of certification, and additional system weight requirements. We expect a more conservative structural weight reduction applied as a technology factor of 20% from the use of advanced composite materials and Lithium Aluminum alloys on the aircraft’s skin panels, nose, and leading and trailing edge surfaces. The material choices for major aircraft components are listed below in Table 9.Table SEQ Table \* ARABIC 9: Aircraft Component Material ChoicesComponentMaterialAdvantagesFuselage skins and wing stringersAluminum Alloys (Al-Li)Better Fatigue Crack Growth (FCG) performance reduces structural weightCanard, Control surfaces and wing skin panelsCarbon-Epoxy Composite LaminatesResistant to damage at high temperaturesNose, Leading and Trailing edgesCarbon-Epoxy Composite LaminatesHigh fracture toughness and yield strengthLanding GearSteel Alloy (Ferrium S-53)High strength, corrosion resistantWEIGHTS AND BALANCEAs discussed in past reports, the empty weight for the aircraft was calculated on a component weight buildup method using equations from the Transport Aircraft category of Daniel Raymer’s design textbook.8 For our final and best configuration; this led to the weights table listed below as REF _Ref260671573 \h Table 10.Table SEQ Table \* ARABIC 10: Component Weight TableComponentWeight (lbs)ComponentWeight (lbs)Gross Weight71,331Fuel System308Empty Weight38,079Flight Controls683Fuel Weight31,481Installed APU440Wing6,913Instruments262Canard598Hydraulics180Vertical Tail1,085Electrical943Fuselage7,224Avionics471Main Landing Gear873Furnishings3,054Nose Landing Gear103Air-conditioning199Nacelle1,628Anti-ice139Engine Controls90Crew800Engine Starter130Payload3,960CG Travel Diagram REF _Ref260242805 \h Figure 35 below shows the location of the aircraft’s center of gravity for various loading conditions. Four points in particular are listed in REF _Ref260242887 \h Table 11 and their corresponding locations are shown on the plot. Point A represents the maximum takeoff weight. Flying the design mission corresponds to a movement from point A to B, and the various points between A and B on the plot correspond to the endpoints of each mission leg. Moving from points B to C corresponds to the unloading of all passengers and crew. Moving from points C to D corresponds to the refueling of the aircraft, and moving from points D to A corresponds to the loading of all passengers and crew to full capacity. Over the course of aircraft operation, the center gravity travels in approximately an eight percent range of the total fuselage length.DBCAFigure SEQ Figure \* ARABIC 35: Aircraft Center of Gravity For Varied Loading ConditionsTable SEQ Table \* ARABIC 11: Four Loading Conditions and Center of Gravity PointsPointNumber of PassengersNumber of CrewFuel (lbs)A16437,750B1640C000D0037,750STABILITY AND WEIGHTSControl SizingCrosswindsQuite often a runway is oriented in such a way that landings must be made while the wind is blowing across the runway rather than parallel to it. When this is the case, additional complexities are introduced into the approach and landing process. If left uncorrected, the crosswind acting upon the aircraft will continually cause the plane to drift in the direction of the wind, making a conventional landing impossible. One method for counteracting the effects of the crosswind is the crab method. In this method, the pilot is required to yaw the aircraft into the wind at such an angle that the flight path remains aligned with the centerline of the runway, as seen in REF _Ref259992381 \h Figure 36. However, to ensure the pilot will have adequate control to prevent the continual drifting and handle crosswind landings, special considerations must be taken with regards to the vertical tail size. Figure SEQ Figure \* ARABIC 36: Crosswind LandingIn order to maintain a steady heading into the wind, the moment generated by the crosswinds must be opposed by a moment of equal magnitude generated by the deflection of the rudders on the vertical tail. The force each vertical tail can generate is a function of its coefficient of lift, its distance from the center of gravity, and its planform area. If everything other than the planform area is considered fixed, the resulting vertical tail area is the minimum size the vertical tail must be to adequately perform under crosswind conditions.One Engine InoperativeIn addition to crosswinds, the one engine inoperative case can potentially place large constraints on an aircraft’s minimum vertical tail size. The one engine inoperative case is characterized by the loss of one engine during flight which results in asymmetrical thrust. This asymmetrical thrust generates a yawing moment that must be balanced by an opposing force created by deflecting the rudders on the vertical tail, as seen in Figure 37.Figure SEQ Figure \* ARABIC 37: One Engine InoperableMuch like the calculations for crosswinds, the necessary force each vertical tail needs to generate an equal opposing moment will determine the minimum area needed for the vertical tails. Having sufficient vertical tail area will allow the aircraft to maintain a constant heading angle with only one engine operable.For this aircraft specifically, the one engine inoperable condition is less constraining than the crosswinds condition with respect to vertical tail sizing. The primary reason that the one engine out condition is less constraining is because the engines are mounted directly onto the fuselage near the centerline of the aircraft. Therefore, when one of the engine stops generating thrust, the resulting moment from the operable engine is still relatively small due to its small moment arm. With a smaller moment, a smaller amount of force from each vertical tail is required, and ultimately smaller vertical tails are needed. By using the vertical tail size as determined by the crosswind condition, the airplane will be designed to withstand both adverse flight conditions.Horizontal Stabilizer Sizing In the specific case of the canard design, the canard is the horizontal stabilizing surface. An equation from Raymer, shown below, was used as a preliminary sizing condition since the horizontal stabilizer was also a lifting canard and thus required a larger surface to provide lift.8 The final size was calculated by multiple iterations through the sizing code to find an ideal surface area. CHT-Tail Volume Coefficient (1.00 for Jet Transport)LHT-moment arm (quarter chord of wing to quarter chord of tail)SHT-surface area of horizontal tailC-main wing chordSw-surface area of the main wingControl Surface SizingThe aileron, rudder and elevator control surfaces were sized based on historical data provided in Raymer. This data is presented in REF _Ref260243219 \h Figure 38, and Table 12, shown below.8 By initially calculating the surface size of the wing, horizontal stabilizer, and vertical stabilizers, the ailerons, elevators, and rudders could be sized respectively. Figure SEQ Figure \* ARABIC 38: Raymer Figure Number 6.3Table SEQ Table \* ARABIC 12: Raymer Table Number 6.5Static MarginFinally, it was important to predict the stability of the aircraft by determining its static margin. This calculation required two values, the location of the center of gravity and the location of the aerodynamic center. The center of gravity was calculated using the weights and locations of each major piece of the aircraft. The aerodynamic center was less straightforward. For conventional configurations, the aerodynamic center can usually be approximated at the quarter chord with acceptable accuracy. For an aircraft with a lifting canard, the aerodynamic center will undoubtedly shift forward. Since no in-depth aerodynamic evaluation could be performed, it was assumed that the aerodynamic center of the canard configuration moved from the quarter chord of the main wing to the leading edge of the main wing. With this assumption, the center of gravity was calculated and the static margin was found to be 9.54% of the main wing chord. Research shows that many transport aircraft have a static margin of 5-10%. Therefore our predicted static margin value of 9.54% was within the normal range. It was also noticed that the center of gravity and static margin varied widely with small changes in the location of the engines. This indicates that a more thorough analysis of engine placement would be required during the preliminary design phase, in order to ensure acceptable stability characteristics.NOISEOne of the main N+2 goals is to significantly reduce noise. Because of that, Team 2 predicted the noise of our aircraft using methods and equations commonly used in acoustic engineering. In order to carry out simple noise estimation, a few assumptions were necessary. First, it was assumed that the primary noise source for the aircraft was the engine, and that the engine noise came from the fan and the jet exhaust. Additional information about the engine needed to be approximated, such as the volumetric flow rate through the core, the exhaust temperature, and the exhaust velocity. All necessary quantities were approximated based on research of similar engines and on simple cycle analysis results. After establishing these assumptions and approximations, the following procedure was used to predict the noise of the aircraft.Step 1 – Establish the power of each noise sourceIn this step, the power of both the engine fan and exhaust jet were calculated. The jet sound power can be calculated using the equation listed below.Pjet= εM5ρoV3A2 [Watts]In the above equation, ε is radiation efficiency, A is nozzle area, ρo is density of ambient air, M is Mach number, and V is flow velocity. Since the radiation efficiency is difficult to predict or approximate, an approximation for the entire term εM5 can be found from REF _Ref260243333 \h Figure 39 below.Figure SEQ Figure \* ARABIC 39: Radiation Efficiency Factor vs. Mach NumberObtaining the sound power of the fan is not necessary, as the noise estimation process for fans starts with the sound power level rather than just the sound power.Step 2 – Obtain the sound power level of each sourceFor the jet, the power value calculated in step one can be converted to a sound power level with the following equation. This equation uses the standard acoustic reference power level of 1x10-12 watts and accounts for the difference in temperature between the ambient air, Ta, and the exhaust temperature, T.Lw=SPL=10log (Pjet10-12)+ 20log (TTa) [dB]The sound power level of the fan can be predicted with a slightly more complicated scheme. First, the type of fan must be selected and the corresponding octave band sound power levels must be selected from REF _Ref260243429 \h Table 13 listed below.Table SEQ Table \* ARABIC 13: Band Sound Power for Varied Fan TypesA correction factor must then be calculated and added to all values taken from the above table. The correction factor, Δ, can be calculated according to the equation below, where Q is volumetric flow rate in cubic meters per second and p is the total pressure in kPa. ? =10logQ+20logp [dB]Notice that the last column of REF _Ref260243429 \h Table 13 is titled “BFI,” which stands for Blade Frequency Increment. This value is measured in decibels and must be added to the octave band level in which the frequency of the rotating blades lies. So after pulling the appropriate data from the table, then adding the correction factor to all octave bands based on volumetric flow rate and pressure, step three in this process is to add the BFI value to the single octave band that contains the rotating blade frequency. This frequency can be calculated according to the following equation, where n is fan speed in revolutions per minute and N is the number of blades.f= nN60 [Hz]The values obtained by summing steps one to three are the total sound power of the fan. In order to obtain the sound power level radiated from the inlet or outlet, the final step in the process is to subtract 3 dB. After completing all four steps, the radiated sound power level of the fan is found for each octave band. The individual octave bands can be summed into a single sound power level according to the following equation.LW,fan=SWLfan=10log(Σ10Loctave/10)Step 3 – Calculate the sound pressure level from the sound power level and distance from sourceWith the sound power levels of the fan and the jet known, the sound pressure level can be calculated based on the distance from the source. This calculation also assumed spherical wave propagation and includes the reflected wave from the ground since the noise sensors would likely be placed near the ground. The step also includes an estimation of the aircraft’s altitude 6000m after takeoff and 2000m prior to landing. These values were estimated using the sizing code and standard approach requirements. With all of the necessary information acquired, the sound pressure level for the fan was calculated with the following equation where r is the distance from the source, and will vary depending on the three standard measurement locations. Lp,fan=SPLfan= Lw,fan- 20logr-10log4π+3 [dB]Additionally, the factor of 4π is a result of the spherical wave propagation assumption, and the +3 dB correction accounts for the reflected wave.The equation for the exhaust jet is similar but includes another factor called the directivity index, which predicts a difference in sound pressure level based on the sensor’s angular position to the exhaust jet. The sound pressure level of the jet can be calculated from the equation below, and the directivity index, DI, can be found from Figure 40.Lp,jet=SPLjet= Lw,jet+ DI- 20logr-10log4π+3 [dB]Figure SEQ Figure \* ARABIC 40: Sound Directivity Index vs. Angle from Jet AxisStep 4 – Apply A-weighted noise correctionWith the sound pressure level of each source known, the decibel values must be adjusted to better represent the response of the human ear. A-weighted adjustments were used, which require that the dominant frequency be known. The fan frequency has already been calculated, and the jet frequency can found according to the following equation where α is the Strouhal number and assumed to be 0.15, V is the velocity, and D is the jet diameter.f=αVD [Hz]A-weighted adjustments can be made with the standard acoustic engineering adjustments provided in Table 14.Table SEQ Table \* ARABIC 14: Standard Acoustic Engineering AdjustmentsStep 5 – Calculate loudness measurementUsing a plot of equal loudness contours, the loudness of each source can be found based on its sound pressure level in decibels and dominant frequency in hertz.11 This is done quite simply by using the plot shown in Figure 41, with the decibel level along the Y-axis and the frequency along the X-axis, to find the loudness in Noy of each source from the approximate location between loudness curves.Figure SEQ Figure \* ARABIC 41: Band Sound Pressure vs. FrequencyIt is at this step in the process that the two sources, fan and jet, can be arithmetically summed to obtain a single numerical value representing the engine as a whole. Loudness values in Noy can be accurately summed to combine sources.Step 6 – Convert loudness values to Effective Perceived Noise LevelFinally, with the engine noise represented as a single loudness value measured in Noy, the Effective Perceived Noise Level (EPNL) can be found according to the following equation.12EPNL=33.3log (Noy)+40+10log t20+ FHere, Noy is simply the loudness value found in the previous step, t is the time in seconds at which the sensor is exposed to the noise source within 10 dB of its maximum sound pressure level, and F is a correction factor for pure tones, which is typically found to be +3 dB. The peak noise level of an aircraft flyby usually occurs over a very small time interval, so t used in calculation can be approximated by any value between 1 and 20 seconds. In this specific case, 20 seconds was used in order to maximize the predicted sound level and obtain a conservative estimate.Following the process outlined above, the noise of each engine considered was found. The results of this analysis are provided in Table 15, which shows the EPNL noise estimation of the geared turbofan and unducted fan. From this comparison, it is clear that the geared turbofan is considerably quieter than the unducted fan. This was one of the primary reasons why the geared turbofan was selected over the unducted fan.Table SEQ Table \* ARABIC 15: Engine Noise for Various Measurement LocationsGeared Turbofan (EPNL dB)Unducted Fan (EPNL dB)Sideline97102Takeoff9095Approach97100It is important to note that the above table lists the EPNL decibel levels for the uninstalled engine. This does not include the sound shielding devices used on our aircraft of the outboard vertical tails and low wing. Because of this, it was important to predict the noise of the installed engines. This was done by following the procedure outlined above, but subtracting 15 dB from the sound pressure levels calculated in step three. The subtraction of 15 dB was chosen because in acoustic engineering, the noise control method of placing a barrier between the source and the sensor can be approximated by a drop in the sound pressure level of 10 to 20 dB. The exact value would depend on the material of the barrier, the size of the barrier, and the distance from the source. The median value of 15 dB was chosen, and applying this correction yielded the installed noise estimation of the aircraft presented in REF _Ref260243963 \h \* MERGEFORMAT Table 16 below.Table SEQ Table \* ARABIC 16: Corrected Airplane Noise EstimateAirplane Noise (EPNL dB)Sideline87Takeoff80Approach87Total254COST PREDICTIONAcquisition CostThe cost estimating relationships that were employed to predict the development, manufacturing, and certification costs of this system were part of the Development and Procurement Costs of Aircraft model (DAPCA IV). This model was developed by the RAND Corporation for a variety of aircraft types, and is discussed in Daniel Raymer’s Aircraft Design: A Conceptual Approach.8 The model consists of a series of relationships that were determined from a statistical analysis of previous program costs. The DAPCA IV model predicts the aircraft’s research, development, testing, and evaluation costs (RDT&E); as well as the “flyaway” cost, which includes airframe production, and the integration of engines and avionics. The costs output by the model were in constant 1999 dollars, and were converted to 2009 dollars using the U.S. consumer price index. To apply the relationships in the DAPCA IV model, seven aircraft specification inputs were required: empty weight, maximum velocity, production quantity, number of flight test aircraft, total number of engines needed for fleet, the cost of each engine, and the cost of avionics. The model uses these inputs to predict the number of labor hours required for engineering, tooling, manufacturing, and quality control; and the direct costs associated with development support, flight test, and manufacturing materials. From these predicted values, the total cost of developing and producing the aircraft was estimated by taking into account the costs of labor. The values used as the inputs for the cost model were determined using the team’s design specifications, data from previous development programs, and market forecasts. The aircraft’s empty weight (38,100 lbs.) and required number of engines (2) were readily available details of the team’s concept. The maximum velocity was approximated as 527 knots based on values from similar aircraft. In addition, since the number of flight test aircraft typically used in past programs ranged from two to six, three was chosen as a conservative number for a small volume production run.8 The cost of avionics was estimated by assuming a price of approximately $3,000/lb and using the estimated weight of the avionics on board (471 lbs); the cost per pound was based on typical values supplied in Raymer.8 The number of aircraft to be produced was estimated using the business aircraft market forecast supplied by Honeywell.2 This forecast predicted a relatively stable demand for ultra-long range (ULR) business jets with moderate growth in the high-speed ULR segment. With these market conditions in mind, the production runs of the Gulfstream G550 and Bombardier Global Express family were studied since these aircraft were within the same category and had been in production for more than 5 years. These production runs were in the neighborhood of 180 to 250 airframes. Therefore, it was decided to use 150 airframes as a conservative value for cost estimation. This value accounted for the projected stability of the market, as well as the impact of being a new entrant.The cost of each engine was approximated using a separate cost estimating relationship (CER) originally developed for turbojet engines.8 This CER approximated the production cost of an engine from its maximum thrust, maximum operating Mach, and turbine inlet temperature. The maximum thrust used was the value that had been determined by the team’s design (11,900 lbs). The maximum operating Mach for the engine was approximated as Mach 0.95, and the turbine inlet temperature was approximated as 2600? Rankine; which was based on typical values for modern turbofans. In addition, a correction factor of 1.2 was used to compensate for the increased cost of production of a Geared Turbofan (GTF) relative to a turbojet. This resulted in an estimated engine production cost of $3.6 million per engine (2009 dollars).Using all of these specifications, numbers, and costs as inputs for the model, the result was an estimated purchase price of $49.7 million per aircraft with a total program cost of $6.71 billion (2009 dollars). Aside from summing all of the costs included in the model, an investment cost factor of 1.1 was also used to account for the cost of money and the team’s profit.8 The projected price of the design was compared with comparable aircraft currently on the market, and appeared to be a reasonable estimate. The Gulfstream G550 and Bombardier Global Express XRS had new list of prices of $49 million and $48 million, respectively.6Operating CostsAfter an approximate purchase price had been determined, the operating cost of the aircraft was estimated. The cost of operating the aircraft was broken down into fuel costs, crew salaries, maintenance labor, maintenance materials, insurance, and depreciation. Each of these costs were estimated separately and then used to determine an overall cost per flight hour. Fuel costs were estimated from the aircraft’s expected fuel flow and utilization. An approximate fuel flow of 2,300 pounds per hour was determined using the aircraft’s design mission, predicted fuel weight, and a correction factor for off-design performance. The utilization of the aircraft was then approximated as being 500 flight hours per year with 200 cycles. This was the lower bound of the range of flight hours provided in Raymer, and was thus expected to yield conservative hourly cost figures.8 In addition, the number of cycles used implied an average flight duration of 2.5 hours, since the typical operating mission of the aircraft was expected to be domestic. Next, the price of jet fuel was considered. After reviewing the current price of Jet A at numerous airports, a price of $4.50 per gallon was chosen for the purposes of cost estimation. The resulting yearly fuel cost was $755,000, assuming 500 flight hours per year, or $1,510 hourly.Crew salaries were approximated from a cost estimating relationship developed using data from Boeing.1 This relationship estimated crew cost per block hour for a three person crew in 1999 dollars. The annual number of block hours for the aircraft was determined by assuming that an additional 20 minutes would be spent on each flight taxiing and complying with air traffic control procedures. The crew cost per block hour was then estimated from the appropriate CER, which took into account the aircraft’s cruise velocity (487 knots at altitude) and gross takeoff weight (71,300 lbs). The resulting crew cost was $537 per block hour, or $268,000 a year (2009 dollars).Maintenance expenses were estimated using the expected maintenance requirements of the aircraft, the cost of labor, and CERs. All maintenance to be performed on the aircraft was included in an average figure of Maintenance Man hours per Flight Hour (MMH/FH). An MMH/FH of 3 was assumed for this design, which was in the range considered typical for a business jet.8 The number of maintenance man hours per year was then estimated from the MMH/FH and the expected number of flight hours (500). The cost of a maintenance man hour was then assumed to be approximately equal to the hourly manufacturing wrap rate. This resulted in an estimated maintenance labor cost of $282 per flight hour. The cost of maintenance materials was then estimated using CERs for civilian aircraft provided in Raymer from RAND.8 These CERs predicted the material cost per flight hour and per cycle using the estimated cost of the airframe, the cost of the engines, and the number of engines on each aircraft. The result was an estimated maintenance materials cost of $619 per flight hour.Depreciation and insurance costs for the design were estimated from data gathered on similar aircraft models. Depreciation was approximated as a linear schedule. After reviewing the asking prices of numerous ULR business jets on the used market, a depreciation rate of 10% per year was selected. This rate would be close to the amount expected within the first 5 years of ownership, but would be considered a fairly aggressive estimate after that. Insurance was estimated assuming that the aircraft would be insured for its full hull value, at a rate of 0.32 percent. This rate was taken from a cost evaluation of the Gulfstream G550 performed by Conklin & de Decker.9Adding together the estimated costs of fuel, crew, maintenance, insurance, and depreciation, an operating cost of $8,500 per flight hour was calculated (2009 dollars). Without depreciation, the predicted cost was $3,400 per flight hour. For comparison, a detailed cost analysis of a Gulfstream G550 performed by Conklin & de Decker predicted a total cost of $5,500 per flight hour excluding depreciation, with an assumed utilization of 400 hours per year.9 Although this is significantly more expensive, the estimates for the G550 used an average fuel burn of 3,300 pounds per hour; which is significantly higher than the fuel burn projected for this team’s concept. In addition, the cost analysis completed here ignored the impact of hangar rent, landing fees, catering, and crew training on operating costs.SUMMARYThe final concept of our designed aircraft can be seen in REF _Ref260244118 \h Figure 42. The concept is a high speed, long range aircraft which is also environmentally friendly. The aircraft has lifting canards, two vertical tails, geared turbo fan engines and spiroid wingtips. Figure SEQ Figure \* ARABIC 42: Final Concept DesignThe requirement compliance matrix associated with the aircraft is shown in REF _Ref260244208 \h Table 17, and was used to verify that all the requirements of our customers were met. Table SEQ Table \* ARABIC 17: Requirements Compliance MatrixPerformance CharacteristicsTargetThresholdCurrentStill Air Range7100 nm6960 nm7100 nmMTOW Takeoff Ground Roll4000 ft5000 ft3900 ftMax. Passengers16816Volume per Passenger per Hour (Design)13.3 ft3/(pax?hr)2.28 ft3/(pax?hr)20.7 ft3/(pax?hr)Cruise Mach0.850.80.85Initial Cruise Altitude41000 ft40000 ft41000 ftCumulative Certification Noise Limits274 dB274 dB254 dBCruise Specific Range0.3 nm/lb0.26 nm/lb0.31 nm/lbLoading Door Sill Height4 ft5 ft5 ftOperating Cost$4100/hr$4300/hr$3400/hrBased on REF _Ref260244208 \h Table 17, the majority of the requirements in the requirement compliance matrix were met with the only one requirement not meeting the target value, which was the loading door sill height. However, for this height, the target value of 5 feet was met. In addition to the above requirements, it was also important to our team to meet the NASA N+2 goals if possible. As seen in the table below, none of the NASA N+2 goals were completely met. The NASA N+2 goals were difficult to meet due to their overly ambitious targets. In order to meet these goals, more development would be needed for engine performance, aerodynamic design, and materials used.Table SEQ Table \* ARABIC 18: Compliance with NASA N+2 GoalsCriteriaGoalOur AircraftAchievedNoise-42 dB below Stage 4-20 dBNoEmissions-75%-50%NoFuel Burn-40%-25%NoTakeoff Field Length-50%-33%NoFurther work would be required to develop the maturity of this concept to a level that is acceptable for the preliminary design phase. This work would include more detailed aerodynamic and structural analyses, and a more thorough approach to predicting engine performance. The aerodynamic analysis that was performed for our aircraft was predominantly conceptual and theoretical, and more experimental data would be helpful. More in depth detail for a final design would be needed, using tools such as computational fluid dynamics tools, and wind tunnel testing. The engine performance could also be more detailed through further analysis and a more accurate engine deck could be built. Further analysis on the structural design, its load paths, and the materials used needs to be incorporated. Sizing of the spars and ribs along with fatigue and temperature analyses would need to be done using finite element modeling to evaluate the stresses and strain present in the aircraft. With this additional work, the initial design phase could be completed. REFERENCES1 "Avionics Magazine :: Outlook: High Hopes for General Aviation." Breaking News and Analysis on Aviation Today. Web. 11 Feb. 2010. < "Honeywell Aerospace Business Aviation Outlook Forecasts $200 Billion inGlobal Business Jet Sales Through 2019." Web. 11 Feb. 2010. <, Egbert. Synthesis of subsonic airplane design an introduction to the preliminary design, of subsonic general aviation and transport aircraft, with emphasis on layout, aerodynamic design, propulsion, and performance. Delft: Delft UP, Nijhoff, Sold and distributed in the U.S. and Canada by Kluwer Boston, 1982. Print.4Great Circle Mapper. Web. 11 Feb. 2010. <“Subsonic Fixed Wing Project”. NASA. 08 February 2010. 's All The World's Aircraft. Web. 11 Feb. 2010. < Norris, Guy. “Rotor Revival”. Aviation Week & Space Technology. 14 December 2009. pages 54-57. 8 Raymer, Daniel P. Aircraft Design A Conceptual Approach (Aiaa Education Series). New York: AIAA American Institute of Aeronautics & Ast, 2006. Print.9 Conklin & de Decker, “The Aircraft Cost Evaluator: Gulfstream G550.” Conklin & de Decker Aviation Information, Fall 2005.10 Raichel, Daniel R. The Science and Applications of Acoustics. New York: Springer, 2000. Print.11 "Effective_Perceived_Noise_Level." SFU.ca. Web. 15 Apr. 2010. <;. 12 "Perceived_Noise_Level." SFU.ca. Web. 15 Apr. 2010. <;. 13“Airfoil Investigation Database” Airfoils. 30 Jan 2010. <, Kenneth W., “Low Speed Aerodynamic Characteristics of Rectangular SlottedSupercritical Airfoil Having Several High Lift Flap Systems” NASA Technical Memorandum, Langley Research Center, August 197115Sun, C.T., Mechanics of Composite Materials and Laminates, School of Aeronautics and Astronautics, Purdue University, West Lafayette, Indiana, 200116 “Cast Nonferrous: Aluminum-Lithium Alloys” – Key to Metals<Article58.htm>17 “Latrobe Specialty Steel Licenses Corrosion-Resistant Landing Gear Alloy”-Red Orbit<; ................
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