Gijung BaeJoe BlakeJung Hoon ChoiJack GeererAlison Jean ...



-1156290-880952Gijung BaeJoe BlakeJung Hoon ChoiJack GeererAlison Jean GongSang Jin KimMike McCarthyNick OschmanBryce PetersonLawrence RauxHwan SongMay 2010Concept Design ReviewI Contents TOC \o "1-3" \h \z \u II Executive summary PAGEREF _Toc261038726 \h 4- III Mission Statement PAGEREF _Toc261038727 \h 6- IV Design requirements PAGEREF _Toc261038728 \h 6- V. Selected “best” aircraft concept PAGEREF _Toc261038729 \h 7Walk-around PAGEREF _Toc261038730 \h 7Values of major design parameters PAGEREF _Toc261038731 \h 9- VI. Results of aircraft sizing and carpet plots PAGEREF _Toc261038732 \h 10Brief description of sizing code PAGEREF _Toc261038733 \h 10Validation & Fudge Factor PAGEREF _Toc261038734 \h 14Fixed design parameter values PAGEREF _Toc261038735 \h 16Drag Prediction PAGEREF _Toc261038736 \h 16Results of engine modeling PAGEREF _Toc261038737 \h 17Mission modeling PAGEREF _Toc261038738 \h 20- VII. Major design trade-offs PAGEREF _Toc261038739 \h 21Trade off overview PAGEREF _Toc261038740 \h 21Summary of carpet plot studies PAGEREF _Toc261038741 \h 23 - VIII. Aircraft description PAGEREF _Toc261038742 \h 24Dimensioned three-view, to scale PAGEREF _Toc261038743 \h 24Representative internal layout PAGEREF _Toc261038744 \h 26 - IX. Aerodynamic design details / justification PAGEREF _Toc261038745 \h 28Drag Polar PAGEREF _Toc261038746 \h 29- X. Performance PAGEREF _Toc261038747 \h 30- XI. Propulsion PAGEREF _Toc261038748 \h 33- XII Structures PAGEREF _Toc261038749 \h 34Configuration layout and highlights PAGEREF _Toc261038750 \h 36Important Load Paths PAGEREF _Toc261038751 \h 36Wing-fuselage intersection PAGEREF _Toc261038752 \h 38Engine mounts and landing gear integration PAGEREF _Toc261038753 \h 39Materials PAGEREF _Toc261038754 \h 41- XIII. Weights and balance PAGEREF _Toc261038755 \h 42Overview discussion on prediction methods PAGEREF _Toc261038756 \h 43Center of gravity PAGEREF _Toc261038757 \h 45- XIV. Stability and control PAGEREF _Toc261038758 \h 47Preparations for malfunctions and extreme situations PAGEREF _Toc261038759 \h 49Engine failure PAGEREF _Toc261038760 \h 49Extreme crosswind PAGEREF _Toc261038761 \h 50- XV. Noise PAGEREF _Toc261038762 \h 50- XVI. Cost PAGEREF _Toc261038763 \h 51- XVII. Summary PAGEREF _Toc261038764 \h 54APPENDIX A: Bibliography PAGEREF _Toc261038765 \h 55APPENDIX B: Sizing code PAGEREF _Toc261038766 \h 56APENDIX C: Compliance Matrix PAGEREF _Toc261038767 \h 57 II Executive summaryTeam XG International, Inc. sought to develop an environmentally-sensitive aircraft which will provide our customers with a 21st-century transportation system that combines speed, comfort, and convenience while meeting NASA’s N+2 criteria.reduce 40% fuel consumptionreduce noise by 42 dBreduce NOx emissionreduce takeoff field length The XG Endeavour’s range of 3,700 nautical miles enables it to make almost any continental flight, as well as many intercontinental flights, and the Mach 0.80 cruising speed makes the trip in a hurry. The convertible cabin layout option also allows the customer to add their own personal touch. While XG International sought to create a fast and convenient air transportation system, they also required an environmentally sensitive aircraft. Therefore, the XG Endeavour offers considerable reductions in noise and exhaust emissions, fuel consumption, and runway field length.The power of three GE Honda Aero HF120 variant turbofan engines, capable of outputting a combined 10,800 pounds of thrust, lifts the XG Endeavour into the sky. At the cruising altitude of 45,000 feet, fuel consumption is greatly reduced when one internal turbofan is turned off, while the two remaining engines, along with the sleek, swept back wings, and cruciform tail allow the XG Endeavour to soar to its destination. During daylight hours, solar film on the upper surfaces of the fuselage and wings collect energy and provide power for in-flight systems.XG International engineers designed an all-new sizing code for the Endeavour project. The jet’s body, geometry, weight, airfoil, design mission, and many other factors were incorporated in the development of this new code. In order to compete in the business jet market, the advantages and disadvantages of every design aspect needed to be weighed against one another. While the three turbofan engines added weight and considerable maintenance costs, having all three engines would be safer in an engine-out situation, and it would also be fuel efficient when turning one off during cruise. The cruciform tail allows for an aft fuselage engine. Solar film without adding significant weight is possible power conservator. The size of the fuselage was also taken into consideration. Inside the cabin, there is ample space to comfortably seat nine people and the two required crew members, though the final layout will ultimately depend upon the customer’s specifications.Formers, ribs, stringers and longerons made of a titanium aluminum alloy and composite materials make the XG Endeavour reinforced structure to withstand the forces experienced during flight. The two horizontal stabilizer-mounted engines have been designed to meet CFR36 Stage 4 noise requirements. Also, more than ample takeoff thrust allows for a decreased time to climb which reduces ground signature during the flyover stage of takeoff. After considering these and many more features, the team has decided upon the present design for the XG Endeavour.Through careful cost analysis, and using the Rand DAPCA IV cost model, XG International estimates that the Research Test Development & Evaluation and Flyaway cost be about $2.4 billion. With a profit per aircraft around $750,000, the breakeven point is at 20 aircraft, which makes this a very practical business plan for the future. These estimations include considerations for depreciation, insurance, crew, fuel and maintenance. The XG Endeavour has been expertly designed to lead the way into the future of noise, exhaust, gas, and field length reduction while providing a fast and convenient mode of air transportation. - Team XG - III Mission StatementThe XG team strives to develop ambitious advancements in green performance and technology by fulfilling the promise outlined in our mission statement:“Develop an environmentally-sensitive aircraft which will provide our customers with a 21st century transportation system that combines speed, comfort, and convenience while exceeding NASA’s N+2 criteria.”ivAn aircraft can damage the environment in many ways; from the consumption of finite fossil fuels, to harmful emissions, and noise; all of which must be considered in a complete green project. For the aircraft must be environmentally-sensitive, the manufacturing details such as material choice and engine details; and the operating lifespan of the final product must be considered equally. Though it was expected that the Endeavor would be designed with several elite performance capabilities in mind. Tight constraints forced several parameters to be changed from the previous document. The aircraft will still strive to incorporate a shorter takeoff length than similar sized competitors’ aircraft while still making an effort to minimize the sacrifice in speed and range. - IV Design requirements 3667125167005The design requirements of XG Endeavour were mainly guided by the NASA’s N+2 goals. In order to be one of the main competitors of the business jet field, the aircraft had to have the green image, and NASA’s requirements state them very well. A table of NASA’s N+2 goals for a twin aisle aircraft are copied and displayed to the right. Figure IV.1: NASA’s N+2 Design RequirementsThe most important design requirement that NASA proposed is to reduce the fuel consumption of the aircraft by 40%. It was unlikely that any one component of any aircraft could make a 40% impact by itself, so major analysis was conducted to improve efficiency in multiple areas such as reducing areas responsible for drag, or through weight reduction. Cumulatively, it is possible to achieve the future standards. This bodes especially true if engine technology will follow an improvement trend. NASA goals also state that the noise emitted during flight has to be reduced by a total of 42 dB. In order to meet this requirement, various changes in the aircraft configurations must be considered. The main consideration of this was the placement of the engines. Depending on engine location, noise can be reduced during certain conditions of flight. For example, when the engines are placed on top of the wing, the sound wave bounce off the wings and and back into to the sky, which in turn reduces the noise detected by the sensors located on the ground, which means that the noise emission during takeoff will be dramatically reduced.The third requirement was to reduce the NOx emission by 75%. This requirement can be met by reducing the fuel consumption, and possibly by using a new technology that may be developed by the year 2020 that will emit the much less NOx. One of the final NASA requirements set by the N+2 goals stated above was to shorten takeoff field length by 50%. But because the afore stated NASA N+2 goal are intended to be applied to large twin isle aircrafts that are limited by their longer take off need, this particular goal was less practical for a small aircrafts that can already operate in relatively small airport. Because XG Endeavour is not a large twin aisle aircraft, the airport field length is short enough to consider this requirement as lower priority in relationship to the others. it is clear that all these requirements have at least and indirect relationship to one another. For example, if the takeoff field length is reduced, the aircraft reduces the rate of fuel consumption as well as resulting in a reduction of NOx emission. In addition, an improvement in take off distance enables the aircraft to rise off of the ground in a shorter period of time, reducing the time of which sampled noise will also be detected. All things considered, The team has concluded that the foremost priority is to reduce the fuel consumption, because this will enable the aircraft to emit much less NOx; reduce the takeoff gross weight; and eventually reduce the takeoff distance. - V. Selected “best” aircraft conceptWalk-aroundThe three engine configuration is designed to provide more fuel efficient flight at cruise condition, the point at which the most of the fuel is used. The total fuel consumption of the flight will be reduced by using engines optimized for cruise condition.The closeable duct, which feeds air to the center engine, will decrease the drag when the aft fuselage mounted engine is not in use. The engine inlet duct will be deployed out of fuselage when the engine is in use, and when the engine is not being used, the duct can fold into the fuselage. Because it does not coincide with the pressurized section of the fuselage, empty weight added to incorporate this feature will not be significant; the team thinks that the benefit from the reduction of drag during the long cruise will be greater than the increased empty weight to close the ducts.The 2 engines that are on the horizontal stabilizers will produce 2000 lbs of thrust each, producing maximum thrust of 4000 lbs during cruise. These engines will operate at the most efficient throttle setting during cruise to maximize fuel efficiency.The aft fuselage engine will produce 6800lbs of thrust. This is needed for takeoff and climb sections of the flight. This engine is designed to be turned off during cruise section of the flight mission. The sizing code calculated the static margin of the airplane. It depends heavily on the location of the main wing because fuel and landing gear will be housed in the wings. The calculated static margin was 16% for the final design. The CG travel graph was also generated using the sizing code.Below (Figure V.1) are two rendered views of the XG Endeavour. The image indicates examples of the concepts mentioned above. 3rd EngineFigure V.1: (top) rendered top view of the airplane in flight (bottom) bottom view displaying third engine.??Values of major design parameters The aircraft needed some of the input variables to start design. Table VI.1: Design parameter constantsDesign ParameterValuet/c0.1428.13°0.5CLmax1.6AR7.8Wo /S92.6TSL/Wo0.33The design parameters for the airplane were set according to historical values or calculated in the sizing code. Wing loading and thrust-to-weight ratio at the sea level were found in the current constraint diagrams for the final concept. Aspect ratio for the concept was chosen based on the size of similar sized business jets. The sweep angle was chosen so that the center of lift is located behind the center of gravity. Thickness to chord ratio was calculated based on the NACA model. In this instance, CLmax is assumed based on the historical data and the shape of the XG Endeavour. - VI. Results of aircraft sizing and carpet plots Brief description of sizing code?????The sizing code was constructed using the Microsoft Excel spreadsheet. Six separate sheets were combined to calculate the empty weight, fuel weight, static margin, drag, flight time, and the shape of the concept aircraft. The ‘Main page’ is where all the major input and output values are displayed. Below is a screen shot of our main page for the final design. Figure VI.1: Sample image of the ‘Main’ worksheet of the Sizing CodeThe length, diameter and locations of fuselage, nose, and empennage were input first, then the leading edge angle, taper ratio, aspect ratio of the various lifting surfaces were input. The area of the main wing was calculated by team the weight of the aircraft, so it was not an input value in the main page. The locations of the engines were input. Because there were two separate types of engines, the calculations for thrust were divided for analysis. The thrust of engine used in cruise was first calculated then the aft fuselage mounted engine was used to carry rest of the thrust the aircraft needs to takeoff. With above information, the sizing code displayed a rough sketch of the aircraft in a graph format. (See figure above) This allowed us to quickly notice any major mistakes with location of the major components. Geometry worksheet is where the aircraft’s general shape is calculated from the input values that were input in the main worksheet. Three-view of the airplane is displayed to easily view if any structural conflict exists. Below is an example of our geometry work sheet. In the wing geometry calculation both canard and yield; were inputted in case we need extra lifting or control surfaces: but because we did not home canard # Div/0 appeared for canard configurations.Figure VI.2: Sample image of the ‘Geometry’ worksheet of Sizing CodeThe sizes of the engines were calculated based on the series of similar turbofan engines previously used in the business aircrafts. The diameter and the length of the engines were calculated from the amount of trust each engine needs to generate. Also, some of the drag coefficients related to the geometry of the aircraft were also calculated in geometry worksheet, and the values were then used in mission details worksheet for a more detailed calculation of the drag. The finalized 3-view of the aircraft was then transferred to the main page as an output.The constraint diagram worksheet is generated based on the constraint diagram tools provided. The input value of the cruise Mach number from the main page was used in this worksheet to generate 5 constraint lines. These lines were used to calculate the wing loading and thrust to weight ratio. These values were also copied to the main page. The thrust to weight ratio used in the finalized concept was 0.36 and wing loading of 92.6 lb/ft2 was used. Figure VI.3: Sample image of the ‘Weight’ worksheet of the sizing codeThe weight worksheet is where empty weight, fuel weight and the total weight of the aircraft are calculated. The initial guess of total gross weight is used to calculate the calculated gross weight of the concept. When the guessed total gross weight is different from the calculated value, the iteration process is used to match the guessed value to match the calculated value of weight. The equations from Raymer’s Aircraft Design: A Conceptual Approach Fourth Edition were used to calculate the various components’ weights. Equations for Cargo/ Transport weights were used; however, because these equations are outdated, validation of the equations was necessary. The validation of the sizing code will be discussed in the later part of the report. The center of the lift of the aircraft was also calculated in the weight worksheet. Miscellaneous components weights were found from the various data sources. For instance, weight of APU, passenger seats, were found from various manufacturers’ websites and historical data. Figure VI.4: Sample image of the ‘Airfoil’ worksheetThe airfoil worksheet was used to define the geometry of the airfoil according to the NACA 4-digit airfoil series. The original intention was to generate drag polar of the airfoil within the sizing code; however, due to time constraint, the sizing code was limited to calculate the thickness to cord ratio and the cross sectional view of chosen airfoil. Figure VI.5: Sample image of the ‘mission’ detail of the sizing code.Mission detail worksheet is where drag, weight, and center of gravity were calculated at the specific points of the mission. The mission was divided into 11 sections which were initial, taxi, takeoff, climb, cruise, loiter, attempt to land, climb, divert, loiter, and land. The atmospheric conditions at different mission altitudes were also calculated. The velocities were inputted mostly from historical data. The original intention was to find the best rate of climb for the climb condition. But the calculation part was not completed hence the inputted value of the climb velocity was used. Validation & Fudge FactorFor CoDR Report, validation and fudge factor of XG Endeavour were done by inputting the data of existing aircraft and compare the empty weight. The aircraft that was used for the validation must have similar features compared to XG Endeavour. To fulfill the requirements, Bombardier Challenger 300 was used to validate the code. Basic dimensions such as lengths of the fuselage, height and width of the cabin and other 50+features of the aircraft could be found from Bombardier website. Few specific dimensions were estimated by using the picture because of the lack of information that can be obtained.After the validation was completed, it was possible to find the fudge factor by using the following equation,Fudge Factor= Actual Empty Weight of Challenger 300Estimated Empty Weight from the Sizing CodeFudge factor could be inputted in the XG Endeavour to better predict the empty weight. Once the validation was completed, empty weight based on the sizing code was predicted 17500lb, while the actual empty weight of the Challenger 300 was 18500lb. Based on these data, fudge factor of XG Endeavour was calculated using the following equationFudge Factor= 1850017500=1.05For the reference, specific details of dimensions of the Challenger 300 are shown on the following page.Figure VI.6: Code constants worksheet as seen in the sizing code.Also, the picture shown below is the comparison between the drawing of Challenger 300 from Bombardier website and drawing from the sizing code.Figure VI.7: Sample image of the plane sketch from the sizing codeFixed design parameter values Drag PredictionDrag was predicted by using the subsonic drag build-up methods. Three different types of drags had to be determined in order to determine the drag as a whole: induced drag, parasite drag, and wave drag. To calculate the estimated parasite drag, all component drag values were added up by using the equation on the following page (equation VI.1).CDP= i=1# compKiQi CfiSwetiSref Equation VI.1where Sref is the reference areaCf is the skin friction coefficientK is the form factorQ is the interference factorThe net surface wetted area was found by summing up all the surface areas using the equations stated in the lecture notes given. In addition, form factor and interference factor were calculated and assumed respectively. Because the atmospheric conditions are different due to the different altitude the aircraft flies at, the drag was calculated for each of the mission segments. The drag was calculated for each mission segments separately because atmospheric conditions vary due to altitude. Below is the table containing the values of drag calculated for each of the segments of the flight mission:Table VI.2. Values of Drag for the aircraft designsTotal CDDragTaxi0.1583220Take off0.0516381Climb0.0703397Cruise0.0673709Loiter0.0535273Attempt to Land0.0478049Climb0.0497239Divert0.0594366Loiter0.1112753Land0.0862528Results of engine modelingEngines were selected then analyzed for the project. Details on the actual engine selection can be found in the later ‘XI. Propulsion’ section of the report. Some details for the two different engine sizes were identified. The 2000 pound thrust model was not scaled because data is already available for this model. The 2000 pound thrust model has a bypass ratio of 2.9, takeoff thrust of 2050 pounds, and a compressor pressure ratio of 24. In order to calculate the data for the 6800 pound thrust model, an Excel sizing routine was used to scale the engine for the appropriate thrust level. The 6800 pound thrust model has a bypass ratio of 3.2, can produce 6800 pounds of thrust, and has a compressor pressure ratio of 26. In order to scale the engine, several assumptions were made. First, technological improvement factors were taken into consideration to determine what engine performance would be like in the year 2020. Since the 2000 pound model was already the correct size, no efficiencies were needed and the larger engine was simply scaled from the smaller one.Several sources were used to generate thrust available and thrust required for the flight of the final designs. These plots show velocities will be attainable in certain configurations at specific altitudes. In order to calculate the thrust available, a basic cycle analysis with the atmospheric conditions at various altitudes were inputted. The results of the calculations are shown below for takeoff, several altitudes that a typical mission climbs through, and for the cruise condition:Figure VI.9: Thrust vs. velocity at take off.The takeoff diagram shows that thrust available exceeds thrust required up to the takeoff speed of 120 mph. Therefore the propulsion system is sufficient for takeoff.Figure VI.10: Thrust vs. velocity at 15,000 ft segment climb.Figure VI.11: Thrust vs. velocity at 25,000 ft segment climb.Figure VI.12: Thrust vs. velocity at 35,000 ft segment climb..The above three diagrams are for the climbing configuration at three different altitudes. All show that thrust available exceeds thrust required up to about 480 mph. Therefore, the maximum climb speed is approximately 480 mph.Figure VI.13: Thrust vs. velocity at 45,000 cruise.At the cruise condition, thrust available always exceeds thrust required with three engines running. After shutting down the big engine to conserve fuel, the thrust available drops significantly. With two engines running there is enough thrust available to enable the aircraft to travel up to 530 mph which exceeds the design cruise speed of 484 mph.Mission modelingMissionTaxiTakeoffClimbCruiseLoiterAttempt to LandClimbDivertLoiterLandAltitude(ft)01000040000450001600020001000040000160002000Accumulative Distance(nmi)0.7582.75832.763443346334643480368037003700A/C Weight(lb)29093289452654426463262472614226003259582592523732Fuel Weight(lb)9837968972887207699168866747670266694476The aircraft will take off at sea level, and climb to cruise altitude of 40,000 ft. when it reaches 40,000 ft, aft fuselage engine will shut down to save fuel. The thrust will be sufficient with 2 smaller engines mounted on horizontal stabilizer for the cruise segment of the mission. As the aircraft reaches the vicinity of the destination, it will descent to loitering altitude, then attempt to land. In case of possible situation when aircraft cannot land at the primary airport, aircraft will climb and cruise to the secondary airport with most efficient fashion. The aircraft is designed for maximum range of 3700 nmi with maximum payload of 9 passengers and their baggage. - VII. Major design trade-offsTrade off overviewTable VII.1: Concept options (overhead view)Previous conceptsFinal settled designThe most important design requirement that NASA proposed is to reduce the fuel consumption of the aircraft by 40%. This requirement forced the team to work very hard in changing various configurations from modern aircraft. The group thought of some possible methods to save fuel efficiency. The hybrid engine configuration is one of the few design ideas of which the team explored during this process. In the SDR report, the team considered applying one turbofan engine and two turbo prop engines, turbofan being used during takeoff and turboprops for cruise. Ultimately, this idea was not used due to its significant loss in loss of the cruise speed and little significant gain in fuel efficiency. Instead, three turbofan engines of different sizes were implemented in the final design. Two smaller engines will be used during cruise segment of the mission, and one bigger engine will be used for takeoff. In this way, fuel consumption may be reduced by approximately 25% to 30% even with today’s technology. In addition, with the constantly advancing and improving trend of engine technology, the team is confident that 40% reduction of fuel consumption can be achieved by 2020. Turbo prop concept was under consideration until the last minute of the design process. It was very attractive concept because of its fuel efficiency compared to the turbofan engines; however, because of its slow cruise speed of Mach 0.6, it will not be able to perform as fast as competition aircrafts. Also the fuel saving will not be great compared to other turbo prop engine mounted engine aircraft such as Beechcraft Kingair 350i. The unducted fan engine idea was not used in the final concept because of increase of noise. The supersonically spinning fan blades will create shock wave which will increase the noise and possible exceed the minimum noise requirement. Also because there were not enough data we were able to obtain, enough engine specification data, unducted fan engine was less attractive despite the fuel efficiency. The issue of not being able to turn the engine after once it is shut off has to be addressed. However, by 2020, the engines will be reliable so that engine can be shut one and off as it is needed. We believe it will be able to operate similar to the gasoline engines in the hybrid cars where gasoline engines are not turned on while battery powered motor is running. The NOx emission will be decreased from less fuel used during the mission. NOx is a byproduct of combustion of fossil fuel. Therefore, with close to 40% fuel reduction, NOx will be successfully reduced. Previously we were considering employing catalytic reduction technology on aircraft for the first time. However, because of weight increase and reduction of thrust, the concept of employing catalytic reduction was not implemented in the final concept.The emergency door was removed in the final concept because, for the size of the aircraft, emergency door is unnecessary and will increase structural weight needed for reinforcing the pressurized cabin.The crucifix form horizontal stabilizer will have small engines mounted on the because of empennage engine. T-tail was used for one of the concepts for the previous report; however, because we have engines mounted on the horizontal stabilizers, T-tail mounted engines would create nose pitching down moment and increase trim drag during the cruise.Canard also was used for concept design, however because we believe aircraft will be successfully controlled with current size of horizontal stabilizer, extra canard control surface will increase empty weight of the aircraft as well as drag during the flight.Summary of carpet plot studiesThe carpet plot was another way to represent the constraints of XG Endeavour. Based on the current constraint diagram and the weight prediction from the sizing code, it was possible to make different plots and construct the carpet plot. Considering the important features of XG Endeavour, there were four main constraint plots:Estimated Gross Weight vs. Wing loading at;a) Different thrust to weight ratio at the sea levelb) 2g maneuver c) Takeoff ground roll d) Landing ground roll. For the Estimated Gross weight vs. Wing Loading, thrust to weight ratio at the sea level and the wing loading were 0.36 and 92.625, respectively. In order to make a carpet plot, both thrust to weight ratio at the sea level and wing loading varied by around QUOTE 20%. For easier calculation, thrust to weight ratio at the sea level (T/W) was estimated at 0.4, while wing loading was estimated at 90. The final result of carpet plot of XG Endeavour is shown below.FigureVII.1: Gross Weight vs. Wing Loading based on various thrust to weight ratio at the sea level.For the next step, estimated weights at takeoff ground roll and landing ground roll were calculated. These calculations were made from first taking exponential equations based on the takeoff ground roll of 4500ft and landing round roll of 2500ft, then followed by using those equations to substitute the values of wing loading. A new graph was generated by plotting these new results with the with the previous carpet plot above. The ultimate carpet plot of XG Endeavour is shown below. FigureVII.2: Ultimate Carpet Plot of XG Endeavour.From the carpet plot above, it was possible to ensure that estimated gross weight calculation was reasonable. Also, it is possible to conduct the off design. - VIII. Aircraft description Dimensioned three-view, to scaleBelow is the 3-view documentation of the XG Endeavor modeled generated in Catia. FigureVIII.1: Scaled 3 views diagram of the XG aircraft with retractable ducts open.The exact values of surface location and center of gravity can be found in the sizing code section of the report. The following figures are just brief overviews of the sized representation of the airplane 22”Wing Leading Edge36”Tail leading edge37”Vertical Stabilizer38”Tail Mounted Engine40”Center Engine50”Total aircraft lengthFigureVIII.2: Overhead dimensions of the XG EndeavourFigureVIII.3: Front view dimensions of the XG Endeavour.3590925114300 Representative internal layout FigureVIII.4: View of the cabin cross section to demonstrate depressed floor.To maximize cabin headroom, the floor of the cabin was lowered (See figure right). In order to contain all the equipment required for a comfortable flight; spaces in the nose of the craft, the aft of the craft and the wings had to be optimized. On the following page are the available images representing space availability. Not shown:The fuel tank extends to the small area under the cabin. Fuel TankWheel housingEquipment compartmentAvionics compartment and nose landing gear housingFigureVIII.5: Representational internal lay out overhead.FigureVIII.6: Generated 3D view of the available space in the aft of the aircraft. -571504445Engines had to be supplied with both fuel and a steady free stream of air when operated. For all three engines, fuel would be pumped from the fuel reserves in the wings, tail and lower aft storage of the cabin into the engines. Those engines mounted on the tail were located in positions with uncontested access to air. The center engine, however, had to have a duct designed to feed the air. To create laminar flow, the ducts had to be sized and contoured appropriately. Raymer (pg 458) has an example of a duct for an air-fed breathing engine as well as explains the methodology and requirements of duct sizing. Because the duct demands specific dimensions, restrictions for the useable space are then imposed (see left). In Figure VIII.6, The image shows the aft engine, the air duct, area for fuel reserve and fuel pump, and area for other components in red, green, salmon and purple respectively. Careful arrangement of all other equipment (APU, air conditioner, motor to close and open air ducts, etc.) must be fitted within these constraints. In previous business modeling reports, it was stated that the internal layout of the aircraft was customizable. As a result, the internal cabin layout could be altered, provided that the center of gravity of the entire internal layout remained between 22-24 ft from the nose of the craft and within the weight tolerance of 300-350 lbs (See Appendix B: the ‘Center of gravity’ section ‘Sizing code’ for details). Below is the layout of the aircraft test case. In this case, the moment of each individual seats and other equipment for validation. For the test case, the cabin center of gravity was located 22.75 ft from the nose of the craft, and the weight of the interior totaled 306 lb. Below is the floor plan of the cabin interior generated in Catia for the test case.Figure VIII.7: Cabin floor plan used for project analysis. Dimensions are in feet. As a professional business jet, the number of expected passengers falls within the range of 6-10 people. Above, in figure 6, is a possible 8 passenger, 1 crew floor plan of the cabin layout. During taxi and take off, the lavatory seat can double over as seat.The lavatory location and exit locations are fixed and cannot be moved, but any other aspect of the system is changeable. - IX. Aerodynamic design details / justificationThe airfoil was selected from historical data. Although it is not the optimized airfoil selection, the aircraft was designed with NACA 2414 airfoil. Technical data of NACA 2414 airfoil is attached in the drag polar section below. During the takeoff and landing, the aircraft requires CL= 1.6 which is more than NACA 2414 can produce. Therefore, high lifting device such as flaps and slats are considered to achieve required CL. Figure IX.1: High lifting devicesDrag PolarDrag polar is a way to represent the coefficient of lift and coefficient of drag of the airplane. For XG Endeavour, NACA 2414 was used for the main airfoil and the drag polar is shown below.Figure IX.2. Drag Polar for NACA pare to XG Endeavour’s Reynolds number, which is around 1.02e6, purple line of the drag polar was used for the drag polar of XG Endeavour. For the landing, parasite drag and induced drag at the landing were added to the cruise drag, while for the take off, parasite drag and induced drag at the takeoff were added to the cruise drag. Resulting drag polar at takeoff, cruise and landing are shown below.Figure IX.3. Drag Polar of XG Endeavour - X. Performance A very important aspect of an aircraft’s performance is the aircraft’s maneuverability at various flight speeds, and the aerodynamic loading that is applied to the aircraft during various points in the design mission. The aircraft’s abilities and limitations regarding its performance provide for a set of parameters that comprise what is known as the flight envelope of an aircraft. One technique that allows for a better understanding of the aerodynamic loading on the aircraft is the V-n (also known as V-g) diagram. The V-n diagram better illustrates the performance characteristics of an aircraft during various different maneuvers, and different flight speeds and aircraft orientations. A standard V-n diagram of an unknown aircraft is shown below in Figure X.1. Figure X.1: A standard V-n (or V-g) diagram is shown to provide a visual understanding of aircraft loading at various flight speeds. The V-n diagram reveals much about an aircraft. The region between the red lines, including the green and yellow ranges, is a typical flight envelope. The envelope is bounded on the left by a pair of curves. These curves constitute the stall velocities at various load factors. At these speeds, any load factor above the corresponding value is aerodynamically unavailable. The point marked maneuvering speed is the speed at which the aircraft may perform a maneuver that results in the maximum possible load factor. Any maneuvers made to the left of the curve result in a stall of the aircraft. This holds true for both the upper and lower sections of the curves that make up the left bound of the envelope. The envelope is bounded at the upper limit of the envelope by the maximum positive load factor (which in this case, is approximately 4.5g, according to the figure). It follows then, that the lower bound of the flight envelope is the negative maximum load factor (in this case, nearly -2g). Between these load factors, the stall speeds, and the maximum structural cruise speed is the normal operating range. Unless the aircraft comes under extreme aerodynamic duress for one reason or another, the entire flight design mission should be contained within the normal operating range. If the aircraft’s maneuvering results in a load factor that exceeds either the positive or negative limit shown in the V-n diagram, structural damage is possible. Subsequently, further disregard of the load factor limits and maneuvers that exceed the upper and lower bounds of the acceptable operating ranges will eventually lead to structural failure in one or more critical components of the aircraft, although the flight crew and passengers will most likely experience black-out, or red-out before structural failure occurs, depending on whether the upper or lower limit is exceeded. So, exceeding these parameters may lead to the loss of the aircraft, or even human life, for a number of reasons. This is what makes the V-n diagram and its parameters so important. The right side of the flight envelope is bounded by the maximum flight speed of the aircraft, which may also be considered the aircraft’s dive speed. Even steady level flight at speeds greater than the maximum structural cruise speed but less than the dive speed is not advised. Flight at these speeds cannot be sustained for extended periods of time without eventual structural damage. When the aircraft exceeds the right-most limit of the flight envelope, structural failure will occur. It is crucial that the aircraft not exceed any of the boundaries of a V-n diagram. Figure X.2: Team XG's V-n diagram is shown to allow for the visualization of the aircraft's flight envelope.232410064770Figure X.2 shows the V-n diagram for Team XG’s flagship business jet. The V-n diagram allows for several performance specifications to be illustrated, and allow for a visual interpretation of the aircraft’s limitations. For steady level flight at exactly 1g, the aircraft’s normal stall speed occurs at 220 ft/s. At the upper limit of the stall curve, the maneuvering speed can be determined from the graph. Maneuvers creating a maximum load factor of +3.333g may be performed once the aircraft reaches the maneuvering speed of approximately 400 ft/s. The stall curve intersects the lower bound at the same speed as the normal stall speed, as the maximum negative load factor of the aircraft is -1.000g. Between these curves and our design cruise speed of 710 ft/s is the normal operating range for the Team XG aircraft. Between the upper and lower maximum load factor limits, the design cruise speed, and the aircraft’s maximum dive speed of 1020 ft/s is the caution range for our aircraft. These values were determined using a number of other predetermined performance specifications. Some of those specifications and the aforementioned flight characteristics determined from the V-n diagram are tabulated in Table X.1. In Table X.1, the aircraft’s normal stall speed, dive speed, cruise speed, and maneuvering speed (denoted as the stall speed at maximum positive load factor in Table X.1) are shown, courtesy of the V-n diagram. However, Table X.1 also provides the take off distance, landing distance, best endurance velocity, and best range of our aircraft. The takeoff and landing distances of 4000 ft and 2500 ft, respectively, are important to note as they will determine which airports our aircraft will be able to access. Also very important to note are the best range and best endurance velocity. Table X.1: Performance specifications for Team XG's business/personal aircraft are tabulated to provide clarity.Performance SpecificationsValues (units)Takeoff Distance4000 (ft)Landing Distance2500 (ft)Best Range3700 (nmi)Best Endurance Velocity710 (ft/s)Cruise Speed710 (ft/s)Dive Speed1020 (ft/s)Normal Stall Speed220 (ft/s)Stall Speed at Maximum (+) Load Factor400 (ft/s)The best range of 3700 nmi for the aircraft is, coincidentally, the maximum design mission range of our aircraft, which can be achieved at the best endurance velocity. Similarly, the best endurance velocity is our aircraft’s cruise speed. The best endurance velocity of 710 ft/s was chosen for the design cruise speed for the obvious reason that it is the most fuel efficient cruise speed, and at the same time, a timely arrival at our customers’ destination is also ensured. These parameters were some of the focal points of Team XG’s performance goals. - XI. Propulsion The final design utilizes three turbofan engines. Two of these turbofans can produce up to 2000 pounds of thrust, while the remaining one can produce up to 6800 pounds of thrust. To incorporate these engines into the final design, they were modeled after an existing turbofan engine. The baseline engine was the HF 120 turbofan, which is currently manufactured by GE Honda Aero Engines. Below are a picture and a schematic of the 2000 pound thrust version of the engine and the dimensions of that model:Figure XI.1: (left) Photograph of the GE Honda Aero Engine, the HF 120 turbofan and (right) a simple schematic of the same engineWhat makes this engine stand out as our final engine choice is environmentally friendly design which minimizes noise and emissions. In addition, it is very fuel efficient sporting a low SFC, and it has a very high thrust to weight ratio. It was very important to generate thrust available and thrust required versus velocity at various altitudes for the final design. These plots are crucial because they show what velocities will be attainable in certain configurations at specific altitudes. The data for these plots was generated from several different places. First, the thrust required was available from the team’s sizing code. In order to get the thrust available, some calculations were necessary. Using a basic cycle analysis with the atmospheric conditions at various altitudes inputted, it was simple to calculate the maximum thrust the engines could produce at a given velocity. The numerical results of the analysis are shown in the “VI. Sizing Code; results of engine modeling” section of the report.Overall, the propulsion system for the aircraft is very effective. It meets the thrust required criteria at all altitudes and configurations. Stability and control-related propulsion considerations, including emergency conditions, are found in section XIV below. Also, the three engine configuration allows for unnecessary thrust to be cut out while cruising therefore saving fuel. The propulsion system also meets all noise and emissions requirements. - XII StructuresStructural integrity is a major design factor in aircraft construction. But before the structure can be designed, one must determine the types of loads that would be imposed on the aircraft. In general, each part of the aircraft is subject to many different loads. In the final design structure, there might be thousands of loading conditions. Of them, hundreds could prove to be critical for parts of the airplane.{cite 1} Generally, the strain allowances on aircrafts are defined by load factor, n. In simple terms, load factor is defined as the component of force on a structure divided by the aerodynamic weight. The factor of safety, which is the ultimate load over the intended load, is also a crucial consideration in aircraft design. The FAA establishes two kinds of load conditions:{Site 1} Limit loads are the maximum loads expected in service. FAR Part 25 (and most other regulations) specifies that there be no permanent deformation of the structure at limit load.Ultimate loads are defined as the limit loads times a safety factor. In Part 25 the safety factor is specified as 1.5. For some research or military aircraft the safety factor is as low as 1.20, while composite sailplane manufacturers may use 1.75. The structure must be able to withstand the ultimate load for at least 3 seconds without failure. In addition to these strength factors, designers have to also deal with fail-safety, fatigue, corrosion, maintenance and inspect-ability, as well as production value. Modern airplane structures are usually designed as load carrying shells that are reinforced by a series of frames, formers, longerons, and other ridged shapes that make up the airplane skeleton. Skin stringers, spars, ribs and boxes then enforce the thin airplane skin. Figure XII.1: Overhead view of a wing showing ribs, spars, and cavities for appliance storage. 2152650123825To the right is a diagram of how a wing’s structural support can be arranged to contain components necessary for the aircrafts operation. Making the most of negative space in airplane control surfaces is crucial to increase aerodynamic efficiency as well as allow more room for additional freedom. Below (Figure XII.2) is an example of a typical airplane airframe. Notice that the windows, the doors and the cockpit glass have a different method of re-enforcement than the rest of the craft. Figure XII.2: The structural skeleton of a similar type aircraftA typical real world airplane will have several areas of which calculations are crucial. For the sake of the design project, however, these calculations will be generalized.Configuration layout and highlightsIn order to maintain a safe and light weight aircraft, the airplane’s skeletal structure, or airframe, has to be built with the loads in consideration. Such loads not only include the internal mass forces (such as payload, fuel load, equipment and passengers etc.) but also include flight forces, and ground forces. Ground forces induced by taxiing and landing can be easily estimated, given the airplane’s weight and performance parameters. Flight forces, such as propulsion thrust, lift, drag maneuver and wind can be mathematically estimated and tested digitally or in a wind tunnel. Important Load PathsBecause very few details are known about the particular structural components of the Endeavour, a general sketch was created to visualize the airframe of the craft. Below are the estimated locations of the ground forces on the aircraft from an overhead view. Figure XII.3: Force distribution sketch and estimated locations of major stress factors.Here, ‘areas of concern’ are marked. These areas were considered important to focus on because of their unusual load situation. Moments formed on the joints or connectors of the wings and stabilizing surfaces of the aircrafts (they are marked with tan hatching) must be addressed through re-enforcement of the main bulkheads at the wing roots and areas like the joints for the landing gear, where unusually high of pressure is applied, has to be resting on framework that is design to properly distribute the entire craft’s weight. In general, moment M is a function of force, and distance. For a cantilever, the moment created on the attached M= F'xdxEquation XII.1Where F is the cross product of force per unit length, or, F(x) = force x distance. When designing locations where airplane features intersect, this equation is invaluable. With these forces in consideration, a skeleton of the major ribs, stringers, formers and frames was then sketched. The results of the analysis for the theoretical load paths and the locations of the applied forces have been overlaid the aircraft diagram below. Figure XII.4: Major paths of load and the location of crucial formers, ribs, longerons and stringsTo further reinforce the airframe, other frames and formers are needed. Ideally the entire plane will be manufactured by the lightest and strongest, durable material such as composite structure. Below is the example of the framework of the aircraft. Figure XII.5: Generated 3D model of the formers and stringers as they would be located on the actual craft?Wing-fuselage intersection2886075708660The aircraft wings are mounted under the fuselage and secured through two of the main supporting formers. Mounting the wings low on the aircraft yields several benefits. One such benefit is the ability to reinforce the joints connecting the wing to the fuselage without disrupting the pressurized portion of the aircraft. The structures must be strong enough to hold the wings and the aircraft together, despite the forces generated by the lift of the wing at the wing root. Most aircraft built in this formation today use a carry-through method of securing the craft to the airplane similar to that shown in the image to the right (Figure XII.6). Figure XII.6: Cross section demonstrating how a low wing can be affixed to the aircraft.Engine mounts and landing gear integrationBecause of the locations of the engines, creative measures were made when addressing the engine mount options. As stated in the previous section of the paper, the engine collection currently consists of three separate turbo-fan engines. There is one central turbofan engine which can produce 6800 lbf of thrust, and two cruise fans that can produce a thrust of 2000 lbf. The two smaller engines are mounted through the horizontal stabilizers. Options on how an engine can be mounted through the control surface are demonstrated in the image below. Figure XII.7: Side view of the horizontal stabilizers with engines.With both, the engines are mounted relatively close to the fuselage to reduce moment. However, unless mounted exactly on the centerline, moment about the c.g. can never be completely eliminated. Further analysis would be needed to determine the most efficient method to keep the engine secure. When more details (such as control surface sizing requirements) of the aircraft become available a more definitive decision will be made between the two options. General analysis tactics can give an estimation of the structural stability required to resist the force caused by the thrust of the engine. Below is an example of a simple truss-type structural option of re-enforcing the stabilizers of the tails with the locations of the engines currently consideration. Figure XII.8: A simple single plane truss formationThe center engine will be mounted using ring supports to allow for modular operation and maintenance. The engine flow duct must be accommodated for. In Raymer (pg 315), dimensions of the duct are stated, the structure then has to work around the required lengths for laminar airflow. Similarly, the structural requirements of the landing gear will be measured. To reduce moment, increase stability and increase the use of negative fuselage space, inwardly-retracting wheels (i.e. toward the centerline of the aircraft) were used. By having wheels that swing outward during landing (like that shown in the image below, far left) you increase the yaw stability of the aircraft while taxing by creating a stable triangular base. FigureXII.9: examples of inwardly retracting landing gear (right-hand figure not to scale)The rear landing gear is located on the intersection of a main wing and a fuselage. The stringer itself is slightly located on the frame of the aircraft of which the empty weight center of gravity is. This way, the center of gravity is forward of the rear landing gear at all times. Below are some example diagrams of the situation and methodology on how to solve for their location. Figure XII.10: (top) Wheel loading geometry taken from Raymer page, (bottom left) side view and (bottom right) bottom view landing gear locations in relationship to the center of gravity.Materials Historically, airplane airframes have been designed from materials chosen for their strength. But as material production evolves, it becomes more safe and affordable to make modern craft out of materials with several other properties in mind such as memory-shape or thermal-expansion capability. When making material choices, it is important to pay attention to more than just the strength to weight ratio. On-board materials often have unique applicability, and designers will often select them based on their: Fracture toughnessCrack propagation rateNotch sensitivityStress corrosion resistanceExfoliation corrosion resistanceGamma Titanium Aluminides (GTA), a new class of lightweight alloys has been in development over the past few decades, and are said to sustain temperatures higher than 600oC. According to InnovaTiAl, a European commissioned project which addresses “NanoTechnologies and nano-sciences, … materials and new production process…”, gamma titanium aluminides have proven to meet the requirements for aeroengines and could easily be applied for use in airfoils, wheels, pylons and other types of supports. GE’s announcement in 2005 to use intermetallics in the GEnx engine underscores the appeal of such materials. Although the cost to implement such materials have been a major hurdle for existing aircrafts, it is likely that this material will be common in crafts of the future, including the Endeavour. Other materials that would prove useful for the future of this craft include the more common composites. Among the many materials available, below are a few possible options and their use. Table XII.1: Material selections for structural use.Material type/nameUseFiber Glass Control surfaces such as wings and ailerons and ruddersCompositesWings and fuselage skin. ThermoplasticsRib and frame structures in the craftAluminum based alloysEngine composition and general supporting structures. Currently, the design team is weighing heavily on the idea of using aluminum-based alloys, particularly the GTA type materials described above. Currently, making the entire aircraft from GTA material is impractical largely because of the manufacturing cost and available data. If, in the future, the availability and practicality of using the GTA type materials becomes feasible, then the option to make a more considerable portion of the superstructure of the aircraft from this material will be reevaluated. - XIII. Weights and balance The estimation of the weight of a conceptual design aircraft is a critical part of the design process. The group used the sizing code and equations in Chapter 15 in Raymer’s book to find the best weight of the major components. The aircraft was divided into 3 major sections: fixed structures, inner structures, and payloads. Fixed and inner structures were catalogued as structures that do not move or change weight during the flight. Payload and fuel were of major concern for computing c.g. travel information. The summation of these groups provides the total empty weight. The fixed components section contains the wing, horizontal tail, vertical tail, fuselage, main landing gear, and nose landing gear. The total weight of the fixed components is 13,125 lb. The inner structures group contains avionics, seats, and compartmental fixtures; it weighs 660 lb. The payload group consists of passengers, crew, and luggage. The total weight of the payload group is 2,223 lb. Total empty weight is generated by summing the weight of structures, propulsion, and equipment groups; it sums to 16,800 lb, which is comparably low to other current-generation aircraft of similar size including the Gulfstream G250, at 24,150 lb, and the Bombardier Challenger 300, at23,349 lb. The gross weight of the aircraft is 29,800 lb.Overview discussion on prediction methodsThe weight was calculated using 22 equations with 90 constants in Raymer to compute the best estimate of the empty weight. The component weight equations are put into the sizing code and provide an estimated empty weight. Table XIII.1: Component chart of location, weight and moment of each component?Name#X-Location(ft)Weight(lb)Moment From Zero PointFixed componentsWwing133.1411157.338355Whorizontaltail140.725465.7118966Wverticaltail139.809379.6815115Wfuselage144.54149.4184649Wmain landing gear136.141798.1728847Wnose landing gear111214.152355.7Wnacelle group14300Wengine controls14340.41737.2Wstarter(pneumatic)14384.6583640.3Wfuel system135.141283.589965.4Wflight controls111456.015016.2Wapu installed1431988514Winstruments111105.511160.6Whydraulics133.141102.643401.5Welectrical133.141586.2619429Wavionics17.333311558470.2Wfurnishing233.141269.558933.4Wairconditioning240.725211.338606.4Wanti-ice233.14159.6121975.6Whandling gear235.1418.9418314.23Wsolarfilm144.530013350Wengine3431861.580044Inner StructureAvionics1202344680Seat1116.734567.8Seat2117.734601.8Seat3125.4634865.64Seat4130.335341031.4Seat5115.6334531.42Seat6130.335341031.4Sofa121.6534736.1Laboratory133.9341152.6Galley11334442Environmental Control12030600PayloadPilot252341170Passanger1116.71953256.5Passanger2117.71953451.5Passanger3125.461954964.7Passanger4130.3351955915.3Passanger5115.631953047.9Passanger6130.3351955915.3Passanger7121.651954221.8Passanger8121.651954221.8Passanger9121.651954221.8Luggage211.72342737.8Total ?32.31215680518210Center of gravityLocation (ft)Weight (lb)Figure XIII.1: X-location CG travel diagram plotted in relationship to the airplane sketchAbove is the visual representation of the center of gravity, the location of which changes according to fuel consumption, as well as payload and crew allocation. Since the location of center of gravity (c.g.) shifts during different stages of the mission, it was important to mark the path of the variations, known as the CG travel diagram (Figure XIII.1). As shown the figure above, the location of center of gravity ranges from 32.6 to 34.5 ft from the nose. This travel distance was calculated by locating various weights of components, which is important to know in order to ensure that the center of gravity is always in front of the neutral point when in flight. Because the XG Endeavour engines are placed so far the aft of the aircraft, the initial location of center of gravity is placed further from the nose of the aircraft than on convention jets of similar sizes with a more traditional configuration. In this case, the center of gravity was calculated to be 32 feet from the nose. The initial location of the c.g was calculated by calculating the moment created by various components. To simplify the problem, the aircraft was divided into 3 major sections, as implied in table XIII.1: Fixed structuresInner structuresPayloadsFixed and inner structures were structures that do not move or change weight during the flight. Payload and fuel were major concern for computing center of gravity travel information. The weight of the aircraft will change constantly during a mission and so important points were identified to find location of center of gravity. The mission was divided into 12 major segments:Take offGear up Fuselage tank use Wing tank use Gear down LandReserve fuelPassenger disembarkCrew disembarkAdd fuel Crew boardingPassenger boarding During the design mission, the payload and fuel weight changes. First, the initial center of gravity was located at 32 feet from the nose, when the aircraft was at its maximum gross weight of 29000 lb. During the take off, the fraction of fuel is used from the tank installed in fuselage under the cabin. During the climb and cruise, more of the fuel stored in the wing box is consumed. Because these fuel locations are behind the initial location of the center of gravity, the c.g. will travel forward as the weight of the aircraft decreases. Therefore, as the travel progresses the aircraft becomes more stable. Because more stability is advantageous during the cruise, this situation is very favorable. With more stability, the passengers will then experience a smoother, more comfortable ride. When passengers and pilot disembark the weight is at its minimum, with only trapped fuel and oil. At this point, the weight of the aircraft is approximately 21500 lb. The fuel is added, and passengers and crew board for the next flight. Below are numerical representations of the X-location of the Center of gravity, and the total weight of the aircraft and its contents.Table XIII.2: The values of CG, and Weight for separate segements of the missionWight SegmentCG (ft), from the noseWeight (lb)Take off32.129163Gear Up32.129011Fuselage Tank32.026468Wing Tank32.025951Gear Down32.025916Land32.025916Reserve Fuel31.923669Passenger Off33.021680Crew Off33.421446Add Fuel33.426940Add Crew33.027174Add passenger32.129163 - XIV. Stability and control The basic idea behind providing static stability for the aircraft is to come up with a design which keeps it from yawing, rolling, or changing pitch too easily as forces from the surrounding fluid. Figure XIV.1: Static stability diagrams for equilibrium flight, unstable flight and neutral static stability.Much of this has to do with the plane’s static margin, which is a percentage of the distance between c.g. and neutral point versus the mean aerodynamic chord (6.6 ft) that tells how far back the neutral point is from the center gravity (32 ft down from the nose). Historically, we found that the static margin is between 4 and 16 percent. Lower static margins makes the plane more responsive to the pilot’s control, while higher static margins result in a plane which is more stable, but slow to respond. For low altitude flights (during takeoff and landing) responsive controls are desired, and for cruise a comfortable, ‘smooth ride’, more stability is wanted. Because of this, a static margin of 18% was chosen, thus putting the neutral point 1.6 ft behind the center of gravity.Figure XIV.2: Wing and horizontal stabilizers diagram and areas of concernTo size the elevators, we used conventional business jet values found in Dan Raymer’s textbook. Each elevator is 90% of the tail span and 32% of the tail chord; thus, each elevator has a chord length of 1.43 ft and a span of 10 ft. this gives them a planform area 14.3 ft2. We were also able to find trim diagrams for elevators with these (percent-wise) dimensions, shown below:Figure XIV.3: These trim diagrams for NACA 1410 were taken from historical dataPreparations for malfunctions and extreme situationsIt is very important to consider potential issues that may occur during flight when working the plane’s stability. Engine failureIn a situation where one of the exterior engines fails resulting in only one of the wing engines proving thrust, a large moment which makes the plane unstable in terms of yaw. This particular problem can be directly resolved by restarting the center turbofan. This eliminates the stress of single outboard engine yaw on the structure of the aircraft. In the event of the center engine failure, the outboard engines will need to provide adequate thrust to gain maneuverable altitude and land, as this would be the only phase of flight affected by such a failure.Propulsion analysis in the “Results of engine modeling” section of the “VI. Results of aircraft sizing” portion of the report (above) discovered that this engine alone could produce the thrust required for safe controlled flight long enough to make an emergency landing. In the event that the center engine is to lose thrust, the two wing engines could also perform all maneuverability requirements of the airplane. In this case, assuming the aircraft would need to rise to a height of 25,000 ft (in the event of a missed landing, for example) the smaller engines provide the necessary thrust to do so. Even though the fuel consumption and time to climb would be increased, the aircraft would be able to adequately perform. Extreme crosswindAnother major concern was that of ‘Crosswind landing’. In this situation, a crosswind which alters the aircraft’s orientation is applied. Such an impact makes yaw control much more difficult, which is especially a problem when the jet prepares for landing. The control surfaces must be sized such as to be able to adequately combat the induced sideslip, and responsive enough to make the final adjustments to align properly with the runway Historical values taken from Dan Raymer’s text book were used to size the rudder and ailerons. The rudder has a span that is 70% of the vertical stabilizer’s and a mean chord length that is 30% of vertical stabilizer’s. Both are 7.07 ft, giving the rudder a span of 4.95 ft and a chord of 2.12 ft. The planform area is 10.5 ft2.The total span is 51.82 ft, and the ailerons are sized such that they comprise 37% of the span. They were sized to have chords that are 20% the length of the wings’ mean aerodynamic chord, which, as stated earlier, is 6.6 ft. As a result, the ailerons are designed to have 9.6 ft spans and 1.33 ft chords. They each have planform areas of 12.73 ft2 and aspect ratios of 7.22. - XV. Noise Aircraft noise is produced by aircraft during all phases of the flight; on the ground when parked, taxiing, during takeoff, landing, and while en route. One of the primary goals of the design is to reduce this noise pollution as much as possible.As the engines will undoubtedly be the largest noise-producing fixtures on the aircraft, special consideration was given to their type (propfan, turbofan, or turboprop), their shrouding, and their orientation on the aircraft. The two stock HF120 turbofans are rated as individually compliant with the Stage IV guidelines, and it is expected that the single 6000 lb thrust-scaled variant will operate in a decibel band within a reasonable margin of those same guidelines. The shrouds for all three engines is intended to muffle a considerable portion of the noise produced, though at the higher-frequencies produced at takeoff they are likely to be less effective. The EPNL noise recording for any airborne platform is measured at three distinct points to reflect three important phases of flight: sideline, takeoff flyover, and landing flyover. Reducing takeoff distance and time to climb considerably decreases the Endeavour’s takeoff flyover noise signature, while the housings for all three engines are designed to reduce sideline noise. The overall noise signature estimate of the aircraft would be expected to fall between 200 and 220 decibels (cumulative). Specific data is available only on the two stock HF120 engines themselves and then scaled for the larger variant, plus aerodynamic noise attributable to the landing gear, flaps, and control surfaces accounted for by similarly-sized competitors and historical estimates.The two smaller engines are mounted rearward on the horizontal stabilizer to reduce pressure-wave noise in the cabin while also allowing for a possible future retrofit with fuel-saving propfan engines. A significant drawback of the propfan engine is its high-amplitude dynamic frequency vibration, and this is one of the reasons it was not pursued for development in conjunction with the Endeavour project. The placement of the largest engine in the baffles of the aircraft acts to reduce sideline and ground noise during takeoff, where its noise output will be highest, but this compounds the issue of internal vibration and noise. To reduce the amount of oscillatory motion transferred to the cabin and passengers, the aircraft is outfittable with an Active Vibration Control System, placed on the structural engine mounts in the case of the internal engine, and at the root of the horizontal stabilizer in the case of the two external engines. Such a system operates by producing noise/vibration at opposing frequencies to its input, effectively dampening up to 90 percent of the vibration and noise given off by the engine. This system will be especially necessary in tandem with the propfan system but adds weight and was deemed a luxury in the face of industry data which indicated passengers expect some measure of engine noise during all stages of flight. A notable feature of the AVC system is that, in addition to the overall dampening it is capable of providing, it also is tunable in that a specific region or compartment of the aircraft can be effectively quieted to near zero vibration at the expense of weaker performance in the other regions of the aircraft. This will allow for absolutely quiet video-conferences and maximally comfortable VIP accommodations in-flight. - XVI. Cost To estimate the cost to develop and manufacture the proposed aircraft, the Rand Corporation’s “Development and Procurement Costs of Aircraft model” (DAPCA IV) model was used from Raymer’s textbook. This model has been used primarily on fighters, bombers, and transports but can apply to a broader class of aircraft. The DAPCA IV model guesses the hours required for Research, Development, Test, and Evaluation (RDT&E) and production by the engineering, manufacturing, tooling, and quality control groups. This takes into account everything from technology research to ground testing. The four above control groups are multiplied by the hourly rates to yield costs. Table XVI.1 below shows a table summary of those hourly costs that are used.Table XVI.1: Hours Estimated for DAPCA IV.Engineering hourly rate$86Tooling hourly rate$88Quality control hourly rate$81Manufacturing hourly rate$73Given equations were used to calculate the four control groups along with the following: development support cost, flight test cost, manufacturing materials cost, engine production cost avionics cost. Together, these costs form the RDT&E + flyaway cost. This total cost includes all aircraft being produced. Based on the definition given in the textbook and how many aircraft could actually be produced in 5 years, 160 aircraft were determined to be produced in that 5 year period. A summary of the results is shown in Table XVI.2 below.Table XVI.2: Cost Prediction Summary Type Price RTD&E + Flyaway Cost $2.4 Billion Production Cost $15 Million Profit Per Aircraft $750,000 Breakeven Point 20 aircraft Production Run 160 units in 5 years The RTD&E + flyaway cost was determined to be $2.4 billion. This is the cost for the entire fleet of 160 aircraft to be built in a five year period. This translates to a production cost of nearly $15 million per aircraft built. Keep in mind; this doesn’t include new technology factors that may need to be added into development. The breakeven point along with the profit produced per aircraft was where the DAPCA IV seemed to struggle to capture the correct values. The equation used to calculate this was the production cost divided by the profit per aircraft. Or simply;Breakeven point= Cost of Production Profit per aircraft sold Equation XVI.2If the profit was raised then the breakeven point became a seemingly unreasonable value. From the presentation, Gulfstream mentioned that normal business aircraft turn a 30% profit while our aircraft was near 5%. To get the profit to 30%, the breakeven point would become absurdly low which doesn’t make reasonable sense. This could be due to the DAPCA IV model using empty weight as a main input into the equations and the aircraft designed has a low empty weight. After all, the DAPCA IV is just a model so it need not apply to all aircraft. This is the drawback from using given equations and historical data. Because of these issues, the profit was determined to be $750,000 an aircraft producing a good breakeven point of 20 aircraft. Also applying to the cost spectrum is the variable cost and fixed costs of the aircraft. Table XVI.3 shows the variable operating cost. Of special note, a two man crew equation was applied from the textbook and the current fuel/oil costs were factored in to produce the results. The direct operating cost per seat-mile was found to be $.22. This seems reasonable as the business jet is not meant to load up on passengers. A business jet is more for traveling in comfort and on a strict time schedule without commercial airline delays. Table XVI.3: Variable operating cost Rates Values Depreciation 6.6% / year Insurance $30,000 / year Crew $230 / block hour Fuel/Oil $1158/flight hour Maintenance $764 / flight hour DOC $2274 / flight hour DOC/Seat-Mile $0.22 Some of the determined fixed costs are shown below in Table XVI.4. These costs are going to be on the consumer that purchases the aircraft. The costs shown below are educated guesses on how much it will be to put the aircraft up in a hangar along with training and landing fee per certain airports which is based on weight. This list is not all inclusive but it gives a good idea of what is to be expected. Overall, the determined costs seemed reasonable when applied to the business aircraft model.Table XVI.4: Fixed Operating CostsType Cost Hangar $80,000 / year Training $40,000 / year Landing $386 / landing - XVII. Summary XG International, Inc. sought to create a cutting edge, environmentally sensitive yet fast and convenient air transportation system. The smartest minds available working with state-of-the-art technologies have answered with the XG Endeavour. This business jet maximizes proven designs with some of the latest advancements in noise and fuel consumption reduction and power conservation to offer the ideal, small-size, mid-range business jet. The compliance matrix can be found in appendix c. Through much research and heavy consideration of aircraft designs and features, XG International decided upon a standard body aircraft with swept back, low wings, and a cruciform horizontal stabilizer. The XG Endeavour’s three GE Aero Honda HF120 variant turbofan engines, two of which are located on the ends of the horizontal stabilizer, and a third up-scaled model located inside aft of the aircraft body, enable flight speed of Mach 0.8 and a range of 3,700 nautical miles. The majority of the cabin layout of the aircraft will be set to the customer’s specifications, which makes the XG Endeavour an elite form of sky transportation. The relatively small HF120 turbofans and the aerodynamic design make the XG Endeavour more than compliant with NASA N+2 requirements. The XG Endeavour is able to reduce fuel consumption by using all three of its turbofan engines during first and second segment climb, then turning the central engine off and relying solely on its two external engines for the cruise. The use of solar film furthers the XG Endeavour’s environmental friendliness. Though the XG Endeavour does not meet its target takeoff distance of 3,400 ft., its actual takeoff distance of 4,000 ft. allows access to nearly any small airport that the customer may desire. The XG Endeavour boasts a cruising speed of Mach 0.80, which will cover its 3,700 nautical mile range with haste. Add that to a takeoff distance of a mere 4,000 feet, and this business jet is as versatile as any other jet in its class. The two horizontal stabilizer-mounted engines are the GE Honda Aero HF120 variant turbofans. Each of these engines is capable of producing 2,000 pounds of thrust. The aft fuselage engine is an up-scaled model of the HF120, capable of 6,800 pounds of thrust. All three of these engines will be used for takeoff and the climb to the XG Endeavour’s cruising altitude, where theAPPENDIX A: BibliographyFrei, Stefan. Urban Mader. “Technology Speed of Civil Jet Engines”, Techzoom publications. 2006Findlay, S.J., N.D. Harrison. “Why Aircraft Fail.” Material Study Magazine. November 2002. QinetiQ Ltd. Famborogh, Hampshire. pp 18-25.Heet, Erika. ‘FrontRunners: Above and Beyond.’ Robb Report. Ed. April 01, 2010. New York, NY. pp 23-24Megson, T.H.G. “Aircraft Structures for Engineering Students.” Third edition. ‘Part II: Aircraft Structures’. Butterworth Heinemann, Burlington, MA. 1999. Raymer, Daniel P. “Aircraft Design: A conceptual Approach” Fourth Edition, American Institute of Aeronautics and Astronautics, Inc. Virginia, 2006.Tamura, Marcelo Satoru. “A Simplified Geometric Method for Wing Loads Estimation” 2009 Brizilian Symposium on Aerpsoace Eng. & Applications. September 14-16, 2009. S. J. Campos, SP, Brazil. Tien, John K., and Caulfield, Thomas, eds.? “Superalloys, Supercomposites and Superceramics”. New York: Academic Press, 1989.APPENDIX B: Sizing codeSee attached file. APENDIX C: Compliance MatrixRequirementTargetThresholdXG EndevourCompliantMaximum Mach Number0.850.80.82NoEmpty Weight (lb)18,50020,00016,000YesGross Weight (lb)28,00032,00029,100NoTakeoff Distance (ft)3,4003,8004,000NoMaximum Range (nmi)3,7003,6003,700YesDesign Mission Range (nmi)3,7003,6003,700YesNoise (dB)4250>42YesSeats1089YesVolume Per Passenger (ft3)656060YesTSFC (% of avg)556565YesN0X Emissions (% of avg.)255010YesCharge Time - 220V 80A* (hr)241.5YesCharge Time - 125V 15A** (hr)354Yes ................
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