SYSTEM DESIGN REVIEW



SYSTEM DESIGN 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 2010EXECUTIVE SUMMARYTeam Two is designing a business jet aircraft that is both environmentally responsible and meets the needs of a selected target customer. These two major drivers for the design of the aircraft were examined in the previous System Requirement Review document which yielded the main requirements of a long range, high cruise speed, and low environmental noise and emissions aircraft. Many possible designs and advanced technologies were analyzed in order to develop an aircraft that best meets these requirements. The initial aircraft designs considered included a hybrid wing aircraft, several conventional aircraft with different engine placements and tail configurations, and canard aircraft with aft mounted wings and various engine and vertical tail placements. These aircraft were evaluated and compared until the two best were ultimately selected. One of these remaining aircraft was a conventional aircraft similar to a G550 except using an advanced engine and some composite materials. The second was a canard aircraft with an aft mounted swept back wing using the same advanced engines and composite materials.In order to meet the NASA N+2 environmental goals, drastic changes will have to be made in comparison to modern business jet aircraft. To do this, advanced technologies were evaluated for their potential use in 2020. Two new engine types that are currently in research phases were examined: an unducted turbofan and a geared turbofan. Using one of these two engines, and the addition of light weight materials, may provide enough benefit to meet the goals. Current calculations show that the conventional aircraft with the advanced technologies will weigh about 71,000 lbs and the second design using canards and an aft wing, also with the same advanced technologies will weigh 73,000 lbs. Both aircraft have been determined to have adequate static stability. Further analysis will be done to determine more precisely which of the two aircraft designs and which of the engines will best provide an environmentally responsible aircraft to the customer.Table of contents TOC \o "1-3" \h \z \u MISSION STATEMENT PAGEREF _Toc257711967 \h 1DESIGN AND OPERATING MISSION PAGEREF _Toc257711968 \h 1AIRCRAFT CONCEPT SELECTION PAGEREF _Toc257711969 \h 2ADVANCED TECHNOLOGIES PAGEREF _Toc257711970 \h 13SIZING CODE PAGEREF _Toc257711971 \h 16ENGINE SELECTION PAGEREF _Toc257711972 \h 30SUMMARY PAGEREF _Toc257711973 \h 39NEXT STEPS PAGEREF _Toc257711974 \h 39REFERENCES PAGEREF _Toc257711975 \h 41APPENDIX A PAGEREF _Toc257711976 \h 42APPENDIX B PAGEREF _Toc257711977 \h 46APPENDIX C PAGEREF _Toc257711978 \h 50MISSION STATEMENTIt is the mission of Team 2 to be the primary systems integrator of a cost effective, long range, environmentally friendly business aircraft that is able to serve as a high speed transportation solution for its customers. The design will have a reduced impact on the environment by meeting NASA’s N+2 criteria while still remaining competitive in the ultra-long range business jet market. DESIGN AND OPERATING MISSIONThe 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. The design mission consists of eight mission legs between nine points as illustrated in the following figure.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. From points 1 to 2 is the climb portion of the mission where the aircraft climbs at best rate to an altitude of 42,000 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, climb to an altitude of 5000 feet at best rate climb, points 5 to 6, and commence cruise to an alternate airport 200 nautical miles away. Once 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 when it comes to a stop.While the design mission is the optimal, most efficient use of this aircraft, several other operating missions can be made by this aircraft as well. One such operating mission would be flying from New York to Van Nuys. The distance between the two cities, which is 2146 nautical miles, falls well within the maximum still-air range of 6350 miles. To compensate for the largely unused range, the aircraft can then be flown at its maximum Mach number of 0.9 at a maximum capacity of 16 passengers. This range tradeoff allows for tremendous flexibility in speed and capacity for shorter ranged flights.AIRCRAFT CONCEPT SELECTIONFollowing the system requirements phase of design, concept generation and selection was required to further the design process. After great thought and consideration, the team was able to generate eight initial designs, as shown in Figure 2, for further analysis. Rather than subjectively choosing one of the eight concept designs, the team implemented Pugh’s method as a means to add objectivity to the decision making process. The first step was to establish a set of criteria against which all of the proposed aircraft would be evaluated. The house of quality that was constructed during the design requirements phase helped in the establishment of evaluation criteria. This provided criteria that were centered on the needs of the customer and not the desires of the designers. Ultimately, ten different criteria, as listed in Table 1, were used in the first iteration of Pugh’s method. Figure SEQ Figure \* ARABIC 2: Eight Proposed Aircraft DesignsThe next step of Pugh’s method was the creation of a matrix relating the established criteria to each of the proposed designs. The matrix’s purpose was to help clarify how each aircraft concept rated under each criterion relative to a benchmark aircraft. Based on the selected benchmark aircraft, all other concepts were evaluated on whether or not they were better, worse, or equivalent to the benchmark aircraft for a given criterion. These relationships were illustrated in the matrix using the symbols ‘+’, ‘-’, and ‘s’ respectively. Each decision was made with cursory calculations whenever possible, relying on the team’s best collective engineering judgment when not possible.After evaluating each aircraft in the matrix against every specified criterion, the total number of positive, negative, and equivalent marks were separately compiled in order to maintain evaluation details. For instance, if positives and negatives were allowed to cancel each other out, it would be unclear whether or not an aircraft received a score of zero due to average qualities or because all of its positive aspects were offset by negative ones. Evaluating the positives and negatives allowed the team to clearly define the favorable and unfavorable aspects of each individual aircraft. The matrix comparing the eight initial designs to the ten criteria is shown in Table 1.Table SEQ Table \* ARABIC 1: Concept Selection, Pugh's Method 1st IterationAfter studying the different ratings given to each aircraft, the team decided to repeat Pugh’s method using the top four aircraft designs determined by the first iteration. This was done because the first iteration of Pugh’s method did not define a single best design for the given mission, but rather judged each aircraft qualitatively against the chosen benchmark. The benchmark aircraft clearly cannot receive a positive or negative rating for any criterion, since it is equivalent to itself. Thus, two iterations of the method were completed, each with a different benchmark aircraft.Examining the outcomes of the first Pugh’s method iteration, it was decided that the canard equipped aircraft could offer a potential advantage in cruise drag over conventional tube and wing designs. The overall weight of each design was compared by counting the number of control surfaces used and observing the general structural complexity of the configuration. No material considerations were incorporated into these decisions as they could apply to any of the designs. An increased number of surfaces, or additional structural requirements such as those found on a blended wing body, were deemed less favorable for overall weight than a conventional aircraft configuration of the same size.In order to meet the project’s goal of having a reduced environmental impact, it was desired that the airframe be able to utilize a variety of high speed engine technologies, including unducted fans. To this end, aircraft with rear pylon mounted engines similar to the baseline model were assigned equivalences to the baseline, whereas designs that would place the engines in locations interfering with the cabin from vibration, noise, and safety perspectives were assigned negative ratings. Specifically relating cabin noise relative to the baseline design, a negative rating was assigned to engine locations in line with the center of the main cabin area.Further in regard to addressing environmental impact, the environmental noise of each proposed design was considered. As every design was evaluated with the possibility of using a variety of engine types, the aircraft structure was the primary consideration for this criterion. Designs providing noise shielding due to wing placement, vertical stabilizer placement, or having shielding due to a blended wing configuration were all given positive ratings relative to the baseline. The placement of each design’s landing gear was also considered. Designs requiring excessively long and robust landing gear were given a negative rating, whereas the blended wing body design with a structural framework capable of mounting the landing gear in a variety of locations was assigned a positive rating relative to the baseline.Window placement and an unobstructed cabin view were also evaluated for each design, and aircraft with aft-mounted wings provide a clear window view, which was deemed superior to the baseline configuration. Difficulties associated with placing an abundant number of windows in a blended wing body configuration resulted in a negative rating for the design relative to the baseline. Though a subjective and hotly debated criterion, each design’s attractiveness and visual appeal was also considered. By majority vote, negative values were assigned to the high wing design, as well as aircraft with less sleek lines. The team agreed that canard designs with aft wing sweep and more futuristic configurations were more visually appealing than the baseline configuration. In addition, it was decided that the forward wing sweep included in design one made it less visually appealing. Each of the eight designs were investigated in relation to two remaining criteria, the first of which was the complexity of cabin pressurization. Every design with the exception of the blended wing body contained a relatively easy to pressurize cylindrical fuselage and these designs, therefore, received equivalency ratings. The blended wing body was assigned a negative value relative to the baseline due to its non-cylindrical shape. Finally, static margin and stability were considered for each design. It was decided that canard aircraft, with the exception of the tri-surface aircraft, would provide a smaller static margin and contain potential center of gravity challenges, which earned these designs a negative value relative to the baseline. The limited historical data available suggests center of gravity difficulties and a small static margin on existing blended wing body designs, thereby also providing it a negative value relative to the base design. The tri-surface aircraft was given a positive value in relation to static margin, as data available for the minimalistic number of tri-surface aircraft suggested an ability to design for a larger static margin.After evaluating each of the eight designs relative to the initial ten criteria, design numbers two, four, six, and seven were chosen as the best designs and were used in the next iteration of the method. Out of these four designs, number six was chosen as the new benchmark. Design six was chosen because it was very dissimilar to the remaining aircraft. It was important to choose a new baseline model so that the outcomes from the second iteration would not be identical to the first, and would allow the previous baseline, design four, to be rated. It was noted during the first iteration of Pugh’s method that there were no criteria related to development cost. In order to improve the value of the second iteration, development cost was added to the matrix for the four remaining designs, as shown in Table 2. Each design was then reevaluated for every criterion based on the new benchmark aircraft. Ratings were assigned in a manner similar to the first iteration, with the same considerations taken into account. Table SEQ Table \* ARABIC 2: Concept Selection, Pugh's Method 2nd IterationUsing the results of the second Pugh’s method iteration, two design choices were selected as best meeting the evaluation criteria and requiring further analysis. Designs four and seven had the fewest number of negative ratings, with design seven having the most positive ratings during the second iteration. Since the goal was only to narrow the field of designs down to two options for additional analysis, further iteration was avoided and designs four and seven were selected. Designs four and seven, selected by the Pugh’s method process, are henceforth referred to as concepts one and two, respectively, and are shown in Figure 3. Figure SEQ Figure \* ARABIC 3: Concept One, Formerly Design Four (left) and Concept Two, Formerly Design Seven (right)Concept one, the first iteration’s baseline design, presented a number of advantages that caused it to be worthy of further investigation. Concept one is configured similarly to a conventional business jet, thus providing a pre-optimized platform for technological improvement and analysis. Through the evolution of aircraft design, this conventional configuration has been refined to the extent that a reinvention of the wheel may not be required to meet the desired end goals. In other words, the application of new technologies and additional improvements to current aircraft configurations may be able to meet the required design goals with minimal developmental costs and a conservative amount of risk. Concept one requires no additional control surfaces, and the aft fuselage engine mounting allows for a wide variety of engine choices. In addition, this mounting location would facilitate an unducted fan by keeping internal cabin noise and vibration to a minimum. Concept one’s T-tail configuration, which can be seen in Figure 4, can provide a necessary amount of horizontal stabilizer and elevator surface area while still being able to facilitate an aft mounted unducted fan. Additional benefits of concept one include a low wing design that allows for minimal spar interference with the main cabin, shorter and lower weight landing gear struts, ease of maintenance, and can serve as an emergency exit platform. Concept one’s circular fuselage also facilitates cabin pressurization and eliminates the need for any special structural reinforcements. A circular cross section also allows for a large interior cabin volume and adequate storage area with no increased development costs. Concept one also provides adequate measures of stability, and incorporates the wide static margin common to current business jets. A large number of windows can be incorporated into this concept, and its appearance is both attractive and familiar to the end user. Figure SEQ Figure \* ARABIC 4: Concept One Three ViewConcept two also presented a number of advantages warranting further analysis. Sharing some conventional aircraft characteristics, but capitalizing on many innovative configuration changes, concept two promises potential improvements effectual of a more pioneering design toward the system’s overall goals. Similar to concept one, concept two may incorporate several new technologies to meet the desired mission goals, but also requires higher development costs and is associated with a higher amount of risk. As in concept one, concept two features an aft fuselage engine placement that enables a wide variety of engine choices while reducing internal cabin noise. Though requiring additional structural weight, concept two’s design specifically addresses environmental noise through the placement of wings below the entirety of the engines. Lateral engine noise is also shielded using twin outboard-located vertical stabilizers. Again as in concept one, concept two shares the same advantages of a low wing configuration with a circular fuselage. The same material technology factors that can be utilized on concept one are equally applicable to concept two. Unique to concept two is the potential cruise drag reduction associated with a lifting canard, which can benefit the environment through reduced fuel consumption. Concept two’s development costs promise to be higher than concept one, but are again variable based upon the degree of improvement desired and number of new technologies incorporated. The use of a lifting canard requires a more aggressive design process with regard to stability, and may potentially contain a smaller static margin. Window placement in concept two is similar to concept one, however the cabin view through the windows in concept two provides a clear view of the ground from nearly the entire cabin. Concept two’s swept wings, lifting canards, and twin vertical stabilizers provide it with a futuristic yet familiar set of lines that are visually stimulating to end users, as shown by the conceptual sketches in Figure 5.Figure SEQ Figure \* ARABIC 5: Concept Two Three ViewADVANCED TECHNOLOGIESThe two advanced technologies that were investigated were an unducted propfan for a propulsion system and use of composite materials for the structure of the aircraft. The reduction in fuel consumption is just one of the many goals proposed by NASA’s subsonic fixed wing program. NASA has set four fundamental goals referred to as N+2 for a subsonic fixed wing business aircraft set for production in the 2020 timeframe. The four goals are to achieve a noise reduction down to 42 dB below stage 4 certification, 75% reduction in NOx emissions, 40% reduction in fuel consumption, and a performance field length reduction of 50%.1 The first three design goals will hopefully be met by using an innovative propulsion system. An unducted propfan is an advanced engine design that would incorporate two counter rotating fans that would be directly connected to the engine’s turbines. General Electric explored the unducted propfan concept in the 1980s and 1990s and even flew a design named the GE36 on a Boeing 727 test aircraft.2 The project had problems with noise levels and vibration due to the wave drag created by the high speed fans. The noise levels and plummeting fuel costs of the 1990s caused the cancellation of the project. Currently, Rolls-Royce, General Electric, and NASA are working together to achieve a noise reduction level in the unducted propfan concept. Robert Nuttal, Rolls-Royce’s vice president of future programs strategic marketing, says that “The outcome of this work is that we are now confident that open-rotor-powered aircraft will be quieter than any equivalent aircraft flying today and that it will comfortably meet Stage 4 noise legislation.”2 The unducted propfan technology is the most promising to meet NASA’s N+2 goals within the specified time frame. In fact, NASA’s environmentally responsible aviation program (ERA) is devoting much of its research in the subsonic fixed wing project to the unducted propfan technology. Guy Norris states “ERA is focused on the goals of NASA’s N+2, a notional aircraft with technology primed for development in the 2020 time frame as part of the agency’s subosonic fixed-wing program.”2 Nuttal was quoted in Aviation Week’s December 14, 2009 issue as saying, “So far the GE-NASA experience seems to echo that of Rolls-Royce. We are able to confirm that the fuel burn will be 25-30% better than today’s products. And, because of the engine cycle of the open rotor; the nitrous oxide will be 20% lower than another engine with an equivalent combustor technology. We are now preparing for the next tranche with the next build of the rig taking place in Q2 2010”.2 Because of the current development being made on the concept and NASA’s faith in the technology, the team feels it is an appropriate decision to anticipate using unducted propfans as a propulsion system for the design project. On the other hand, there is current engine technology from Pratt & Whitney in the form of a geared turbofan that shows promise. More importantly, the technology is slated for production before the 2020 production date for N+2. 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- 24,000 lbf depending on engine selection. The thrust range is ideal for our aircraft and allows the team to continue using the rubber engine assumption in sizing.The other possible advanced technology will be 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 designs that are currently in production. The technology is likely to keep improving and deliver 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.SIZING CODEConstraint DiagramsTwo separate constraint diagrams, one for each design concept, were used in this portion of the design process. The constraint diagram associated with the conventional configuration changed very little since the System Requirements Review; the only significant changes were updating plane geometry inputs such as aspect ratio and adjusting landing and takeoff distances through design iterations. A second constraint diagram was developed for the canard pusher configuration. All the necessary plane sizing inputs were changed, but many values remained the same. For example, values for CLmax at landing and takeoff are the same for both concepts because we have assumed that these values are limited primarily by the flaps included on the plane; both configurations currently use single slotted flaps. Future analysis will consider adding Fowler flaps to determine if they are beneficial. Our design group currently does not definitively know how adding a canard will change many of the values associated with the canard pusher configuration. Because of that, many values have been assumed to be the same as the conventional configuration until further differentiation can be achieved. REF _Ref257359329 \h Table 3 lists some inputs for both configurations.Table SEQ Table \* ARABIC 3: Aircraft Input ValuesW0/STSL/W0(CL)maxT/O(CL)maxLandingCD0eARConventional76.331.52.0.018.88Canard-Pusher84.331.52.0.018.810.5Some values, such as CD0 and SFC, are calculated many times in the sizing code at different flight velocities and thrust settings. Because of that, theses values vary widely depending on the location of interest in the flight mission. The corresponding values reported in the above table are representative of the appropriate flight condition.It is also clear from this table that both configurations have identical input values for some significant variables. This is based on an early design-phase assumption necessitated by a lack of knowledge that both configurations will have similar constraining variables. It is anticipated that many more differences will exist between the two models as progress is made in how to mathematically and systematically represent the differences between the two configurations. Also the engine type will be differentiated between the geared turbofan and the unducted turbofan jet engines. Both have varying benefits with regards to fuel efficiency, noise, and emissions. Future additions to the sizing code will attempt to quantify these variations of the two engines to see which is most beneficial overall to both aircraft concepts. REF _Ref257648270 Figure 6 below shows the current constraint diagram for the conventional configuration. This model is limited by second segment climb and landing ground roll. REF _Ref257648293 Figure 7 shows the current constraint diagram for the canard pusher configuration, which is limited primarily by landing ground roll, takeoff ground roll, and second segment climb. The first two columns of REF _Ref257359329 Table 3 above summarize the initial starting points used in the code for the thrust to weight ratio and wing loading. The wing loading for both configurations has changed since the system requirements review, in which the wing loading was estimated at 100 lb/ft2. Both values have dropped based upon research of exiting planes and because of iteration results in the design process.Figure SEQ Figure \* ARABIC 6: Constraint Diagram for Conventional ConfigurationFigure SEQ Figure \* ARABIC 7: Constraint Diagram for Canard Pusher ConfigurationSizing MethodsA detailed Matlab script has been developed in order to size both configurations. This code is used to estimate the size of the aircraft required to complete our design mission, which is provided in Figure 1. The code takes in approximately one hundred inputs for each configuration, utilizes a drag calculation at each step of the mission, predicts the empty weight with a component build up method, and incorporates an early but detailed engine model. While the code is functioning and provides reasonable estimates, it is under constant revision in an attempt to improve estimations, to make assumptions more accurate, and to further differentiate between the two configurations. The general process of the sizing code is shown below in REF _Ref257359578 \h Figure 8, which displays the solution process graphically in a flow chart.Figure SEQ Figure \* ARABIC 8: Flow Chart for Sizing CodeAfter every significant change, the sizing code is calibrated twice, once for each configuration. The conventional model is calibrated against the Gulfstream G550 and the canard pusher model is calibrated against the Beechcraft Starship. Both of the chosen calibration models were used because their configurations closely match the design configurations our group has chosen. Calibration is carried out by running our sizing code with all inputs describing the corresponding calibration aircraft. The code’s predictions of empty weight, fuel weight, and gross weight are then compared to published data for each calibration aircraft, and the appropriate calibration factors are calculated. The calibration factors for the G550 are used in sizing our conventional model and the calibration factors for the Starship are used in sizing our canard pusher model. The final calibration factors are listed below in REF _Ref257359659 \h Table 4.Table SEQ Table \* ARABIC 4: Final Calibration Factor ValuesCanardConventionalEmpty Weight0.770.77Fuel Weight1.331.28Gross weight1.051.01The calibration factors shown above are larger than desired. This is most likely caused by two significant sources of error. The first issue in the current sizing code is that the component-build-up equations used best describe cargo and transport planes from the 1970’s. This will likely cause some degree of error, in that a business jet in 2020 will have many significant differences. Additionally, the engine model in use attempts to scale a 50,000 lbs thrust engine down to approximately 15,000 lbs of thrust. This process, described later in the report, will also likely cause some degree of error. As the code is continually refined and made more and more accurate, it is hoped that the calibration factors will decrease in magnitude and become continually closer to one.It is also important to note that both calibrations yield the same value for the fuel weight. This is because the Starship, which is used to calibrate the canard pusher configuration, uses a turboprop engine while our code predicts the engine performance of a turbofan or unducted fan. Because the turboprop engine is so fundamentally different from what the codes analyzes, the G550 calibration is carried out first, and it is assumed that the fuel weight calibration factor for the Starship will be the same as the G550. So after the Starship is calibrated, its fuel weight calibration factor is replaced with the fuel weight calibration factor of the G550.The inputs for each of our two configurations are provided in Appendix A and Appendix B. The variables have been sufficiently commented such that the variables and their values should be self explanatory. A final weights table for all components of our two planes has been provided below as REF _Ref257359761 \h Table 5.Table SEQ Table \* ARABIC 5: Final Weight Values?Canard (lbs)Conventional (lbs)?Canard (lbs)Conventional (lbs)Wing5,3676,232Flight Instruments260254Canard/Horizontal Tail728753Hydraulics179171Vertical Tail879662Electrical System9431,125Fuselage7,4457,392Avionics471394Main Gear922986Installed Avionics359359Nose Gear107162Air Conditioning171466Nacelle1,6281,656Anti-Ice147145Engine Controls9098Crew800800Engine Starter130130Payload1,7601,760Fuel System352360Fuel Weight27,30028,383Flight Controls683766Empty Weight39,74241,257Installed APU440275Takeoff Gross Weight72,94471,365Stability and ControlIt is important for any business jet to be statically stable. Even at the early design stage, our sizing code predicts the static stability of each configuration very simply by calculating the static margin. The static margin measures the distance between the aerodynamic center and the center of gravity of the airplane and normalizes this distance relative to the chord of the wing. In order to be statically stable, the static margin must be positive. The sizing code calculates the static margin according to the equation listed below.SM= XAC-XCGCmacThis calculation is carried out as one of the final steps in the code because the final sizes and locations of each component must be known to accurately find the center of gravity. If the location of any component cannot be calculated, it is given directly as an input. The aerodynamic center for the conventional model is assumed to be at the quarter chord. The aerodynamic center for the canard pusher is slightly more complicated since it is not explicitly known how much the canard will affect its location. The effect of the canard on the aerodynamic center will also depend greatly on whether or not the canard is a lifting canard. For now, we have assumed that our canard pusher configuration includes a lifting canard, and that this will cause the aerodynamic center to shift from the quarter chord point of the main wing to the leading edge of the main wing. With these assumptions, the static margin and center of gravity have been calculated for each configuration. These values are listed below in REF _Ref257359888 \h Table 6. It is important to note that the center gravity is listed as a percentage of the fuselage length while the static margin is listed as a percentage of the chord of the wing. As expected, the center of gravity for the canard pusher model is much further aft. Because of this, it also makes sense that the static margin is smaller for the canard pusher model.Table SEQ Table \* ARABIC 6: Static Margin and Center of GravityCenter of Gravity (% Fuselage Length)Static Margin (% Cmac)Conventional51%37%Canard74%29%Additionally, the code also calculates the general sizes of the wing and tail surfaces. These calculations are based primarily on the aspect ratio, the wing loading, and the tail volume coefficients. The tail volume coefficients are provided in Daniel Raymer’s textbook and are based primarily on historical data. It was assumed that the canard can be sized using the horizontal tail volume coefficients from Raymer’s textbook.4 The current sizes for the wing, horizontal tail, vertical tail, and canard surfaces are listed below as REF _Ref257359965 \h Table 7. Table SEQ Table \* ARABIC 7: Wing SizesWingVertical TailHorizontal tailCanardCanardAR10.52.4-9b96.128.1-40.7c10.510.5-4.5S876.5329.5-183.6ConventionalAR8.01.02.0-b87.213.723.8-c12.012.38.9-S950.7186.6284.2-Drag Component BuildupThe drag for the aircraft may initially be estimated by considering component contributions. Essentially, the aircraft may be broken down into a handful of major sections with each section’s drag estimated through a combination of theoretical and empirical equations. Each component’s drag may also be modeled as a sum of different types of drag. There are four major types of drag being considered in our analysis. The first is parasite drag, Dp. Parasite drag is comprised of skin friction drag, pressure drag and interference drag. This drag is a result of surface interactions with a viscous fluid, in this case, air. The coefficient of parasite drag is defined as: CDp = i=1# compKiQiCfiSwetiSref+kCL2However, the second term will be used in a different manner described later. Sref is the reference area, in this case the wing planform area. K is the form factor, a function of component shape that takes into account pressure drag. Q is interference factor. Depending on the placement of the component, Q may range from 1 to 1.5. In this case of this aircraft, Q will be usually be 1. Swet is the wetted surface of the component, described as being the surface area of the object which would be wetted by dropping the aircraft into a bucket of water.The second major type of drag is induced drag, Di. Induced drag is a product of finite-span airfoils, whereby wingtip vortices flow from the lower to upper surface of the wing, lowering the airfoil’s effective angle of attack. The presence of wingtip vortices leads to a component of the lift acting as drag since drag is defined as the force produced parallel to the free stream. The induced drag coefficient is defined according to the equation below.CDi= CL2πAR?Here, ? is the span efficiency factor, defined to be 1 when the wing produces an elliptic lift distribution. Wing shape is very important in determining the lift distribution and many aircraft employ near-elliptic lift distributions. Current general aviation aircraft generally have an efficiency factor ranging between 0.75 and 0.85. The aircraft’s induced drag coefficient is currently being estimated using an efficiency factor of 0.80. The third type of drag is wave drag, Dw, also known as compressibility drag. Wave drag results from the formation of shock waves produced around the aircraft at sonic or near-sonic speeds. The energy of the shock waves radiating off the aircraft is perceived as drag. While the aircraft will not ever fly at Mach 1, localized airflow will most certainly reach sonic speed. As of yet, no method for predicting or estimating wave drag has been found or included for this application, thus it is currently assumed that the aircraft experiences 20 counts of wave drag. While likely incorrect, the inclusion of a constant value of wave drag serves as a place holder for future development and as a rough estimate for the early design stage analysis. This omission will be corrected in the future.The last type of drag being considered is miscellaneous drag, Dmisc, which is a factor taking into account leakage, protuberance, and imperfections in manufacturing. This drag accounts for issues such as gaps in control surfaces, external lighting, antennae, rivets and other surface blemishes. Based off of sample drag prediction information, conservative estimates put the aircraft’s miscellaneous drag coefficient at 5% of the aircraft’s parasite drag coefficient. This approximation is used exactly in the current drag prediction. To vary slightly from the described list of drag coefficients, it is convenient to redefine two aspects of the total drag into a new term called ‘drag due to lift’. This term is comprised of the parasite drag term dependent on CL, kCL2, and the induced drag term, CL2πAR?. Together these terms are combined to produce the term CL2πARe, the drag due to lift, where e is Oswald’s efficiency factor. As stated previously, e is estimated to be 0.8. For the purpose of this report it is reasonable to determine CL at cruise based off of the assumption that lift equals weight during cruise. Accordingly, Wcruise = CLqS. Sample calculations are provided in APPENDIX C. Now that the types of drag have been discussed, it is time to consider the contributions from the aircraft components themselves. The major components being considered on the aircraft are the fuselage, main wing, vertical and horizontal tails, canards, nacelles and pylons. The fuselage is currently being modeled as a cylinder with a nose and tail section added on. The fineness ratio of the aircraft, λf, was determined via a combination of current similar aircraft’s fineness ratios and mission design criteria such as ceiling height and volume per passenger. The wetted area of the fuselage is a function of the fineness ratio, and can be calculated according to the following equation.Swetf=πDflf1-2λf23(1+1λf2)Here, Df and lf are the diameter and length of the fuselage, respectively. The wetted area is necessary to calculate the parasite drag, as skin friction only affects the wetted area of the aircraft. The wetted area for the nacelles uses a simple model of a cylinder. For this the wetted area is computed with the following equation.SwetN=πDNlNSimilarly to the fuselage D is the diameter and l is the length. Since the wings, horizontal tail, vertical tail, and pylons all all similar shapes, it is possible to use the same equation. For these surfaces, the wetted area was calculated with the following equation.Swet=Sw2cSw is the plan form area, c is the curvature of the surface, and multiplied by 2 to take the top and bottom surface into account.For the component drag build up, the form factor for the shape of each component was needed. The form factor is shape dependent and varied for each component according to the equations listed below.Kf=1+60λf3+λf400KN=1+0.35(lNDN)Kw=1+Ztc+100(tc)4In these equations, the subscript f is for the fuselage, N is for the nacelle, and W is for the wings. The equation for the wings was also used for the pylons, vertical tail, and horizontal tail. Here, λf is again the fineness ratio, Z is the sweep correction factor, and t/c is the thickness to cord ratio. To find this form factor, the sweep correction factor is needed. This factor was found using the following equation.Z=(2-M2)cosΛc/41-M2(cosΛc/4)2With the form factor and wetted area known, the last required piece is the interference factor. The interference factor is determined by where different components are located relative to each other. For this aircraft the nacelles have an interference factor with the fuselage. This factor is based on the distance from the fuselage to the nacelle. If the distance is greater than the diameter of the nacelles, the interference factor, Q, equals one. If the nacelle is resting directly on the fuselage then the interference factor, Q, is 1.5. The wing and the fuselage also interfere with each other. The position of the wing, being a low wing with fillets, results in an interference factor that is also 1. For the pylons and tail, it is similar to a midwing set up, which will result in an interference factor also of 1. With these four factors all known, it is possible to use the following equation to sum up the different components and their drags to form a total parasite drag.CDp = i=1# compKiQiCfiSwetiSrefThe total drag coefficient, CD is very simply defined to be a sum of the four types of drag discussed previously. CD= CDp+ CDi+ CDmisc+ CDwCD= i=1# compKiQiCfiSwetiSref + CL2πARe + CDmisc + CDwENGINE SELECTIONTwo separate and distinct engine concepts are being explored by the team: an unducted propfan and a geared turbofan. Both concepts offer significant savings in both emissions and fuel consumption, and incorporate the most modern designs in the aircraft propulsions industry. The details of each particular engine model were discussed previously in the Advanced Technologies section of this report, but a table of the estimated performance improvements is provided below as REF _Ref257707989 \h Table 8.Table SEQ Table \* ARABIC 8: Engine Model Projection and ComparisonPredicted Engine Performance by 2020?% Fuel ReductionNoise Reduction (dB)% Emission ReductionUnducted fan25-30Meets Stage 420Geared TurboFan152050An in depth analysis process is currently being used to model the engine as part of the sizing code. The main purpose of the engine model is to predict how much fuel will be required to complete the design mission, and more specifically the fuel weight. This process starts with the engine data provided in Raymer’s textbook for a “High Bypass Turbofan.”4 Unfortunately, Raymer is not more specific than that as to what model the engine actually is. The engine data plots are scanned directly into a MATLAB script which converts the data to tabular form and interpolates where necessary. This engine is much larger than required, and must be scaled down from 50,000 lbs of thrust to approximately 15,000 lbs of thrust in order to be useful. The scaling process is performed simply by calculating a scale factor, which equals the desired amount of thrust at any given time divided by 50,000 lbs, and multiplying the thrust data by this value. As engines shrink in size and performance, however, their fuel consumption and efficiency generally worsen. Because of this, the specific fuel consumption was scaled inversely proportional to the established scale factor.With the data provided and the scaling operations performed, the fuel consumption can then be predicted as a function of Mach number, altitude, and throttle setting. Different calculations are used for each segment of the flight mission: taxi, takeoff, climb, cruise, descent, loiter, and land. The appropriate equations provided by Professor Crossley in lecture were used exactly for each mission segment. Some sections are simple and straight forward, such as taxi, takeoff, and landing, while the others require some looping to find optimum climb rates or lift to drag ratios. The climb calculations predict the fuel consumed at a velocity of 250 knots up to 10,000 ft followed by the fuel consumed at best rate of climb up to the initial cruise altitude of 41,000 ft. The calculations are broken down into 500 vertical foot increments in which it is assumed that atmospheric properties are constant. The cruise section is broken into twenty equal-portion increments, and it is assumed that cruise altitude will remain constant at 41,000 ft. The aircraft is now also assumed to be flying at its maximum lift to drag ratio. This same performance assumption is used for loiter calculations. Descent is given a zero-range credit, which has not changed since the System Requirements Review.Each engine is analyzed using the above process with the technology factors from REF _Ref257707989 \h Table 8 included. When research states that a particular engine will decrease fuel consumption by X%, that factor is used directly when applied to the specific fuel consumption value in calculations. For example, the unducted fan reportedly reduces fuel consumption by 30%, so all SFC values taken from the engine model in Raymer’s textbook are scaled and then multiplied by 0.70. While noise emission is not currently predicted, this same method will be used when the calculations become possible.SUMMARY OF AIRCRAFT CONCEPTSThe major goals of this design are to provide an environmentally conscious, ultra-long range corporate aircraft, which meets NASA’s N+2 goals, is fast enough to be marketable, and can operate out of general aviation airports. Over the course of the design process, these broad goals have given way to a detailed set of requirements. These requirements are shown in the left column of the requirements compliance matrix in Figure 9. In addition to specifying the name of the criteria being measured, this matrix also specifies the target value of each criterion, the minimum threshold value necessary to satisfy each requirement, and the current value based on the present state of the design. It is important to note that several of the current values shown are estimates, and may not be constrained by the current level of system detail. These estimated values are denoted by a trailing asterisk in the matrix.The range, cruise mach, and initial cruise altitude requirements were derived from the design mission. While the locations of the design mission have not changed, the still-air range of the mission has been adjusted since the System Requirements Review (SRR). Originally, the aircraft’s design mission range was specified as the great circle distance between Los Angeles (LAX) and Hong Kong (HKG). This range ignored the fact that strong headwinds would likely be unavoidable when traveling west to Hong Kong over the Pacific. As a conservative figure, a headwind of 60 knots was assumed for the duration of the flight. This resulted in an equivalent still-air range of 7100 nautical miles. A threshold value of 6960 nautical miles was then determined from a less conservative, but still realistic, headwind of 50 knots. The cruise mach and initial cruise altitude remain on target since the SRR, and are incorporated into the aircraft’s current sizing and weight estimates. The target and threshold values of these requirements are both based on producing an aircraft that is fast enough to be marketable. The cruise mach is representative of most modern jet aircraft; and the initial cruise altitude is high enough to allow for flight above most of the cruise altitudes used by part 121 carriers.The maximum passenger capacity and the volume per passenger per hour requirements were obtained from the aircraft’s cabin layout and design mission. The maximum number of passengers has remained on target since the SRR, and is factored into the cabin’s seating configuration. The threshold value of 8 listed in Figure 9 signifies the number of passengers carried on the aircraft’s design mission. The volume per passenger per hour requirement was a new system requirement added to provide a measure of passenger comfort. The current threshold value of 2.28 ft3/(pax·hr) represents the minimum level necessary to ensure the design mission is not arduous to passengers.5 The current and target value of volume per passenger hour is 13.3 ft3/(pax·hr). This value was calculated from a CATIA model of the aircraft’s cabin, and is considered to be comfortable over the course of the design mission, which is approximately 13 hours.5 The balanced takeoff field length at maximum takeoff weight (MTOW) is also a new requirement, which replaced the previous and more ambiguous requirement of “takeoff distance” featured in the SRR. Balanced takeoff field length is the minimum length of runway required such that a twin engine aircraft can operate safely with one engine inoperative during any stage of takeoff. It was chosen as the metric for the design’s takeoff performance, since it specifies the minimum amount of runway necessary to comply with safety regulations. The target value of 6000 ft was selected since it is close to the length of runways found at many general aviation and regional airports. The threshold of 7000 ft was chosen since it was approximated as the maximum required field length at which the aircraft could still be marketed as having reasonable airport access. Since the current level of design detail was not sufficient to calculate balanced field length, the ground roll values for each concept were included in the matrix. These ground roll values were calculated using the constraint diagrams for each concept. While the current ground roll values are promising in comparison to existing aircraft, this requirement will continue to be of great concern until field length can be approximated.The cabin noise, cumulative noise, emissions, sill height, and variable cost requirements remain unchanged since the SRR, and are currently only estimates unconstrained by the present level of design detail. Reasonable cabin noise values were estimated from existing aircraft data. However, it is not yet possible to calculate cabin noise, or the difference in cabin noise that is expected to exist between the two concepts. Cumulative noise limits were also estimated from existing aircraft data, while emissions levels were placed relative to the NASA N+2 goals. Cumulative noise limits cannot yet be calculated, but emissions levels will be able to be refined once more detail on the power plants is obtained. Sill height was estimated based on the need to facilitate aircraft loading, and also by studying the sill heights of existing business jets. This is expected to be governed by takeoff rotation considerations in the future. Variable costs were estimated from existing aircraft cost data, and the cost of operating a current aircraft with a similar design mission was used as the threshold. Clearly, the variable cost cannot be calculated at this point in the design. Lastly, the cruise specific range requirement represents the desired fuel efficiency of the design. The threshold value of 0.26 nm/lb was based on existing aircraft performance.6 Current values were calculated for each concept, using the fuel weight determined from the sizing code and the range of the design mission. This method was a crude approximation, but a more accurate answer will be achievable once more power plant details are determined. The target value was set to the range necessary to fly the design mission with reserves, assuming current fuel weight. Figure SEQ Figure \* ARABIC 9: Requirements Compliance Matrix.Scaled models of each concept have been developed in CATIA. Three view drawings with a few general sizing dimensions and a three dimensional screen shot from CATIA have been provided for the Canard Pusher Configuration as REF _Ref257706365 \h Figure 10 and REF _Ref257706384 \h Figure 11. Similar plots follow for the Conventional Configuration in REF _Ref257706397 \h Figure 12 and REF _Ref257706405 \h Figure 13.Figure SEQ Figure \* ARABIC 10: Three View Drawing - Canard Pusher ConfigurationFigure SEQ Figure \* ARABIC 11: Isometric CATIA Screen Shot - Canard Pusher ConfigurationFigure SEQ Figure \* ARABIC 12: Three View Drawing - Conventional ConfigurationFigure SEQ Figure \* ARABIC 13: Isometric CATIA Screen Shot - Conventional ConfigurationSUMMARY The second phase of this project consisted of developing many designs using various technologies and configurations and then to narrow those designs down to the two best choices for further analysis. The two choices that best met the developed requirements are a conventional aircraft and an aircraft with canards and an aft swept back wing. Using these two configurations, advanced technologies will be implemented such as composites and more efficient engines like the unducted turbofan or the geared turbofan. Current calculations show that the aircraft, sized to meet the demands of the target customer and designed to achieve the environmental goals, are in the low 70,000 lb range. Further analysis will be done to evaluate these two designs to determine which is best and which meets those requirements best.NEXT STEPSNow that Team 2 has finished the System Design, there are many more steps that need to be taken before the conceptual design of the aircraft can be completed. For the wing, tail, and canard, airfoils need to be selected to better predict the drag polar and general performance of each piece. In propulsion, the engine selection between the geared turbofan and unducted fan needs to be decided. In structures, a lot of work needs to be done especially in loading of different parts and in integrating the landing gear. The longitudinal stability of each aircraft needs to be finalized and the lateral stability needs to be analyzed for cross-wind landing and one engine out conditions. Using all of this updated data, the best performance data needs to be calculated including all airplane weights and sizes. The internal and external noise of each aircraft needs to be assessed. Finally the estimated cost of each aircraft need to be found. Using all of this and our component sizing code, carpet plots need to be used to graphically optimize each of our designs. Once all of the calculations and optimizations have been completed a final aircraft selection will be made. REFERENCES1“Subsonic Fixed Wing Project”. NASA. 08 February 2010. Norris, Guy. “Rotor Revival”. Aviation Week & Space Technology. 14 December 2009. pages 54-57. 3 "PurePower PW1000G." Home. Web. 30 Mar. 2010. < Raymer, Daniel P. Aircraft Design A Conceptual Approach (Aiaa Education Series). New York: AIAA American Institute of Aeronautics & Ast, 2006. Print.5Torenbeek, 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.6Jane's All The World's Aircraft. Web. 11 Feb. 2010. < AConventional Configuration Input Variables%% Input section % aircraft contant ratiosWo_S = 76; %lbs/ft^2 from initial contrait diagram Cfactor = GetCfactors('G650'); %Tech Factors engine.tech_fact = .7; %Technology factor - SFC (~30%) (~40%=projected) % wing variableswing.AR = 8;wing.taper = .3;wing.sweep = deg2rad(35); %radwing.t_c = .08;wing.Scs = 25; %control surface area on wing (ft^2) %vertical tail variablestail_vt.Cvt = .09;tail_vt.Lvt = 40; %feettail_vt.taper = .8; tail_vt.sweep = deg2rad(25); %radtail_vt.t_c = .08;tail_vt.Ttail = 1; %1 for t-tail, 0 otherwisetail_vt.AR = 1;tail_vt.Scs = 40; %control surface area on vertical tail (ft^2) % canard variablescanard.Lc = 40; %feetcanard.taper = .5; = 1;canard.AR = 2;canard.sweep = deg2rad(40); %radcanard.t_c = .08;canard.Kuht = 1; %all moving = 1.143, otherwise 1canard.Fw = 9; %fuselage width at canard location (ft)canard.Scs = 50; %control surface area on canard (ft^2) %pylon/Nacellesnacelle.L = 20;nacelle.D = 5;pylon.c = 1;pylon.t_c = .3;pylon.S = 10; %fuselagefuselage.L = 90; %length of fuselage (ft)fuselage.D = 9; %diameter of fuselage (ft)fuselage.SDepth = 7; %fuselage structural depth (diam at wing box) (ft)fuselage.Lc = 40; %wing mac to canard mac (ft) %landing geargear.location = 1; %1 = wing, 1.12 for fuselage mountedgear.Nl = 2.5; %ultimate landing load factor (1.5*Ngear)gear.length_main= 5; %ftgear.length_nose = 5; %ftgear.main_wheels = 4; %number of main gear wheelsgear.main_struts = 2; %number of main gear shock strutsgear.nose_wheels = 2; %number of nose gear wheels % mission variablesmission.num_pax = 16;mission.weight_pax = 220;mission.num_crew = 4;mission.weight_crew = 200;mission.range = 7000; %nmimission.q = 200; %dynamic pressure at cruisemission.Nz = 2.5*1.5; %ultimate load factor (1.5*load factor)mission.Mcruise = .85;mission.cruise_alt = 42000; %engine variablesengine.Kng = 1.017; %1.017 for pylon mounted nacelle, 1.0 otherwiseengine.nacelle_length = 20; %ftengine.nacelle_width = 5; %ftengine.weight = 3000; %lbsengine.number = 2; %total number of enginesengine.d2c = 110; %ft - total distance from engines to cockpit (sum for both engines)engine.Kr = 1; %1.133 if reciprocating engine, 1.0 otherwiseengine.Ktp = 1.0; %0.793 if turboprop, 1.0 otherwiseengine.SFC_cruise = .5; %1/hrengine.SFC_loiter = .4; %1/hr %controls variablescontrols.num_fnx = 5; %number of functions performed by controlls (usually 4-7)controls.num_mech = 1; %number of mechanical functions (usually 0-2)controls.Iy = 1001000; %yawing moment of inertia (lb ft^2) % aerodynamicaero.V_stall = 210; %ft/s - stall for the main wing %weightweight.Wo_guess = 100000;weight.APU_u = 125; %lbs - uninstalled weightweight.avionics_u = 250; %lbs - uninstalled avionics weight %miscallaneousmisc.Rkva = 50; %system electrical rating (kV) - usually 40 to 60 for transportsmisc.La = 250; %electrical routing distance, generators to avionics to cockpit (ft)misc.Ngen = 2; %number of generators (usually equal to number of engines)misc.Vp = 4000; %volume of pressurized section (ft^2) %Locations (note, all in feet relative to nose, plane length ~ 90 ft)loc.apu = 82;loc.avionics = 5;loc.fuel = 42.5;loc.wing = 42.5;loc.canard = 90;loc.tail_vt = 86;loc.fuselage = 45;loc.main_gear = 50;loc.nose_gear = 10;loc.nacelle = 72;loc.engine_controls = 5;loc.engine_starter = 5;loc.fuel_system = 42.5;loc.flight_controls = 5;loc.instruments = 5;loc.hydraulics = 50;loc.electrical = 45;loc.furnishings = 35;loc.airconditioning = 6;loc.antiice = 35;loc.crew = 10;loc.payload = 35;loc.engines = 72;loc.AC = 50;APPENDIX BCanard Pusher Configuration Input Variables%% Input section % aircraft contant ratiosWo_S = 84; %lbs/ft^2 from initial contrait diagram Cfactor = GetCfactors('Starship'); %Tech Factors engine.tech_fact = .7; %Technology factor - SFC (~30%) (~40%=projected) % wing variableswing.AR = 10.53;wing.taper = .2;wing.sweep = deg2rad(30); %radwing.t_c = .1;wing.Scs = 10; %control surface area on wing (ft^2) %vertical tail variablestail_vt.Cvt = .09;tail_vt.Lvt = 23; %feet wing mac to tail mactail_vt.taper = .8; tail_vt.sweep = deg2rad(30); %radtail_vt.t_c = .08;tail_vt.Ttail = 0; %1 for t-tail, 0 otherwisetail_vt.AR = 2.4;tail_vt.Scs = 20; %control surface area on vertical tail (ft^2) % canard variablescanard.Lc = 50; %feetcanard.taper = 1; = 1; %volume coefficient canard.AR = 9;canard.sweep = deg2rad(35); %radcanard.t_c = .08;canard.Kuht = 1.143; %all moving = 1.143, otherwise 1canard.Fw = 9; %fuselage width at canard location (ft)canard.Scs = 35; %control surface area on canard (ft^2) %pylon/Nacellesnacelle.L = 15;nacelle.D = 5;pylon.c = 1;pylon.t_c = .3;pylon.S = 10; %fuselagefuselage.L = 90; %length of fuselage (ft)fuselage.D = 9; %diameter of fuselage (ft)fuselage.SDepth = 7; %fuselage structural depth (diam at wing box) (ft)fuselage.Lc = 50; %wing mac to canard mac (ft) %landing geargear.location = 1; %1 = wing, 1.12 for fuselage mountedgear.Nl = 2.5; %ultimate landing load factor (1.5*Ngear)gear.length_main= 4; %ftgear.length_nose = 4; %ftgear.main_wheels = 4; %number of main gear wheelsgear.main_struts = 2; %number of main gear shock strutsgear.nose_wheels = 1; %number of nose gear wheels % mission variablesmission.num_pax = 18;mission.weight_pax = 220;mission.num_crew = 4;mission.weight_crew = 200;mission.range = 7000; %nmimission.q = 200; %dynamic pressure at cruisemission.Nz = 1.5*2.5; %ultimate load factor (1.5*load factor)mission.Mcruise = .85;mission.cruise_alt = 42000; %engine variablesengine.Kng = 1; %1.017 for pylon mounted nacelle, 1.0 otherwiseengine.nacelle_length = 20; %ftengine.nacelle_width = 5; %ftengine.weight = 3000; %lbsengine.number = 2; %total number of enginesengine.d2c = 100; %ft - total distance from engines to cockpit (sum for both engines)engine.Kr = 1; %1.133 if reciprocating engine, 1.0 otherwiseengine.Ktp = 1; %0.793 if turboprop, 1.0 otherwiseengine.SFC_cruise = .5; %1/hrengine.SFC_loiter = .4; %1/hr %controls variablescontrols.num_fnx = 5; %number of functions performed by controlls (usually 4-7)controls.num_mech = 1; %number of mechanical functions (usually 0-2)controls.Iy = 1000100; %yawing moment of inertia (lb ft^2) % aerodynamicaero.V_stall = 220; %ft/s - stall for the main wing %weightweight.Wo_guess = 75000;weight.APU_u = 200; %lbs - uninstalled weightweight.avionics_u = 300; %lbs - uninstalled avionics weight %miscallaneousmisc.Rkva = 50; %system electrical rating (kV) - usually 40 to 60 for transportsmisc.La = 150; %electrical routing distance, generators to avionics to cockpit (ft)misc.Ngen = 2; %number of generators (usually equal to number of engines)misc.Vp = 737; %volume of pressurized section (ft^2) %Locations (note, all in feet relative to nose, plane length ~ 90 ft)loc.apu = 10;loc.avionics = 5;loc.fuel = 78;loc.wing = 78;loc.canard = 10;loc.tail_vt = 83;loc.fuselage = 45;loc.main_gear = 80;loc.nose_gear = 10;loc.nacelle = 85;loc.engine_controls = 5;loc.engine_starter = 5;loc.fuel_system = 78;loc.flight_controls = 5;loc.instruments = 5;loc.hydraulics = 40;loc.electrical = 45;loc.furnishings = 35;loc.airconditioning = 6;loc.antiice = 70;loc.crew = 10;loc.payload = 35;loc.engines = 85;loc.AC = 70;APPENDIX CDrag Sample CalculationsAt a cruise height of 42,000 feet, at a cruise speed of 0.85 Mach, q is found to be q = 0.5*(0.000532 slugs/ft3)*(0.85*968.076 ft/sec)2q = 180.11 psfAs well, S for the aircraft is currently estimated at 1035 ft2. Thus, CLcruise = WqS ≈ 90,000 lb180.11 psf*1035 ft2 = 0.483The drag due to lift will vary slightly based on whether the canard or standard aircraft setup is considered, since the canard aircraft boasts an AR =10 while the standard aircraft currently has an AR=8. Thus, for the standard aircraft,CD = 0.01924 + 0.01159 + 0.0010+ 0.0020= 0.03384For the canard,CD = 0.01924 + 0.009274 + 0.0010 + 0.0020= 0.03151Once CD is found the total drag force, DTotal, may be found very easily by the equationDTotal = CD q SThus, the current total drag for the standard aircraft is found to beDTotal = 0.03384*(180.11 psf)*1035 ft2DTotal = 6308 lbf ≈ 6310 lbfThe current total drag for the canard aircraft isDTotal = 0.03151*(180.11 psf)*1035 ft2DTotal = 5837 lbf ≈ 5840 lbfNote that it is reasonable to approximate the drag to the ten’s place based on the amount of approximation present in the calculations. ................
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