AAE 451 Conceptual Design Review

Transcription

1 AAE 451 Conceptual Design Review May 6, 2010 Team 1 Alex Mondal Beth Grilliot Brien Piersol Heath Cheung Jason Liu Jeff Cohen Jeremy Wightman Kit Fransen Lauren Hansen Nick Walls Ryan Foley Tim Fechner

2 TABLE OF CONTENTS EXECUTIVE SUMMARY... 5 MISSION STATEMENT... 6 DESIGN MISSION... 6 TYPICAL OPERATING MISSION... 7 MAJOR DESIGN REQUIREMENTS... 7 SELECTED BEST AIRCRAFT CONCEPT... 8 WALK AROUND CHART... 8 IMPORTANT DESIGN FEATURES RESULTS OF AIRCRAFT SIZING DESCRIPTION OF SIZING CODE CCD Taxi and Takeoff Climb Fuel Burn Approximation Cruise Fuel Burn Approximation Landing and missed approach Approximation Reserve Approximation Constraint calculations DESCRIPTION OF SIZING CODE BWB MODELING APPROACHES AND ASSUMPTIONS SIZING CODE VALIDATION MAJOR DESIGN TRADE-OFFS CARPET PLOTS CANARD PLACEMENT TRADE STUDY AERODYNAMICS AIRFOIL SELECTION FOR WINGS AND TAILS HIGH LIFT DEVICES DRAG BUILDUP WEISSINGER STRIP METHOD PERFORMANCE V-N LOADS DIAGRAM PERFORMANCE SUMMARY PROPULSION... 35

3 ENGINE MODELING ENGINE TRADE-OFFS STRUCTURES CONCEPTUAL STRUCTURAL DESIGN IMPORTANT LOAD PATHS CONVENTIONAL CANARD DESIGN STRUCTURAL CONCEPT Fuselage Canard Box Wing Wing Box Tail Engine Pylons Landing Gear BLENDED WING BODY DESIGN STRUCTURAL CONCEPT Pressure Vessel Fuselage Wing Wing Box Engine Pylons Landing Gear MATERIALS SELECTION AND JUSTIFICATION WEIGHTS AND BALANCE STABILITY AND CONTROL LOCATION OF CENTER OF GRAVITY CANARD SIZING NOISE REGULATION AIRCRAFT CLASSIFICATION ENGINE TECHNOLOGY AND AIRCRAFT NOISE NOISE CERTIFICATION OF NEW AND UPGRADED AIRCRAFT NOISE MANAGEMENT NOISE ESTIMATES COST METHODS USED TO COMPUTE COST... 81

4 COST VALIDATION NUMBER OF AIRCRAFT IN PRODUCTION RUN ESTIMATED COST TO DEVELOP AND MANUFACTURE ESTIMATED OPERATING COST RESULTS SUMMARY AND NEXT STEPS WORKS CITED APPENDIX A: HOUSE OF QUALITY APPENDIX B: COMPONENTS OF AIRCRAFT APPENDIX C: NOISE FIGURES APPENDIX D: NOISE CALCULATION AND METHODS APPENDIX E: COST FUNCTION

5 EXECUTIVE SUMMARY The goal of Team 1 s aircraft is the construction of a business jet for the year 2020 that attempts to meet the NASA N+2 goals. From our research, Team 1 determined that a longer range aircraft will be preferred and one that has improved destination flexibility will be more marketable. We created a sizing code that calculates geometry, weights, drag and lift characteristics, and stability sizing considerations. From this sizing code we were able to obtain physical dimensions and performance values for our designed aircraft. We were then able to model its cost, and determine whether the aircraft could not only be marketable, but achieve the N+2 goals. Our aircraft ultimately was successful in achieving the range desired of 7100 nmi still air range while maintaining its weight of 87,000lb; a fuel efficiency improvement of 12% meant that our aircraft had a longer range when compared to planes of the same weight class. By also running the G550 through our code, we determined that our code successfully models real aircraft with only a 1-2% error, so in this manner we were able to validate our code. Despite the improvements of our aircraft, we were unable to reach NASA s N+2 goals. While we made great strides, our aircraft was unable to make the noise and fuel improvement requirements. From this point, we will continue refining the code, specifically the drag estimation, and also begin physical modeling and testing to gain more accurate data.

6 MISSION STATEMENT To engineer a conceptual business aircraft solution capable of transporting esteemed passengers, in luxury, while adhering to NASA s N+2 environmental goals. In order to abide by NASA s N+2 environmental standards, our concept will provide reduced NOx emissions, reduced noise pollution, and increased fuel efficiency. These key topics will help address the primary concerns of environmentally conscious groups. DESIGN MISSION To reach our mission statement goals, we decided upon the idea of a long-range business aircraft. By looking at long-range business aircraft currently in production and choosing attributes that we believe we can improve, we created the design mission summarized below Passengers + 4 Crew - Cruise Altitude > 42,000 ft - Cruise Speed 0.85 Mach - Still air Range of 7,100 nmi - Takeoff Field Length 4,700 5,000 ft - Landing Field Length 2,500 3,000 ft By selecting a cruise altitude of greater than 42,000 feet, our business jet will operate above the majority of air traffic allowing for higher speeds and a cruise-climb method; the aircraft increases in altitude as the aircraft becomes lighter from burning fuel. This method improves the overall efficiency of the aircraft and decreases fuel usage. Timely flights are a desirable characteristic that consumers expect in a business jet. High cruise speed directly correlates to the flight duration. Therefore, we chose the cruise speed of 0.85 Mach from historical data as it offers a high speed while maintaining fuel efficiency. This was later verified to be the best jet range speed. A range of 7,100 nmi, a conservative distance equal to traveling from Los Angeles to Hong Kong with a 60 knot headwind, was the design mission range for the aircraft. Destination flexibility is also important for a desirable business jet. With a takeoff field length of 4,700 5,200 feet and a landing field length of 2,500 3,000 feet, the jet will have access to many small

7 airports. This reduces the aircraft design s reliance on larger and more congested terminals which in turn improves turnaround time and decreases wait times. TYPICAL OPERATING MISSION It is not reasonable to expect the designed aircraft to operate at the full design mission at all times. Therefore, the typical operating mission was chosen to carry 6 8 passengers, with 3 crew, over approximately 2,500 nmi. This mission allows for travel between many transcontinental cities. As a reference, a flight from New York to Los Angeles is 2,139 nmi. While this mission does not fully utilize the aircraft s capabilities, the short takeoff and landing capacity will allow more opportunities for shorter range flights in a given time frame. MAJOR DESIGN REQUIREMENTS From the definition of the target customer, we categorized the assumptions concerning desired characteristics for the business jet by performance, aesthetics, service, and extraneous attributes. After evaluating these characteristics, range, speed, and comfort are of highest priority, followed shortly by destination flexibility and environmental image. While range, speed, and comfort are regularly desired qualities for a transpacific personal business jet, destination flexibility and environmental image are developing concerns. In countries and locations where smaller airports are available, the ability for a business jet to avoid landing at busier airports gives the customer the ability to make deadlines without worry of external factors preventing prompt landings. To address these factors, we selected a variety of performance characteristics. The house of quality is a graphic tool which defines a relationship between customer desires and the business jet s capabilities. From its results, we prioritized range velocity and fuel weight. Comparing the designed jet to the current competitors, the range, price and speed are comparable, as seen in the House of Quality that is located in Appendix A. However, it is forecasted that a desire for technology which is less harmful to the environment will progress in the future. Using this House of Quality, we determined that the following properties were most important for meeting our design mission goals:

8 - Range - Empty Weight - Cruise Speed - Cabin Height - Cabin Volume - Take-off Distance SELECTED BEST AIRCRAFT CONCEPT In order to obtain the various parameters of our concept plane, a sizing code was compiled. The sizing code produced dimensions for the nose, fuselage, tail, vertical tail, canard, wing, engine nacelle and engine pylon. These dimensions included values for the lengths of each part, the chord lengths for the wings, diameters of the fuselage and engine nacelle and location of each part in relation to the nose of the plane. Because this was an iterative process, the modeling was done such that dimensions could be easily changed to reflect the approximate current dimensions of the aircraft. After the optimal design was run through the code, the parameters were then used to create a CAD model of the plane in CATIA. WALK AROUND CHART Figure 1, Figure 2, Figure 3, and Figure 4 are the finalized, dimensioned three-view CAD model of the CCD using the current iteration of the sizing code.

9 Figure 1: Top View of CCD Figure 2: Front View of CCD Figure 3: Left Side View of CCD

10 Figure 4: Isometric View of CCD In addition to the outer shell of the aircraft, the interior cabin was also modeled using a 16 passenger configuration designed at the beginning of the project. As an approximate model used for visual inspection of the aircraft, only the basic dimensions were modeled. Pieces such as chairs and tables were modeled as simple rectangles attached to cylindrical legs while the cabinets, galley, and restrooms were modeled as rectangular prisms to represent the space they would take up in the actual plane. Windows and an emergency exit are not currently modeled on the aircraft, but are expected to be on the final aircraft. The emergency exit would be located near the main wing of the jet in order to allow for the best possible evacuation conditions in case the main door is obstructed. Occupants would exit out the aft of the cabin space and then out onto one of the main wings. Figure 5 and Figure 6 are both a view of the cabin inside the CAD model along with the cabin floor plan.

11 Figure 5: CAD Representation of Cabin Figure 6: Top Down View of Cabin The dimensions for the interior pieces are listed in Table 1.

12 IMPORTANT DESIGN FEATURES Table 1: Interior Cabin Dimensions Chair 28 x 28 seat 28 back height Top of seat 21 off of ground Table 32 x 28 top Top of table 32 off of ground Storage 42 x 36 x 48 Galley 72 x 18 base 72 x 20 top Top of galley 36 off of ground Restroom 60 x 36 floor area Surrounded by walls The aircraft s canards allow for the reduced takeoff and landing field lengths which help to meet the destination flexibility we desire. In addition, we have met our range requirements without increasing our weight more than current aircraft of similar proportions. By applying composite materials, we further reduced our weight, facing tradeoffs with maintenance. Our design also allowed for the wings to have a simpler structural mounting, with the wing running straight through the fuselage. Some of the advance technologies, such as geared turbofans, and the composite materials mentioned above, are also included in our design, and will be discussed in more detail further in this document. Table 2: Major Design Parameters Wing Parameters Wing Loading 117 lb/ft² Aspect Ratio 8 Sweep 40 Thickness to Chord 0.1 Taper Ratio ¼ Canard Parameters Wing Loading 80 lb/ft² Aspect Ratio 8 Sweep 40 Taper Ratio ¼ Thickness to Chord 0.1 Aircraft Parameters Thrust to Weight 0.35

13 RESULTS OF AIRCRAFT SIZING DESCRIPTION OF SIZING CODE CCD The final version of the sizing code consisted of several MATLAB scripts broken up into two major elements, those used to predict the empty weight of the aircraft and those used to predict the fuel weight required by the aircraft. The process was iterative, as an initial guess was required in order to size the aircraft components, such as wing area. The fuel burn portion separated the aircraft s mission into many different sections and computed the fuel burned for each segment. The following outlines each of these segments and the approach used to model the fuel burn: Figure 7: Mission segments TAXI AND TAKEOFF Taxi was taken to be 10 minutes at idle at the airport altitude; business aircraft are assumed to have shorter times for taxi because they operate out of smaller airports. Takeoff was calculated as 100% throttle for one minute. CLIMB FUEL BURN APPROXIMATION The entire climb segment was modeled as 100% throttle until the cruise altitude was reached, as business aircraft attempt to reach altitude quickly. The algorithm for determining the fuel burn is as follows: i) Climb was broken up into segments of 500 feet ii) Under 10,000 ft, we assume that air traffic control would limit the aircraft to 250 knots indicated. The calculation for the actual aircraft velocity is as follows:

14 (1) (2) Over 10,000 ft, the following function was maximized for the best velocity for climb: (3) Where V v = vertical velocity V = horizontal velocity T = Thrust available D = Drag W = current aircraft weight The value was approximated as 0.567M² below 36,150 ft and 0.7M² over 36,150 feet. The code uses the MATLAB function fminbnd to maximize the velocity for each climb segment; the bounds set were 250 knots and Mach 0.92 which is the drag divergent Mach number. iii) With the aircraft s horizontal velocity and altitude, drag was calculated iv) With drag and vertical velocity, fuel burn was calculated v) Increment the aircraft s velocity by the delta h, adjust aircraft weight, and repeat until the cruise altitude is reached. There was an issue with the drag function under-predicting the amount of drag required during climb; therefore, Mach 0.92 was found to be the best climb velocity for each step, as seen in Figure 8.

15 Figure 8: Best climb velocity via fminbnd Therefore, a climb rate was hard coded as plus 0.01 Mach per segment, which was essentially a guess. The overall results were considered acceptable for the level of fidelity of the entire code. Figure 9 shows the climb trajectory of the aircraft. Figure 9: Climb trajectory

16 A climb time of approximately 19 minutes seemed reasonable compared to historical data, shown in Figure 10 and Figure 11. The taper off at the end corresponded with the expectations of increased climb difficulty as air density drops with altitude. Figure 10: Actual flight velocity v. Time during climb Figure 11: Actual aircraft weight v. Time during climb

17 The overall climb segment fuel burn behavior over time is monotonic. The aforementioned values and plots are only for the initial climb segment, as it is in the design mission. This analysis was replicated for the reserve mission. CRUISE FUEL BURN APPROXIMATION The cruise segment was broken up into <x> number of distance steps of about <y> distance. For each of these segments, the aircraft weight and altitude were held constant and the fuel burn rate was simply: (4) Where thrust required was the total amount of drag produced, specific fuel consumption was outputted from the engine model. From there, the fuel weight was computed as: (5) At the end of this calculation, the fuel burned would be subtracted from the aircraft weight and then the next step would be computed. The initial choice for number of segments for the cruise leg was 20, in order to assess the validity of this choice, Figure 12 was created. Figure 12: TOGW % change v. number of steps wrt 20

18 The x-axis represents the number of segments used for the cruise fuel burn calculation; the y-axis shows the percent change from the original choice of 20 steps. From the plot, doubling the number of steps only results in just over two tenths of a percent refinement in the solution. Figure 13: Percent Change in TOGW v. the Number of Steps Figure 13 shows the percent change in gross takeoff weight of the aircraft from the previous number. For example, the change in TOGW for 30 cruise steps is about a 0.01% change from 29 steps. Another metric to consider would be the run time for the code, however the time differences were negligible, therefore the original choice of 20 steps was kept. LANDING AND MISSED APPROACH APPROXIMATION These two segments were treated simply as weight fractions: RESERVE APPROXIMATION Landing W 4 /W Missed approach W 5 /W The computations were exactly the same as the design mission, however the Breguet range equation was used for the loiter segment:

19 (6) (7) The loiter Mach and altitudes were chosen to be where the engine deck was most efficient, which were found to be at Mach 0.6 at an altitude of about 37,000 feet. Table 3 tabulates each individual segment result for the final aircraft configuration. Table 3: Final Fuel Weight Results Takeoff and Taxi 400 lb Main Climb Segment 1500 lb Main Cruise Segment lb Reserve Climb Segment 700 lb Reserve Cruise Segment 800 lb Loiter Segment 1200 lb CONSTRAINT CALCULATIONS Several aircraft performance parameters were calculated in order to facilitate the design optimization process later. From the design requirements, destination flexibility would depend heavily on takeoff and landing distances. The following equations were used to calculate these: Landing Ground Roll: (8) Where W = Aircraft weight before landing (lb) Sw = Wetted area (ft²) T = Thrust (lb) =, thrust of thrust reversers v. normal thrust (0.4 from course notes) = aircraft weight fraction, current weight over TOGW = braking coefficient (0.4 from Raymer) = density at airport altitude Takeoff Ground Roll: (9)

20 Where = lapse rate A Federal Aviation Regulation mandates a climb gradient of 2.4% with two engines. Climb Gradient: (10) Where N = number of engines = Zero lift drag = Increase in zero lift drag due to high lift devices, chosen to be AR = Aspect ratio e = Oswald s efficiency factor (0.75) The FARs also mandates a top of climb rate of 100 ft/min. This value calculated in the aforementioned climb function, therefore the very last value at altitude was used for this constraint. DESCRIPTION OF SIZING CODE BWB The BWB design used similar sizing code architecture as that of the canard design. The primary difference was the replacement of the fuselage with a lifting centerbody. The cabin dimensions constrained the size of the centerbody. The centerbody required a thicker airfoil to accommodate the required headroom in the cabin. While the length and width of the cabin could be flexible, the cabin height could not be less than 6 feet without violating the set threshold for the jet. The cabin layout determined previously roughly set the length while the headroom requirement set the thickness. The centerbody airfoil thickness to chord was bound by these values. The design team iterated through various values for thickness to chord ratio and center body length and compiled a comparison of the resulting gross weight estimate to the lift provided. Because the span was already small, center body width was used as a free variable to adjust the lift generation. From this analysis, a thickness to chord ratio of 12% was chosen, along with a center body length of 70 ft. and a width of 20 ft. These dimensions allowed a

21 minimum of excess lift, and lowered the empty weight of the BWB design by 7.5%. The BWB sizing code sized all other portions of the aircraft similarly to the canard sizing code. MODELING APPROACHES AND ASSUMPTIONS In order to begin sizing the aircraft, a number of physical parameters were initially assumed before sizing the aircraft. By sizing based on the G550, we chose our wing s aspect ratio, sweep, thickness to chord and taper ratio. We assumed a lifting efficiency of 90% compared to the 2 dimensional case. We also assumed the aerodynamic center of the fuselage and wing combined was the same as the aerodynamic center of the wing. For a conservative approximation, we also chose the lower values of efficiency improvement provided by outside sources. Many of these approximate improvements will be discussed later in respective sections. SIZING CODE VALIDATION We initially based our aircraft geometric properties on the G550. In order to validate, we ran our sizing code for the physical dimensions of the G550 to see if our code accurately predicted the weight values of the aircraft. G650 Our Code Error Empty Weight 54,000 lb 52,000 lb 3.7% Fuel Weight 44,200 lb 45,000 lb 1.2% Total Weight 99,600 lb 100,000 lb 0.4% Market Cost $58.5 million $62.7 million 7.2% As seen, our code relatively accurately predicted the weight values of the It should be noted that during our presentation to Gulfstream, they informed us that our cost listed based on the information on their website was actually low, so our code s prediction is actually more accurate than what is listed.

22 MAJOR DESIGN TRADE-OFFS CARPET PLOTS In order to determine the design conditions of the best concept, carpet plots were utilized. The objective for the carpet plot analysis was to determine the configuration which produced the lowest cost aircraft. The main constraints on the aircraft, obtained from the constraint diagram, are top of climb, second segment climb, and landing ground roll. The variables we chose for consideration are thrust-to-weight and wing loading. Values for the requirements are from the House of Quality and the Compliance Matrix which are shown in Appendix A. The first requirement value was a ground roll takeoff distance less than or equal to 3600 ft; this value was taken from the target value of the compliance matrix. The second requirement under consideration was the climb gradient: which should be greater than or equal to The last requirement was the rate of climb at altitude; the Federal Aviation Requirement states this value should be greater than 100 ft per minute. The sizing code was run to include various combinations of thrust-to-weight and wing loading, and we recorded the cost, takeoff distance, and range. From these combinations, we created the carpet plot relating the wing loading, thrust-to-weight, and production cost as seen in Figure 14.

23 Production Cost [Millions $] Cost Trends Wing Loading [lb/ft^2] Power (T/W = 0.28) Power (T/W = 0.3) Power (T/W = 0.32) Power (T/W = 0.34) Power (T/W = 0.36) Power (T/W = 0.38) Power (T/W = 0.4) Figure 14: Cost trends for carpet plot. Next, we created the constraint cross plots by requiring the location on the carpet plot to meet each requirement. We then found the value of wing loading and production cost for each requirement. Using these values, we created a final carpet plot, which includes the constraint curves which can be used to identify the best design parameters. The final carpet plot can be seen in Figure 15. The dark line with a negative slope is the climb rate at altitude constraint; all possible values of wing loading must be higher than this line. The line with a positive slope is the climb gradient constraint; all values of wing loading must be lower than this line. From this chart, the bottom right corner provides the best thrust to weight and wing loading values for this aircraft design. The blue line with the negative slope is the constraint of the takeoff ground roll; any wing loading must be higher than this curve to be able to satisfy the constraint.

24 Production Cost [Millions $] Final Carpet Plot Wing Loading [lb/ft^2] Power (T/W = 0.28) Power (T/W = 0.3) Power (T/W = 0.32) Power (T/W = 0.34) Power (T/W = 0.36) Power (T/W = 0.38) Power (T/W = 0.4) Linear (Climb Gradient Constraint) Linear (Vvtoc Constraint) Linear (d TO Constraint) Figure 15: Final carpet plot to find design parameters. The final carpet plot was used to find the best thrust-to-weight and wing loading to optimize the aircraft design. For these constraints, the best values can be seen in Table 4. Table 4: Design parameters found from carpet plot. Parameter Value Thrust-to-Weight (T/W) 0.35 Wing Loading (W/S) 117 lb/ft² In future analysis of the aircraft, the values for the design parameters found in the carpet plots would be used as the design choices. The next step would include finding the optimal value of aspect ratio, or any other major design variables, using the thrust-to-weight and wing loading found through this analysis. Since these are the best values, they would be the starting point for further optimizations. CANARD PLACEMENT TRADE STUDY In order to determine the optimal placement for the canard, trade studies were conducted. Placing the canard lower on the fuselage mitigated some of the interference between the canard and the main wing. Furthermore, a higher canard placement could hinder

25 the ability of the aircraft to pull up to a terminal if need be by obstructing the jet way. This leaves only the canard distance to be determined in a trade study. This trade study was conducted utilizing the strip method code because the primary value affected by the canard s axial placement should be the induced drag. The study set a location for the canard and then recorded the induced drag for that location. Each iteration of the code recorded the canard distance from the nose and the induced drag caused in flight. The tabulated results are plotted below C Di as Location of the Canard Changes C di Canard Distance from Nose [ft] Figure 16: Effect of canard placement on induced drag The study ran at conditions just after takeoff and gear up. From the study, the canard should be placed as far forward as possible to reduce the induced drag. However, there are other factors to consider such as pilot visibility and structural concerns. If the canard were to be placed at the very front it could obstruct the pilot s view of the runway. It may also be difficult mount the canard to that part of the structure since the nose cone is only lightly reinforced. Because of these concerns, the optimal placement for the canard is such that the root leading edge is fifteen feet from the nose. This places the canard almost under the pilot s seat providing good visibility of the ground and reasonable aerodynamic performance. There

26 may be concerns about damage to the canard caused by collisions but if care is taken moving the aircraft it should not be problematic. AERODYNAMICS AIRFOIL SELECTION FOR WINGS AND TAILS Because our business jet will be operating at transonic conditions, a supercritical airfoil was selected. Supercritical airfoils delay the onset of wave drag at transonic speeds, thus are desirable for our cruise speed. The final airfoil choice was made such that the aircraft could take-off, and minimize its empty weight. Based on our sizing code, an increase in the airfoil s thickness ratio corresponded to an increase in total weight. The change in airfoil performance characteristics also affects the aircraft s lift characteristics. From the code predictions, the increase in lift compensates for the increased weight, allowing for the aircraft to still take-off despite the weight increase. We selected the airfoil that provided a low empty weight and still had the lift abilities for take-off. Thus, we selected the NASA SC(2)-0612 for the wing and the NASA SC(2)-0610 for the canard. It is to be noted that while we chose this airfoil s thickness ratio and it s coefficient of lift for an angle of attack at zero, our aircraft s airfoil will not exactly be the NASA We would be using a similar airfoil tailored to our needs. HIGH LIFT DEVICES While modeling take off performance we determined that high lift devices were not necessary for our configurations. The aircraft designs both could take off at a moderate angle of attack at approximately 150 knots. However, we decided to implement them on the design to help reduce our take off distance and on landing. High lift devices add weight to the aircraft but we felt that the additional lift outweighed the penalty to weight. We decided to only implement a simple flap to decrease the weight penalty and only use them on the main wing. A simple flap system is estimated to add approximately 6 lbs of weight per square foot of area. The expected increase in the lift coefficient is approximately 0.5 with fully deflected flaps on approach.

27 DRAG BUILDUP We modeled the aerodynamics of both designs using a set of functions which predict the different contributions of drag force and combine them together. The sizing code passes current flight conditions and aircraft geometry in to a function called Total_Drag.m which then calls a parasite drag function. The parasite drag function uses a component build up method to predict the viscous and pressure drag. The list of major components is different for the two designs. The components included for the canard design are the fuselage, the main wing, the canard, the nacelles, the pylons and the vertical tail. The components included for the blended-wing design are the same with the following exceptions: the canard and vertical tail components are not used and a center body is used in place of the fuselage. The parasite drag function returns a value to Total_Drag. Total_Drag then calculates drag due to lift using the following equation. (11) Total_Drag.m then calculates the approximate drag divergence Mach number using Shevell s curves supplied by Dr. Crossley. If the flight speed exceeds then Total_Drag uses Locke's fourth power law to compute an estimate of the compressibility drag. The function then sums all of the drag coefficients together and then calculates the total drag force applied to the aircraft using the definition of the drag coefficient. (12) The total drag applied to the aircraft is then returned to the main sizing function in pounds of force. WEISSINGER STRIP METHOD For a prediction of the induced drag on the aircraft, a Weissinger strip method code, weissinger.m, was written. However, this method relies on potential and incompressible flow assumptions which do not hold at transonic speeds. Thus, the strip method code could not be implemented in the main sizing code. Fortunately, the design team still made use of it for modeling takeoff performance which was not specifically addressed in the main sizing code.

28 The strip method uses a collection of horseshoe vortex strips to approximate the lift distributions on the lifting surfaces. The function then predicts the induced drag using the Biot-Savart Law. It also calculates the actual lift generated using airfoil data. This is done by calculating lift per span on each strip and then multiplying by the strip width. These pieces are then summed together to find the Lift force in pounds. The lift prediction was used to determine whether or not high lift devices are necessary and at what speed takeoff could be accomplished. PERFORMANCE V-N LOADS DIAGRAM In order to determine the view of normal load factor versus speed possible for the aircraft and to see how this relationship altered when the aircraft encounters a gust, a V-n (loads) diagram was created. In Figure 17, the equivalent velocity of the aircraft is plotted versus the positive and negative load factors which were used to find the empty weight of the aircraft. The V-n diagram shows maneuver loads and gust loads on the aircraft, but neglects the load on lifting surfaces and gust loads due to control surface deflection. Several assumptions were made during the construction of the diagram; we assume the Mach of the design dive speed equals the Mach number of the structural design cruising speed plus 0.2 Mach, the gusts encountered have a cosine intensity which acts instantaneously and affects the entire aircraft instantaneously, the gust intensity varies with Mach and altitude, and it is possible to interpolate the FAR gust intensity values given for 20,000 ft and 50,000 ft to find the gust intensity at the cruise altitude. The maneuver loads diagram was constructed to represent the equivalent speeds the aircraft will fly versus the expected load factors in flight. This can be seen in Figure 17. This graph used the maximum lift line from zero velocity to design maneuver speed (for a positive load factor). Between the design maneuver speed and design dive speed, the positive load factor is constant. At the design dive speed, the load factor decreases from the given value until it gets to zero. The negative load factors are incorporated into the V-n diagram using the

29 Load Factor (n) curve of maximum coefficient of lift for negative angle of attack (for a negative load factor). From the speed at which the load factor reaches the design value, it remains constant until the speed arrives at the structural design cruise speed. Between the structural design cruise speed and the design dive speed, the negative load factor increases linearly. 2.5 Maneuver Loads Diagram Equivalent Velocity (knots) Figure 17: Maneuver loads diagram. Next, the gust diagram was created to include the effect on the aircraft from turbulence or storms. The gusts encountered by the aircraft change the apparent velocity and increase the angle of attack of the load factor. We calculated the gust speeds using the data displayed in Table 5 by interpolating to find the gusts at each speed at the flight altitude. Table 5: Gust speed assumptions used for calculating gust at altitude. h<=20,000 ft h>=50,000 ft Location: Gust speed (ft/sec): Location: Gust speed (ft/sec): The incremental load factor from the gust can be calculated using Eqn. 13.

30 Load Factor (n) (13) Where is the gust speed in feet per second, is the intensity factor which accounts for cosine shape of the gust, is the equivalent velocity at that point in knots, is the coefficient of lift due to angle of attack, and is the wing loading. These incremental load factors were added at the specified speeds, to alter the graph to match conditions which are more likely to occur during flight. The gust diagram, plotted with the maneuver diagram, can be seen in Figure Maneuver and Gust Loads Diagram Equivalent Velocity (knots) Figure 18: Maneuver and gust loads diagram. The combined V-n diagram is found by combining the outermost points of the maneuver and gust diagrams from the previous two figures. This diagram can be seen in Figure 19.

31 Load Factor (n) 3 Combined Loads Diagram Equivalent Velocity (knots) Figure 19: Combined loads diagram. The values which were used to create the V-n diagram can be seen in Table 6. Table 6: Assumptions for creating loads diagram. Input: Value: Units: W/S 80 lb/ft^ /rad Cruise Altitude ft Positive Design Maneuver Load Factor 2 N/A Negative Design Maneuver Load Factor -1.4 N/A (equivalent) 150 ft/sec (equivalent) ft/sec (equivalent) ft/sec (equivalent) ft/sec In the future development of the aircraft, the V-n diagram would need to be updated to better represent the gust loads which the aircraft feels. More detail should be added into the gust analysis to include a non-cosine gust function, and the gust should be included in the atmosphere function instead of adding it in separately. A study was completed to determine the effect of changing the load factor which the airplane could withstand. If the positive load factor is changed from 2 to 2.5, the upper part of

32 Load Factor (n) the V-n diagram shifts upward to reflect this change. This is illustrated in Figure 20. The red line shows the outline of the V-n diagram when the maximum positive load factor is 2, and the blue line shows the adjusted V-n diagram with the higher load factor. As seen in Figure 20, the slowest speed the aircraft can fly without stalling at the maximum load factor is higher when the maximum load factor is higher. 4 Combined Loads Diagram Equivalent Velocity (knots) Figure 20: Loads diagram with change in maximum load factor. Another study was completed to determine the effect of changing the wing loading of the aircraft. If the wing loading is changed from 80 lb/ft^2 to 70 lb/ft^2, the V-n diagram shifts outward to reflect this change. This is illustrated in Figure 21. The red line shows the V-n diagram when the wing loading is 80 lb/ft^2, and the blue line shows the adjusted V-n diagram with the lower wing loading. As seen in the figure, the lower wing loading will decrease the stall speed at the maximum load factor, but will increase the load factor which can be withstood above 225 knots equivalent airspeed.

33 Figure 21: Comparison of two wing loadings. The complete loads diagram for a wing loading of 70 lb/ft^2 can be seen in Figure 22. Figure 22: Loads diagram for lower wing loading. Additionally, a study was completed to determine the effect of increasing the wing loading from 80 lb/ft^2 to 90 lb/ft^2. This change is illustrated in Figure 23. The red line shows

34 Load Factor (n) the V-n diagram when the wing loading is 80 lb/ft^2, and the blue line shows the adjusted V-n diagram with the higher wing loading. As seen in the figure, the higher wing loading will decrease the stall speed at the maximum load factor, but will slightly increase the load factor which can be withstood above 225 knots equivalent airspeed. Figure 23: Comparison of two wing loadings. The complete loads diagram for a wing loading of 90 lb/ft^2 can be seen in Figure 24. Combined Loads Diagram Equivalent Velocity (knots) Figure 24: Loads diagram for higher wing loading. Although the V-n diagrams shown in this paper do not use the final value for wing loading of 117 lb/ft², the trends displayed above are correct. From the trends analyzed, it is

35 expected that the higher value of wing loading will decrease the stall speed at maximum load factor and will increase the load factor which can be withstood at high equivalent velocities, both compared to the values seen in Figure 24. V-n diagram relates to other portions of the sizing code by being able to use this data and find what loads the aircraft can take at each equivalent airspeed. This is important because if the aircraft flies at a loading higher than the ones shown in these diagrams, the aircraft may fail. PERFORMANCE SUMMARY After running through our sizing code, the final performance values for our aircraft were found and are shown in Table 7. Table 7: Performance Summary Values Best Range Velocity 823 ft/sec Best Endurance Velocity 581 ft/sec Stall Speed 241 ft/sec Maximum Speed during Climb M DD Maximum Speed during Cruise M = 0.9 Takeoff Distance (ground roll) 3,600 ft Landing Distance (ground roll) 600 ft PROPULSION ENGINE MODELING The sizing approach involved using a rubber engine in order to minimize the number of constraints during the preliminary design stage. However, due to insufficient information, an engine deck could not be developed to the desired degree of accuracy. As an alternative, a set of supplied engine data was used to determine fuel burn characteristics of the aircraft. The data was for a high bypass turbofan similar to General Electric s CF34-8 engine, with a sea level static thrust capability of 13,600 lbs. The tables contained various values of thrust and fuel flow rates with their corresponding Mach and altitudes. Curve fits were assembled for each of these in order to help provide more fidelity for the calculations. Fourth order polynomials were used on the suggestion of the instructor, and thrust was used as the independent variable and fuel

36 flow rate for the dependent variable. SFC was not used for the fits because it was found that SFC values were not consistently monotonic, whereas an increase in fuel flow rate would result in an increase in thrust. Thus we decided that the oscillations in SFC made it more difficult to curve fit. The engine model was separated into three sections, 100% throttle, idle, and partial throttles. With the first two, the flight conditions are passed and the corresponding thrust is calculated. The partial throttle section of the model takes a given thrust requirement, either from the drag function or the other two throttle settings, the aircraft flight conditions and returns the fuel flow rate in pounds per hour. In accordance to the rubber engine design philosophy, the thrust and fuel flow rates would have to be scaled with the size of the aircraft. This was done in the following manner: (14) (15) Where T max = SLS thrust requirement for the aircraft N en = number of engines on the aircraft SF prop = scale factor for the propulsion subsystem With a static set of engine data values, the scale factor was used for changing the thrust required from the engines in order to ensure that the fuel flow rate values would be representative of a turbofan of this size. The scale factor used in the final sizing iteration was The supplied data was assumed to be outputs from an engine deck and for an uninstalled engine. A factor of 3% was used in order to take installation losses and power extraction into account. ENGINE TRADE-OFFS Pratt and Whitney s geared turbofan, the PurePower series, aims at improving propulsive efficiency by gearing the main fan to rotate slower. Current turbofans link their fans and low spool compressors to the same low pressure turbine, which forces each component to

37 operate outside of their optimal regions. The gearing of the fan allows for the reduction of the fan rotating speed and subsequently the fan pressure ratio, reducing the exit flow velocity. The gearing also gives the low spool compressor the freedom to operate at higher rotor speeds, therefore in regions of higher efficiency. The reduction in jet velocity is ideal for propulsive efficiency, but requires an increase in the air mass flow being moved. This results in such bypass ratios of 8:1 and 10-12:1 for these engines (Norris), thus larger engines, which presents a problem with engine placement for smaller aircraft. Under wing placement would be restricted for many of them and larger pylons would be required for standard aft fuselage installation. Pratt and Whitney intends on developing a smaller version of their geared turbofan for Mitsubishi s regional jet. The PW1215G and PW1217G are rated for between 14k to 17k lbs at SLS and are claimed to reduce fuel burn by 12-15% versus current engines. NO x and CO 2 emissions are reduced due to an advanced combustor, with NO x emissions being 50% under CAEP (Committee on Aviation Environmental Protection) 6 standards. The smaller geared turbofans are expected to reduce noise levels 15 decibels below stage four regulations (MTU Aero Engines GmbH). The PW1215G and PW1217G are to have fan diameters of 56 inches, which translates to the engine being slightly larger than other high bypass turbofans in this thrust class. Rolls Royce s BR710 is rated for lbs at SLS and has a fan diameter of 48 inches. Their BR725 at lbs SLS is about 50 inches in diameter (Rolls Royce). The BR710 is currently flying with Gulfstream s G550 and Bombardier s Global Express. The BR725 is expected to enter service in 2012 with Gulfstream s G650. A quick study was conducted in order to identify the effects of the increased diameter. The sizing code ran for engine diameters from three feet wide to seven feet wide. Figure 25 shows the results, where TOGW represents takeoff gross weight.

38 Figure 25: TOGW v. Engine Diameter Overall, a 1% change in engine diameter changed the takeoff gross weight by about 0.17%. The only variable changed was the fan diameter, the nacelle and pylon sizes and weights were all adjusted accordingly by each of the respective functions. Additionally, the tail size would be affected by the longer moment arm created by having to push the engines further away from the centerline. The jumps and discontinuities are unexpected and their cause is unknown. The plot was recreated with a much finer step of 0.01 ft increments in diameter and other odd behavior shows up. This can be seen in Figure 26.

39 Figure 26: TOGW v. Engine Diameter (0.01 diameter increments) A higher bypass ratio, and consequently larger fan, does offer an SFC benefit. Therefore, we conducted another sensitivity study for specific fuel consumption. Figure 27: TOGW v. SFC

40 A 1% change in SFC resulted in an average change of 1.3% in gross takeoff weight with everything else held constant. In this case, we varied the technology factor applied to the fuel burn from 100% to 50%. 100% was chosen because it represents no change from today s turbofans; 50% represented the other extreme where NASA s N+2 goal is met. From these results, the weight and drag penalty from the geared turbofan is made up in its improvement in fuel efficiency. The smaller PurePower models are about 11.25% larger than Rolls Royce s BR710, which would increase the TOGW of aircraft by about 1.6%. Conversely, a 12% fuel burn savings would decrease the TOGW by about 15.6%. We modified this study slightly for a scenario involving General Electric s Unducted Propfan. In this case, however, there should be no added nacelle drag. Since the core of the UDF is not a set design, an engine diameter of four feet was assumed. The larger rotors do require more clearance, from UDF test bed in the 1980 s. The fan is about 140 inches in diameter and therefore a pylon length of seven feet was set. An additional 3% was added to the engine weight, as the propfan is expected to be heavier than the geared turbofan. These values are not exact, but the resulting values, Table 8, should provide some insight. Table 8: Propfan results TOGW (lbs) Fuel Weight (lbs) Empty Weight (lbs) Geared Turbofan Propfan As expected, the fuel consumption savings would outweigh the physical penalties of the propfan, thus a 7.5% decrease in empty weight can be seen. This study did not include the weight of extra insulation required for the cabin, as pointed out by A.B. Bauer. Additionally, the assumption is that far field noise issues have been rectified without significant changes to any physical or performance characteristics. While we chose a design cruise Mach number of 0.85 early in the project, the tools for analyzing this choice were only recently completed. The following plot shows the thrust required, or drag, for various Mach numbers at the cruise altitude of 42,000 feet and with an aircraft weight of 70,000 lbs. The thrust available is also shown and the crossing point of the two was determined to be the Mach for best jet range, seen in Figure 28.

41 Figure 28: Thrust Available v. Thrust Required for cruise Coincidentally, Mach 0.85 was about the ideal cruise Mach for the given configuration. The drops in thrust around Mach 0.6 and 0.9 are due to the deficiencies in supplied engine data, therefore data outside of these zones was ignored in order to maintain accuracy. The thrust required versus thrust available at the approach and takeoff configurations are shown in Table 9. Table 9: Thrust Available v. Thrust Required Drag (lbs) Thrust (lbs) Weight (lbs) 250 knot Takeoff knot Approach It should be noted that our drag predictions are likely low on takeoff largely due to modeling inaccuracies. We believe that although our code accurately models at cruise, the drag performances at low altitudes are perceivably inaccurate.

42 STRUCTURES CONCEPTUAL STRUCTURAL DESIGN Conceptual structural design is an important step during the later stages of the design process. Though many of the particular loadings, aerodynamic interactions, material strengths, and finalized geometry layouts are unknown, structural engineers must postulate the relative sizes and locations of major structures and the loadings they must undergo. This base level structure allows the estimation of the internal volume available for fuel, avionics, and other important systems. Though the conceptual structural design may not completely describe the final solution, it will be close enough that major changes to the structure will not be necessary in later design phases. Many of the conceptual approaches to design the internal aircraft structure are taken from observation of previous business aircraft solutions. Unfortunately though, the particular layout, geometry, and locations of major elements (i.e. wings, tail, etc.) are unique to our concept. Though we attempted to adhere to conventional structural design, deviations were necessary to account for differing aircraft. These changes primarily utilized the observation of load paths and using engineering intuition to design structure to account for these load paths. IMPORTANT LOAD PATHS During service and storage, aircraft undergo a litany of various loadings. During conceptual design and early analysis, it is important to theorize many of these important and high magnitude loads. These assumptions are needed to conceptually design the locations and relative size of the load carrying structures. The two drawings below, shown as Figure 29 and Figure 30, demonstrate several of the main loads present for the internal aircraft. Though the internal structure for the two designs will be drastically different, the general loads experienced during the aircraft s life will be very similar. The BWB has many unique features, such as a fuselage that acts as a lifting surface, which requires exotic structure. As there is not a basis for comparison for this design, a new

43 conceptual approach must be created and will be discussed later. This section will focus on the similar structures between the two designs. Figure 29: CCD Loads

44 Figure 30: BWB Loads Aircraft Weight: While on the ground and in operation, the weight of the aircraft must be accounted for. On the ground, the weight is carried by the landing gear and relevant fuselage structure. While in the air, the weight is accounted for by the lift and is, therefore, carried by the wing structures. General Lifting Load: These loads are created by the lift generated by the airfoil lifting surfaces. In both cases, these loads are generally carried in the wings (or canard) by internal wing (or canard) spars. Torsional Load due to Lift: This load is a torque generated by the chordwise lift distribution over the airfoil and is carried by the closed cell geometry of a wing box. The ribs of the wing maintain the aerodynamic geometry of the airfoil

45 Nose-to-tail Moment: When the aircraft is struck with a gust or the CCD performs a climb maneuver, the aircraft experiences a nose-to-tail bending. This load path is carried by the longerons that run along the length of the aircraft. Landing Gear Loading: The landing gear of an aircraft is a very high strength structure. While landing, the gear, and structure that the gear are attached to, must withstand many times the weight of the aircraft. These loads are taken by the landing gear themselves as well as other high strength structures (such as wing spars). Engine Weight: This load is proportional to the static weight of the engines and is carried by the pylons. The pylons are then attached to high strength spars or stiffeners. Rudder Torque: During flight, rudder deflection applies large torques to the tail (and smaller magnitude torques to the pylons). These loads are carried by tail (or pylon) spars and chordwise ribs. Thrust: The thrust generated by the jet engines creates a pitch down moment due to their location above the centerline of the aircraft. The thrust loads are carried by the pylon spars and structures to which the pylons are connected. Cabin Pressure (not shown): To ensure that customers are comfortable, the cabin must be pressurized. The cabin pressure is carried by the pressure vessel skin and circumferential stiffeners or formers. CONVENTIONAL CANARD DESIGN STRUCTURAL CONCEPT A forward canard and rear wing design is not a new concept. There have been several designs that have utilized this configuration to engineer a business aircraft solution. Our design takes many of the conceptual structure elements with small modifications to accommodate our particular size and geometry layout. Figure 31, Figure 32, and Figure 33 below depicts a front, top, and side view of our conceptual design.

46 Figure 31: CCD front view Figure 32: CCD top view Figure 33: CCD side view

47 FUSELAGE The CCD fuselage design uses conventional wing and tube ideology that utilizes thin walls for the pressure vessel and fuselage skin. Longitudinal stiffness is provided by longerons that run from nose to tail. Circumferential stiffness is provided by stiffeners or formers which will also aid in internal pressure resistance and serve as a mounting point for the longerons. CANARD BOX To ensure that the main wing is not directly in the wake of the forward canard, the canard was placed as a low canard. Figure 34 and Figure 35 depict a close up view of the canard box. Figure 34: CCD Canard side view Figure 35: CCD Canard top view This low canard design offers many benefits. As shown in Figure 36, the main spar of the canard goes through the fuselage similar to a low-wing design.

48 Figure 36: Low Canard Benefits The low-wing allows a very simple wing (or canard) box as well as relatively lightweight construction without the need for extra structure to carry the load around the fuselage. The low canard also does not interfere with the rest of the cabin allowing for maximum cabin volume and comfort for the passengers. WING The wing uses a very conventional structural layout. Figure 37 shows a closer view of the CCD wing structure. Figure 37: Wing Structure

49 The majority of the longitudinal stiffness is provided by the large leading edge (LE) and trailing edge (TE) spars. These offer bending resistance due to lifting loads on the wing. The two spar design maximizes the internal volume of the wing that may be utilized for fuel storage while still maintaining the closed cell torsional resistance required for lifting loads. To prevent buckling of the skin and to maintain the wing s airfoil shape, ribs will be used at intervals along the span. The number and size of ribs, spars, and other wing structure, as well as the control surface size and their wing interaction, will be optimized in later design phases. Another design consideration is the pilot s visibility of the wing tips. Our current design would make it very difficult for a pilot to see the ends of the wings during ground maneuvers. A system of cameras would likely be implemented to alleviate this issue. WING BOX One of the aircraft structures that experience the largest loading is the carry through wing box. This structural member takes all of the loading from the wings and transfers it through the fuselage. To withstand this high loading, the wing box must be very robust. Figure 38 and Figure 39 depict this carry through structure. Figure 38: Carry Through Wing Box Figure 39: Wing box Top View

50 The wings are mounted far aft on the aircraft. This places the main wing spars behind the pressure vessel/cabin and allows the aircraft to have a mid-wing design that takes the main spars directly through the fuselage. Figure 40 demonstrates a mid-wing. Figure 40: Mid-wing depiction The mid-wing with spars taken through the fuselage offers several benefits compared to alternative mid-wing designs due to a reduced need for significant turning of the load path. The turning of a load requires extra structure to support the nonlinear distribution through the member and, therefore, a design without the turning offers a lower weight structure and a more simple wing box. The mid-wing also does not experience dihedral effects which allow the wings to be flat, further reducing the need for load turning. TAIL When the rudder of the aircraft deflects, rotational loads are induced on the tail. During flight, especially during emergency conditions such as one-engine-out or landing in a sideslip, it is important that this stability surface remains relatively rigid. Figure 41 shows the tailfuselage interaction.

51 Figure 41: Tail-Fuselage interaction The LE spar of the tail is taken through the fuselage into the rear spar of the wing. The wing box is a very strong structure and would be an excellent mounting point for the tail. The TE spar is attached to a circumferential stiffener near the rear of the fuselage. The rudder size is drawn in such a way as to show structural interaction and is not optimized. Particular rudder sizing will be addressed later. ENGINE PYLONS Engine pylons are subject to very high magnitude cyclical loads and must be cable of withstanding the reaction forces generated from the engine. The LE spar of the tail is a strong member and is connected to the wing box. The LE of the pylons would be connected to the LE spar of the tail. The TE spar of the pylon would also be connected to the same stiffener as the tail TE spar. These interactions were also shown in Figure 41. LANDING GEAR The landing gear of the aircraft is a very important system and is often overlooked during conceptual analysis. The mid-wing design of this aircraft allows the wheels to be stored within the fuselage and further maximizes the wing volume available for fuel. The size of the landing gear was calculated using the equations in the class textbook. The height of the main landing gear was calculated using the angle of attack at take-off and the

52 distance from the main landing gear to the rear of the fuselage plus the back end clearance required for an aircraft. A diagram explaining the equation used can be seen in Figure 42. H x Figure 42: Landing gear calculation method α From this figure, the value for the height of the main landing gear can be calculated using Equation 12 where x is the distance from the main landing gear to the end of the fuselage plus the back end clearance. (12) The tire dimensions can be calculated from the statistical data found in. The weight carried by each wheel can be found using Equations in the same book. These are then used to calculate the tire wheel diameter and width associated with the weight carried by each wheel. A tricycle landing gear has been chosen to be used to minimize weight while preventing tipping and roll-over while on the ground. The tricycle landing gear is the industry standard for business aircraft and will be used in our concept. There are several assumptions associated with using these equations which include that the main tires carry about 90% of the aircraft landing weight, there are two wheels per strut, and the aircraft will not be landing at soft or rough runways. Using the above method, the gear height was calculated to be about 4.05 ft with a 3 ft clearance on roll-out. This gear height is measured from the bottom of the fuselage and would require more height to reach the mid-wing spar. The forward gear will be mounted to the canard and the rear gear will be mounted to the wing box. Figure 43 shows this interaction.

53 Figure 43: Landing gear interaction Figure 44 shows the gear deployment. Figure 44: Landing gear deployment The front gear will sweep forward from under the fuselage and rotate to the appropriate orientation for landing. The rear gear will sweep forward from under the fuselage and separate to a wheel base preventing roll-over on the ground. The rear gear is placed at about 80 ft from the nose of the aircraft. This location is behind the center of gravity and prevents the aircraft from tipping backwards on the ground. The wheels in the above drawings are not to scale and are larger than calculated to better show storage and other interaction with the airframe. Another concern that came up during our conceptual analysis was if the wing tip would contact the ground during takeoff. Figure 45 shows that the wing tips do not extend very far past the underside of the fuselage.

54 Figure 45: Wing tip location With our calculated 3 ft clearance during roll-out, and the fact that the wings will deflect upward due to the high lift at take-off, the likelihood of the tips dragging is very unlikely. BLENDED WING BODY DESIGN STRUCTURAL CONCEPT The concept of a blended wing body aircraft is a revolutionary paradigm shift in the way engineers design aircraft. Since the technology is still in early development stages, there are very few BWB designs, most of which are also still conceptual. Without historical structural designs to base our BWB concept, the conceptual structure is a hybrid of adapting conventional wing-and-tube designs and applying engineering intuition. Figure 46, Figure 47, and Figure 48 show the conceptual structural design layout of our BWB concept. Figure 46: BWB Front View

55 Figure 47: BWB Top View Figure 48: BWB Side View PRESSURE VESSEL One of the unique problems encountered by the BWB aircraft is the pressure vessel/cabin. Cylindrical pressure vessels are the most efficient in resisting hoop stresses and, thereby, minimizing weight. Due to the wide, airfoil shaped fuselage of the BWB and the shorter overall aircraft length, a cylindrical pressure vessel would dramatically reduce the internal volume of the cabin as well as waste a considerable amount of space.

56 Two solutions have been proposed to address this issue. The first is to make the pressure vessel elliptical. This concept will fill more of the space of the fuselage but, due to the ellipse s ineffective pressure resistance, the vessel must have a significantly thicker skin and, therefore, be heavier. The second solution is to overlap several cylindrical pressure vessels, remove the overlapping skin, and replace that material with a thin membrane or several columns. Research by NASA scientists supports this method (Mukhopadhyay and Sobieszczanski-Sobieski). Figure 49 illustrates this design. Figure 49: BWB skin method The vertical columns would not carry pressure loads but must account for the hoop stress that is created by the pressure attempting to force the vessel into a circle. A conceptual approach to estimating this pressure vessel began by calculating the hoop stress, σ θθ, in each of the cylindrical pressure vessels using Eqn. 13. (13) Where P = internal pressure r = radius of the particular vessel t = thickness of the skin. Knowing the pressure, radius, and choosing a conservative thickness, the hoop stress in each vessel can be calculated. The distance that each center point is from the center vessel will determine the amount of hoop stress each column must withstand. The variables that will be discussed in the further analysis are defined in Figure 50.

57 Figure 50: Hoop Stress Where d = distance from the large vessel to the smaller x = distance from the large center point to the chord R = larger radius r = smaller radius a = length of the chord As d increases, the length of the chord decreases and, therefore, the amount of stress carried in the beams would similarly decrease. The equations for the two circles are shown as Eqn. 14 and 15. (14) (15) Solving for the distance x gives Eqn. 16. (16) The chord, a, is then found using Eqn. 17. (17) Using the value a, the arc angle for the two circles that would be removed by their overlap is found using Eqn. 18 and 19.

58 (18) (19) By multiplying the two angles by the radius of the corresponding circle to find the arc length removed as well as multiplying those arc lengths by the corresponding calculated hoop stress, the stress that must be carried by the membrane or columns between the vessels can be calculated. The weight of the pressure vessel is then calculated by multiplying the weight by the total amount of material used (Weisstein). The below figure, Figure 51, shows an example of the weight calculated using the above method. This example assumes a carbon-fiber composite skin with a 3/8 efficiency factor due to randomly oriented fibers (Callister and Rethwisch). The model also assumes a 7 ft diameter main radius, 6 ft smaller radii, 25 ft cabin length, a skin thickness of 0.05 in, and a cabin pressure of psi (6,500 ft).

59 height of column [ft] volume of pressure vessel [ft 3 ] weight of pressure vessel [lb] d [ft] d [ft] d [ft] Figure 51: BWB sample skin weight As expected, as the length d increases (and therefore the amount of skin that is still present in the design), the weight of the vessel increases. Using the maximum available cabin volume, the distance between center points would be around 5.25 ft. However, this would make fitting the vessel into the contour of the fuselage difficult and would also have around a 4 ft ceiling between the middle and side compartments. An optimization against weight, cabin volume, and column height would be performed in later stages to ensure the best design is chosen. FUSELAGE It is difficult to distinguish between fuselage and wing for a BWB concept due to the lifting characteristics of the body. For our concept, the fuselage is defined as the non-wing structure surrounding the pressure vessel and the structure aft of that. Figure 52 shows the structure considered the fuselage.

60 Figure 52: BWB Fuselage Other than the pressure vessel, the fuselage maintains a more conventional design. To help resist the pressure and maintain the airfoil shape of the fuselage, circumferential stringers/formers would be placed along the vessel. Due to the lift loads on the body, nose-totail spars (shown in the lower part of the above figure) would be integrated into the structure to improve the longitudinal stiffness. WING The blended wing-body interaction allows a very large amount of space in the wing for fuel and other components such as landing gear. The two spars would give the wing its required strength similar to the CCD. Figure 53 shows the wing structure.

61 Figure 53: BWB wing WING BOX Just like the CCD aircraft, the wing box is a very important structural element. Due to the complex geometry of the pressure vessel and high loads generate by the wings, we decided it was important to minimize the wing loading on the vessel. To accomplish this, many of the loads are carried around the vessel instead of through it. Figure 54 shows the wing-fuselage interaction. Figure 54: BWB Wing box As shown above, the TE spar is taken through the fuselage as a mid-wing similar to the interaction discussed in the CCD. The LE spar would likely be taken through the cabin; therefore, the carry-through mid-wing box is bent around the pressure vessel. Figure 55 shows this interaction.

62 Figure 55: BWB Wrap around Wing Box Unfortunately, to carry the load around the vessel, the load must be significantly turned. This creates a much heavier carry-through section but maintains many of the other mid-wing advantages. ENGINE PYLONS Just like the CCD, the pylons must sustain the loading caused by the engines. Figure 56 shows the pylons. Figure 56: BWB pylons In this design, the LE spars of the pylons are attached to the TE wing spar. This is a very strong structure and a good mount for the engines. The TE pylon spars would be attached to a span-wise spar that would be through the fuselage. LANDING GEAR The approach to landing gear sizing was the same as the CCD. The tricycle pattern offers a lower weight and prevents tipping and rolling on the ground. The leading gear would be attached to a stronger stiffener and the rear gear would be attached to the TE wing spar.

63 The mechanism for gear deployment would be similar to many conventional aircraft. The front gear will deploy down from the nose. The rear gear, which is stored in the wings, will deploy down to an angle that prevents roll over on the ground. MATERIALS SELECTION AND JUSTIFICATION Composites have been chosen to be the material for constructing the aircraft. There were many details which went into this decision; there are benefits and costs from utilizing this material choice. Benefits include a high strength to weight ratio, strong corrosion and fatigue resistance, lower part counts within the aircraft, and the ability to be aeroelastically tailored; this is especially important for the Blended Wing Body. The strength to weight ratio is important because it what governs how the aircraft responds to loading during flight. Corrosion and fatigue resistance are important to the aircraft having a safe, long life. Using composites can reduce the number of parts which must be connected when manufacturing the aircraft. This lower part count will help reduce the amount of labor costs from tooling parts together. Aeroelastic tailoring is especially important to keep the interaction points as smooth as possible. This will help prevent turbulent flow over the aircraft. There are also some negatives associated with composites. These include manufacturing details, additional cost to produce, and of the composite material specifically: inspection, repair, mass production capability, and the lack of environmental friendliness. The cost to produce the composite aircraft will be higher in the initial stages of the production run since the composite molds must be made. Since the Boeing 787 is already in production a commercial aircraft largely made from composite materials we are operating using the assumption that there will be strategies to inspect and repair the airframe. By comparison, if aluminum were used to produce the aircraft, the time costs of learning how to work with the material would not be as great as expected with composites. The process for designing aluminum aircraft has been used for several decades, so the process has already been optimized. One of the initial design criteria for designing an environmentally friendly aircraft would be to have a certain percentage of recyclable materials; however, this value will likely not be achieved through use of composite materials.

64 Taking all of the benefits and costs for using composite materials, it was decided to use composites for this aircraft. Similarly to the Boeing 787, the main structure components are forecasted to be made of composites. The expectation is to use approximately 50% advanced composites as seen on the aircraft. The breakdown of materials used for construction will be similar to that seen in Figure 57 with other structural elements following current industry standards. Figure 57: Material Selection (Hale) Though composites have seen many uses in the aerospace industry, its many problems have delayed its full acceptance. The high cost of manufacturing and material costs are one of the major setbacks for this excellent material. Unlike aluminum and other conventional materials, composites do not have the same robust infrastructure to accommodate manufacturing on a very large scale. Though this is an issue in the current market, we believe that the large amount of composite structure that the Boeing 787 will require for production will bolster the industry s knowledge and infrastructure on composite materials. Regarding the problems of inspection and repair, since the aerospace industry has begun the transition to high component volumes of composite materials, the industry has realized its need for new methods. The nondestructive evaluation and repair of composite structures is a very prolific research and development topic at the moment. We believe that, due to the research and infrastructure improvements involving composite materials, that by 2020 enough of the

65 problems currently facing composites will be addressed to a point that it will be a very viable structural material for our concept. WEIGHTS AND BALANCE The maximum range mission, where the aircraft was sized, assumes a payload of 10 passengers with 4 crew members. Each person was taken to have a weight of 225 lbs., which included some carry-on items. Along with overestimating for an average weighted person, an additional 5000 lbs. was also assumed to come onto the aircraft as removable payloads e.g., fax machines, computers, other office supplies and foodstuffs. The empty weight of the aircraft was determined by taking a total sum of component weights. The component weights calculated was through empirically derived equations found in Daniel Raymer s text book. Component weights are a function of time and thus required a gross weight to calculate an empty weight estimate. The following diagram helps portray the process. 4) Repeat until consecutive gross weights meet convergence criteria 1) Calculate Fuel Weight from Mission 3) Sum Empty, Fuel, and Payload weights to estimate Gross Weight 2) Pass in gross weight estimate to get Empty Weight Figure 58 - Weight Calculation Flow Chart Using a single source for component weights may have led to a lower than expected total gross weight of both aircrafts. This lower weight was found without the use of a composite material factor. Thus, the empty weight equations are likely the reason for an early convergence in gross weight. The empty weight calculations determined the future gross weight for the coding process, thus, a lower empty weight leads to a lower gross weight. The gross weight calculated for the CCD is approximately 87,000 lb.

66 The weight buildup of the BWB design was formulated using the same methods as the conventional design, with the exception of the centerbody. While equations from Raymer s text were used to calculate weights for conventional features such as the outboard wings, vertical tails, and landing gear, a different equation had to be used for the centerbody weight. For the fuselage and aft body weight estimates, equations from the NASA report A Sizing Methodology for the Conceptual Design of Blended-Wing-Body Transports by Kevin Bradley was used. This report detailed the process of adapting the Flight Optimization System (FLOPS) to provide analysis for BWB aircraft. In this report, equations were provided that related the weight of the fuselage and aft body to the surface area of the cabin, the area of the aft body, and the number of engines. These equations are shown in Eqn. 20 and 21. (20) (21) In general, because of the unfamiliarity of the BWB design and the lack of comparable aircraft, it was assumed that sizing and weight equations for the conventional design would hold true for the BWB design. The gross weight calculated for the BWB is approximately 90,000 lb. STABILITY AND CONTROL LOCATION OF CENTER OF GRAVITY Eqn. 21 states that to determine the center of gravity of the aircraft, about a fixed point and divide by the total weight. (21) The location of wings, canards, tails and overall geometry were chosen during the calculation of drag on the aircraft. The wing center of mass was placed at its aerodynamic center plus a quarter of the mean geometric chord of the wing. Similarly, the canard s center of mass was placed with respect to the distance from the nose to its leading edge while the vertical tail was taken to be the leading edge plus half its chord. The engines were placed at the

67 leading edge of the vertical tail(s) for both aircraft design for noise shielding benefits. The engine, nacelle, starter, APU, and engine controls were placed at the leading edge of the vertical tail. Since no equation for a canard was provided in the text, the weight of a horizontal stabilizer was used instead. The main landing gear for the CCD was placed under the wing where it can be expected to have adequate structural support. The nose landing gear was placed at 10 % of the length of the fuselage. The moments of other weights like furnishings, electronics, hydraulics and the fuselage were taken about the half the length of the fuselage, to approximate a distributed load. The moments of the payload and crew were also taken to be half the distance of the fuselage while the moments of the pilots were taken to be at 5 % of the length of the fuselage. Upon determining the fuel weight, a fraction of it was determined to be able to fit in the wing while the other component would have to be placed in the fuselage. The wing fuel moment was taken about the center of mass of the wing while the fuel placed in the fuselage was placed in between the canard and wing where it would best stabilize the aircraft. This placement of the fuselage and fuel tank was determined to be acceptable because the space beneath the aircraft s floor boards is not used for any other storage or structure. The empty weight and center of gravity (c.g.) for the BWB design was calculated similarly, except it did not include a horizontal stabilizer and had two vertical tails instead of one. For the BWB design, the landing gear was placed not in concern with the structural support, but keeping the center of gravity ahead of the landing gear at 70 % of the length of the aircraft. Although the wings of the BWB are large enough to hold nearly most of the fuel, 30% of the fuel was again put in the fuselage near the front to bring the c.g. forward. For further inquiry about the weights of each component in the CCD and BWB designs, refer to tables 1 & 2 of the Appendix. Table 10 gives important dimensions both aircraft designs.

68 Table 10: Dimensions Component CCD BWB Design Total Length 110 [ft.] 70 [ft.] Wing LE 70 [ft.] 20 [ft.] Wing Aero. Center 86.6 [ft.] 49.7 [ft.] Wing s MGC 10.7 [ft.] 33.6 [ft.] Canard LE 15 [ft.] N/A Vertical Tail 100 [ft.] 70 [ft.] The c.g. moves along the length of the aircraft as the aircraft is loaded in different configurations. It is therefore important to know the location of c.g. for the purposes of static stability of the aircraft. The largest shift in c.g. will occur with the shift in fuel and payload. Table 11 states the aircraft weight definition used to calculate the center of gravity. Table 11: Weight Definitions Definition of Configuration Empty Weight (EW) Structural weight which cannot be removed Operating Empty Weight (OEW) EW + 1% trapped fuel and pilots OEW plus fuel (OEW +F ) EW + all fuel and 2 pilots Max. Take Off Weight (MTOW) OEW +F + 10 passengers and 4 crew MTOW minus Fuel ( MTOW -F ) MTOW with 99 % fuel removed MTOW minus Fuel & passengers MTOW -(F &PAX ) The neutral point is the aerodynamic center of the entire aircraft; the point where all the moments sum to be zero. It is also important in terms of longitudinal stability because placing a c.g. past the neutral point causes the aircraft to pitch up while most aircraft are designed to pitch down. The neutral point for both designs was determined through Eqn. 22, which is given in Raymer. (22) Where, = lift curve slope of the wing = partial derivative of the coefficient of moment of the aircraft with respect to the fuselage angle of attack. = ratio of dynamic pressures on the horizontal tail and wing

69 = horizontal area = wing area respectively F p = propeller force (neglected) = non-dimensional position However, both c.g. and the neutral point were calculated with dimensional values while the static margin was non-dimensionalized by dividing with mean geometric chord of the wing. The following equation for static margin defines it as the difference between the neutral point and c.g. location and is non-dimensionalized by dividing with the wing s mean geometric chord. Historically the static margin of general aviation aircraft can be anywhere between %. The neutral point for the two designs is shown in Table 12. (23) Table 12: Neutral Point of A/C Neutral Point (From nose of A/C) CCD 73.2 ft BWB 49.6 ft The following two tables, Table 13 and Table 14, and figures, Figure 59 and Figure 60, state the c.g. of aircraft and static margin at different loads defined earlier. The plots help visualize the movement of the c.g. at different loads; note that the location of the c.g. is measured from the leading edge of the wing. The aim during this analysis was to produce a positive static margin but below 30% for all loads. A low static margin will make the aircraft too responsive and unstable while a high static margin will not allow the aircraft to pitch up. For the CCD, the largest static margin (CG aft) is with MTOW -F at 24%, but is assumed to be low enough to not cause a problem during landing. Changing this configuration would lead to a higher static margin at other loads, thus this configuration is acceptable.

70 Weight [lb] Table 13: C.G. and Static Margin of CCD C.G. from LE of wing [ft] Weight [lb] Static Margin EW , % OEW , % OEW +F , % MTOW , % MTOW -F , % MTOW -(F &PAX ) , % 8.5 x 104 CCD CG Travel the CCD CG from LE of Wing [ft] Figure 59: CG Travel in CCD Design For the BWB design, the largest static margin occurs again at MTOW -F but is lower than Table 14: C.G. and Static Margin of BWB Design C.G. from LE of wing [ft] Weight [lb] Static Margin EW , % OEW , % OEW +F , % MTOW , % MTOW -F , % MTOW -(F &PAX) , %

71 Weight [lb] 9.5 x 104 BWB CG Travel CG from LE of Wing [ft] Figure 60: BWB CG Travel CANARD SIZING For initial canard sizing, the canard on the conventional canard design was designed to carry 10% of the design s total lift. In order to accomplish this, one must simply take the definition of coefficient of lift and solve for area. This definition is shown in Eqn. 24. (24) Where S = wetted area L = lift of the canard q = dynamic pressure c l0 = coefficient of lift at zero angle of attack c lα =coefficient of lift due to angle of attack α = angle of attack Assuming the aircraft is in steady level flight, the aircraft s total lift is equal to the aircraft s total weight. Therefore, the lift used in this equation is simply 10% of the design s total weight. Dynamic pressure for this was calculated from design cruise conditions. The coefficients of lift were design parameters chosen at the beginning of sizing.

72 In order to obtain a final canard size, one must properly size it to keep the aircraft stable longitudinally during flight. The two most important cases to look at are takeoff rollout and landing flare. The canard must provide enough force to pitch the aircraft up in both cases. This means the canard must counteract the wing s moment about the design s center of gravity. Eqn. 25 shows a simple moment sum about the center of gravity. (25) Where M = pitching moment L = lift d = distance from the center of gravity. c = canard w = wing From this equation, it is possible to find the wetted area of the canard by breaking up the moments and lifts in Eqn. 25 by subbing in the definition of coefficient of moment and lift respectively then solving for wetted area. Eqn. 26 shows Eqn 25. broken apart into the necessary known variables, and then solved for the canard wetted area. In this equation, q is the dynamic pressure, S is wetted area, c m is coefficient of moment, c l0 is coefficient of lift at zero angle of attack, c lα is coefficient of lift at zero angle of attack, and α is the aircrafts angle of attack. The final canard size for our conventional canard design is 340 ft 2. (26) Where S = wetted area q = dynamic pressure c l0 = coefficient of lift at zero angle of attack c lα =coefficient of lift due to angle of attack c m = coefficient of moment α = angle of attack

73 VERTICAL TAIL SIZING Tail sizing is necessary for any flying aircraft to remain stable, or at least controlled to be stable, in yaw. Without a large enough vertical tail, an aircraft could become unstable and unable to fly. In order to prevent this, the vertical tail on our designs must be large enough to be able to overcome any unwanted yawing moments with a maximum of a 20 degree rudder as a safety factor. In order to ensure our aircraft could fly under any circumstance, tail sizing was conducted for two cases. The first case was to make sure our design could fly with one engine out. This being the case, one engine would provide a positive yawing moment around the design s center of gravity from its thrust, while the other engine would also provide a positive moment due to the extra drag. This yawing moment must be overcome by the vertical tail in order to ensure stable flight. The vertical tail can be properly sized by setting the sum of the moments from both engines and the vertical tail to be zero, as shown in Eqn. 27. (27) Where Meo = moment provided by the operable engine Meu = moment provided by the inoperable engine Mvt = moment provided by the vertical tail The moment on the operable engine can be found from the thrust given by the engine and the engine s distance from the centerline of the aircraft. The drag on the inoperable engine can be estimated using an equation from Raymer s textbook shown here in Eqn. 28 (28) Where D = drag on the inoperable engine Aeff = effective area of the front of the nacelle q = the dynamic pressure.

74 From here, one solves Eqn. 27 for the moment of the vertical tail. Then divide by the distance from the aircraft s center of gravity to the vertical tail to get the force the vertical tail must produce to overcome these moments. Once the necessary force for the vertical tail was obtained it became possible to calculate the size of the vertical tail. This was done by taking the definition of a lifting force and solving for the vertical tail area as shown in Eqn. 29 and Eqn. 30. (29) (30) Where N = total yawing force q = dynamic pressure S = vertical tail area c nβ = coefficient of lift due to sideslip β = sideslip angle c nδr = coefficient of lift due to rudder deflection δ r = rudder deflection angle The second case that was needed to properly size the vertical tail was a twenty degree landing sideslip angle. For this case, the vertical tail must overcome the moments provided by the fuselage and the wings about the center of gravity due to this sideslip angle. In a similar manner to the one engine out case, we calculated the necessary vertical tail size by summing the moments around the center of gravity. Before this can be done, however, the moments needed for this case must be calculated. Raymer s book provides equations that allow us to calculate both of the unknown moments needed for this case. Eqn. 31 provides an estimation of the coefficient of yaw for a fuselage surface due to sideslip. Eqn. 32 provides an estimation of the coefficient of yaw for the wings due to sideslip.

75 (30) Where c nβf = coefficient of yaw for a fuselage surface due to sideslip V = fuselage volume S b = fuselage wetted area L = fuselage length D f = fuselage parasite drag W f = fuselage weight (31) Where c nβw = coefficient of yaw for the wings due to sideslip c l = wing coefficient of lift A = wing aspect ratio Λ = wing sweep angle x ac = position of the wing aerodynamic center x cg = position of the center of gravity of the aircraft Once calculated, these coefficients can be used to determine the yawing force using the definition of these coefficients, and then sequentially the yawing moment by multiplying the yawing force by the distance from the force location to the center of gravity. These moments can then be summed to find the necessary moment the vertical tail needs to produce. From here, using the exact same steps as in the engine out case, the size of the vertical tail can be found using Eqn. 29 and Eqn. 30. The main sizing code is programmed to calculate the vertical tail size and then return the largest value for both designs. The final vertical tail size for the CCD aircraft is listed below in Table 15. The BWB s tail is assumed to be the size of the nacelles. Table 15: Final Vertical Tail Size Design Final Vertical Tail Size Conventional Canard 200 ft 2

76 NOISE Noise is considered the most important environmental concern for many people who live or work near an airport. In general, the greatest noise levels are experienced close to the airport. However, disturbance can also occur many miles away under aircraft approach and departure routes. Noise disturbance around an airport is caused by many events. These events include: Aircraft in the air Reverse thrust used by aircraft to slow down after landing; Aircraft on the ground (including taxiing, engine testing and running on-board electrical generators) Road traffic to and from the airport Construction activity Aircraft noise disturbance is commonly measured as a function of the number of aircraft and the noise levels made by each, this provides the noise climate. REGULATION The control of aircraft noise is complicated by the range of ownership and operation of aircraft, involving companies which may not have an office at the airport or even in a particular country. The way in which aircraft are flown is not only determined by the pilot, who is completely responsible for the safety of the aircraft, but also by the safety rules set by the aircraft maker, the airlines and the Civil Aviation Authority. Agreements, laws, regulations and guidelines on the control of aircraft noise are laid down by a number of different organizations including: - Department of Environment Food and Rural Affairs; - Department of Transport; - Civil Aviation Authority (CAA); - National Air Traffic Services (NATS, part of the CAA); - Airports Council International; - International Air Transport Association; - International Civil Aviation Organization.

77 AIRCRAFT CLASSIFICATION Aircraft are classified according to the noise levels they produce by the International Civil Aviation Organization (ICAO). Certification is based upon an international scale of four chapters. These are: Unclassified the first generation of jet aircraft, which are now banned by international agreement. Chapter 2 the older, noisier aircraft which have been phased out or upgraded by Chapter 3 the more modern, quieter aircraft. Chapter 5 the modern propeller aircraft. Our Aircraft will be flying under Chapter 3, due to the engine selection being a turbofan. Even though our business jet can be classified, the noise for an aircraft varies depending on altitude, loading, weather and the nature by which it is flown by the pilot. These variations must be considered when laying down standards. ENGINE TECHNOLOGY AND AIRCRAFT NOISE There are five main types of aircraft engines which contribute to the levels of noise generated in the vicinity of modern airports: a) turboprops low noise emission, small aircraft b) pre-1963 turbojets high noise emission, uncertified aircraft (e.g. Concorde) c) Low bypass turbofan engine aircraft (e.g.dc9) d) High bypass turbofans (e.g. B ) e) Propfans For our jet, we have chosen to use turbofan engines much closer to that of a High Bypass turbofan. This causes our aircraft to emit less noise, since the higher bypass engines are quieter. Specifically, we have chosen to use a geared turbofan with a bypass ratio of around 8. This engine is projected to be 15dB lower than the Stage 4 requirements.

78 NOISE CERTIFICATION OF NEW AND UPGRADED AIRCRAFT Noise levels from individual aircraft are generated by the noise certification requirements given in the Federal Aviation Regulations (FAR) Part 36 in the USA and Convention on International Civil Aviation (ICAO) Annex 16. The parameter used to assess aircraft noise is the effective perceived noise level (EPNL) measured in EPN db. This depends on: 1) Annoyance perceived by the ear 2) Tonal content of the spectrum 3) Time during which the aircraft noise remains within 10 db of the peak noise at the measurement position. Expressing these constraints in a mathematical model nets Eqn. 32. ) (32) Where PNLT is the tone corrected perceived noise level of the flyover. Its derivation uses a 1/3-octave spectrum of the sound. It is derived at ½ second intervals during the flyover. It is approximately equal to the db(a) level + 13dB. t 1 and t 2 are the times between which the PNLT is within 10dB of the peak. In certification the EPNL is measured at specified positions beneath the approach and take off flight paths and on either side of the take-off flight path as seen in Figure 61. Noise certification limits are shown in Figure, Figure B, and Figure C in Appendix C. Figure 61: Noise path

79 NOISE MANAGEMENT The key to the control of noise disturbance is reduction at source i.e. by airlines operating the quietest aircraft. However, the advancements in engine design have also served to reduce noise from aircraft. A significant reduction in noise around airports can be achieved by modification of flight operations. These operations could include: 1. Thrust cut back after takeoff. This is effective close to the airport but levels may increase further away due to the decreased altitude of the aircraft. 2. Two-segment landing approach, incorporating a steeper decent path than the usual approach. This is less significant than a reduced thrust climb because approach levels are quieter than those at takeoff. The higher altitude will cause less noise on ground level. 3. Planned preferred routing. After takeoff the flight path is prescribed over regions least likely to suffer from noise impact, i.e. away from towns and over sparsely populated areas. The reduction in noise around airports that can be achieved by improving technologies on the aircraft is influenced differently by each feature on the aircraft. Shown below in Figure 62 are the different factors that contribute to the noise of an aircraft.

80 Figure 62: Noise Contributions (Federal Aviation Administration) Examining the factors for noise, the best improvement would come from a new engine design. The airframe also contributes to the noise, but at a much smaller degree. The Blended wing body is projected to reduce noise due to noise shielding from the body itself. The specific values for how much reduction it gains, however is very hard to estimate. Therefore a conservative factor of 5% was added into the airframe factor for that section. Other than the noise shielding, the blended wing body has the same features regarding noise as the conventional canard design. NOISE ESTIMATES Shevell: Using an estimation method proposed by Stanford professors Ilan Kroo and Richard

81 Table 16: CCD Noise levels Takeoff [db] Sideline [db] Approach [db] Total [db] Design Condition (15 db below) Stage Engines produce max Stage 4 noise Table 17: BWB Noise Levels Takeoff [db] Sideline [db] Approach [db] Total [db] Design Condition (15 db below) Stage Engines produce max Stage 4 noise COST A cost analysis, as well as other previously demonstrated trade studies, quantified the choice between the Blended Wing Body and the Conventional Canard Design for the best overall design. The methods used to compute cost are from the Raymer s textbook. Aircraft cost estimation is a rough process based on statistical analysis which incorporates research, development, test and evaluation (RDT&E) of an entire aircraft program. METHODS USED TO COMPUTE COST The process is begun by using cost estimating relationships (CER) that output either cost or labor hours that are then converted to cost by multiplying the appropriate hourly rate. Raymer states, RDT&E and production costs are frequently combined to develop CERs. For the cost analysis of the BWB and the CCD aircraft, the RAND Development and Procurement Costs of Aircraft (DAPCA) IV Model, which uses a set of CERs for conceptual aircraft design, is used. DAPCA estimates the hours required for RDT&E and production by engineering, tooling, manufacturing, and quality control groups. These hours are then multiplied by hourly rates to yield costs. The costs were found by creating a MATLAB script Appendix E that takes in

82 information of weights, velocities, fuel burn, and others and inputting them into CERs. It was the only way to be able to frequently update the cost of both the Blended Wing Body and the Conventional Canard Design aircraft. Engineering hours are meant for the entire design of aircraft. This includes airframe design, testing, configurations, and system engineering. Avionics and propulsion systems are not included under engineering hours. Tooling hours are meant for the entire preparation of production. This includes making tools, manufacturing programming, etc. Manufacturing hours are meant for the direct labor to create the aircraft, through the entire process. Quality control hours are a part of manufacturing but estimated separately because it involves tools, figures and aircraft subassemblies. RDT&E also incorporates development support costs, flight test costs, manufacturing materials cost, and engine production cost. Development support and flight test costs are, but not limited to, non-recurring costs that include simulations, flight test aircraft as well as planning for flight operations. An assumed assumption of four flight test aircraft was made through information from Raymer. Manufacturing material costs are dependent on the raw materials as well as purchased hardware and equipment from which the aircraft is built. Engine production cost is 30% higher than predicted because of a more advanced gearbox turbofan engine. The DAPCA equations do not include the cost to develop a new engine. Additionally, an assumed Avionics and interiors cost of 10% was added to the fly-away cost. Avionics are typically $5000 per pound and interiors are $3500 per plane. All of these hours are included into the DAPCA equations, with respective technology factors that include materials as well as difficult designs. Engine hours, tooling hours, manufacturing hours, and manufacturing materials are dependent on the predicted empty weight of the aircraft, maximum velocity, and the production quantity of aircraft (in five years). Flight test costs are dependent on weight, velocity and flight test aircraft. Development support cost is determined by weight and velocity. A quality control hour is a fixed constant and engine production cost is dependent on engine maximum thrust, turbine inlet temperature, and engine maximum Mach number. The equations are located in the cost function in Appendix E.

83 Once those values are determined, the hours estimated with the model are multiplied by the appropriate hourly rates to calculate labor costs. These wrap rates include the direct salaries paid to employees as well as the employee benefits, overhead, and administrative costs. These assumed rates from Raymer are listed as: engineering =$86, tooling = $88, quality control = $81, and manufacturing = $73. DAPCA does not handle the most advanced designs very well, so cost estimates were increased as 20% for the Conventional Canard Design and 40% for the Blended Wing Body. Because of the canards and the composite materials Conventional Canard Design is a 20% increase while the Blended Wing Body is 40% because of the introduction of newer manufacturing techniques, new designs, and tougher engineering work. A production and market price is predicted later in the analysis for both of designs. According to Raymer, the market cost is roughly estimated by 30%. In the figure shown below, the market price of an aircraft is plotted against the amount of aircraft sold. To determine the price of an aircraft, the entire developmental cost is divided by the predicted quantity of aircraft to be manufactured. Figure 63: Market Price vs. Number of Aircraft Now the operating costs are calculated differently than the production cost of an aircraft. Variable costs are dependent on the usage of the jet. These include fuel costs, maintenance costs, and taxes. In the cost analysis, oil costs were assumed negligible and fuel costs were assumed at 60 cents per pound of jet fuel. The higher cost of fuel is expected to

84 increase due to indications that the world will reach peak oil before the year 2015 (Owen, Inderwildi and King). Maintenance costs are based on upon how often the aircraft breaks and the average cost to fix it. Raymer has an equation of material cost per flight hours that is based on the cost of each aircraft, cost per engine, and the number of engines. Furthermore, in the near future, a carbon tax may be implemented throughout the world based on amount of fuel burned by aircraft. Based on the social cost of carbon, the United States Department of Energy claims that, an aircraft may be taxed $0.016 per pound of jet fuel in the future (US Department of Energy; Energy Information Administration). Fixed costs are the costs of owning an aircraft. A typical jet aircraft flies 500 block hours or the total time the aircraft is in use, from when the blocks are removed from the wheels at the departure airport to when they are placed on the wheels at the destination. Fixed costs include insurance, crew costs, hanger costs and training. The equation for the cost of a crew, mentioned in Raymer s text book, is dependent on the cruise velocity and take-off gross weight. The equation should be viewed as rough approximations only. Hanger costs and training are one hundredth of a percent of the entire aircraft, and were estimated using costs assumed from the source Plane Quest. For a G550, hanger costs are $80,000 and training costs are $40,000. An assumption of a higher training cost was used for the BWB since it is a new design to the aircraft industry. Additionally, insurance costs for aircraft are roughly 1.5% of the cost an aircraft per year. This may be higher than a typical aircraft today, but both these designs are non-conventional and may require a higher insurance. Depreciation may be included as a fixed cost; however, the cost analysis completed did not. Depreciation is the allocation of the purchase price over the operating life of the aircraft. A straight line schedule (or ADS deprecation) is a reasonable estimate. The rates for deprecation used were an assumed resale value of 10%, with a depreciation period of 12 years. Thus 0.9 is multiplied by the value of the aircraft, and then divided by 12 for a yearly depreciation value. Most of these estimated costs have lots of variation. Additionally, since these costs are estimated in 1999 dollars, a future worth value calculation is used to determine the costs in 2010 dollars. This consists of averaging an inflation rate from the consumer price index from

85 the years of 1999 to Then the inflation rate is inputted into an equation to determine the future value of money. All of the calculations shown in the next sections are in 2010 dollars. COST VALIDATION The G550 was used to validate the cost estimation script because it contained similar inputs to the Conventional Canard Design aircraft. As shown in the table below, the average market value of a G550 is roughly $59.9 Million/Aircraft. It is averaged because each aircraft can be created to specification for the customer. In January 2009, there were 190 G550 s in service according to Jane s All the World s Aircraft. This was inputted into the quantity variable for the cost function. Additionally, an appropriate total operating cost found is listed in the table below (Conklin and de Decker). It shows in the table that the cost analysis function validates well with current aircraft created today based on information given. Table 18: Cost Validation Cost Breakdown Actual Calculated Cost of Market Value (30% Investment Rate) $59.9 Million/Aircraft $57.6 Million/Aircraft Total Operating Cost $6,000/hour $5,800/hour NUMBER OF AIRCRAFT IN PRODUCTION RUN In order to quantify how many aircraft would be produced in a five year span, a market analysis and estimations from similar aircraft had been done. By observing the recent economic downturn and the rising Asian market, it was concluded that the amount of aircraft sold can be relatively stable and comparative to the past five years. Thus, a determined 200 aircraft manufactured and sold for the Conventional Canard Design deemed reasonable and 150 aircraft for the Blended Wing Body. The Conventional Canard Design estimate was based on the introduction of easier manufacturing methods by using composite molds for aircraft parts. Also, information about the 190 G550 aircraft in service since its introduction in 2004 was used. However, the Blended Wing Body aircraft would possibly be more difficult to produce because of the complex fuselage and design capabilities. The composites would increase the amount of Blended Wing Body aircraft to be produced, but the design work and possibly introduction of new manufacturing tools would decrease the amount produced. Since there are no current Blended Wing Body aircraft in production, it is more difficult to estimate that number.

86 ESTIMATED COST TO DEVELOP AND MANUFACTURE Located in the tables below are the estimates for both the Conventional Canard Design and the Blended Wing Body aircraft, including their respective technology factors. As shown, the Conventional Canard Design has a lower initial purchase cost than the Blended Wing Body. For both designs, 129 aircraft are needed to break even with the current investment rate of 30 percent. Table 19: Cost Breakdown of CCD Cost Breakdown of CCD (1.2 tech factor) Value RDT&E $7.9 Billion Production $51.7 Million/aircraft Market Value (30% Investment Rate) $62.0 Million/aircraft Yearly Depreciation Value $3.9 Million Table 20: Cost Breakdown of BWB Cost Breakdown of BWB (1.4 tech factor) RDT&E Cost of Production Market Value (30% Investment Rate) Yearly Depreciation Value Value $9.5 Billion $82.5 Million/aircraft $107.3 Million/aircraft $6.2 Million ESTIMATED OPERATING COST Estimating the operating cost consisted of the method discussed earlier. Additionally, the cost of crew for both aircraft is for three. As shown below, the Blended Wing Body has a lower variable cost but a higher overall operating cost. The lower variable cost is because the Blended Wing Body uses less fuel than the Conventional Canard Design. The insurance is also double compared to the Conventional Canard Design because of a higher production cost. Furthermore, cost estimation can be an inaccurate analysis because of certain technology factors; however, it is a good starting point for understanding the range of costs that the aircraft may be in.

87 Table 21: Operating Cost Breakdown of CCD Cost Breakdown of CCD Value Total Operating Cost $5,200/block hour Variable Cost $2,900/block hour Fixed Cost $2,300/block hour Maintenance Cost $1177/block hour Crew Cost $428/block hour Cost of Fuel(max design range) $1700/block hour Tax $51/block hour Hanger $180/block hour Training $100/block hour Insurance $1550/block hour Table 22: Operating Cost Breakdown of BWB Cost Breakdown of BWB Value Total Operating Cost $5700/block hour Variable Cost $2500/block hour Fixed Cost $3200/block hour Maintenance Cost $1450/block hour Crew Cost $426/block hour Cost of Fuel(max design range) $1050/block hour Tax $31/block hour Hanger $180/block hour Training $120/block hour Insurance $3300/block hour RESULTS After sizing the aircraft, we were able to determine if our aircraft met our desired target values. While were managed to achieve our threshold values for the aircraft, we were unable to reach our target for some of the dimensions and noise considerations. Since these are both lower priorities for us, however, the results are consistent with our design intentions.

88 Table 23: Resulting Compliance Matrix CCD Threshold Target Target Achieved Range (Still-air) [nmi] 7,100 4,500 7,100 Yes Ground Roll Take off Distance <# [ft] 3,600 4,500 4,000 Yes Empty Weight [lb] 40,000 55,000 45,000 Yes Fuel Weight [lb] 35,000 45,000 25,000 No Cabin Volume [ft^3] 2,000 2,000 2,200 No Cruise Speed [Mach] Yes Cumulative Certification Noise Level[dB] No Cabin Height [in] No Our aircraft was unable to make most of the N+2 requirements, however, we did manage to make significant improvements towards that end, this is shown in Table 24. Table 24: NASA N+2 Goals CORNERS OF THE TRADE SPACE N+2=2020 Achieved N+2 Goals Noise -42 db No: db (cum below Stage 4) LTO NO x Emissions -75% No: -50% (below CAEP 6) Performance: Aircraft Fuel Burn Performance: Field Length -50%** No: -12% -50% Yes

89 SUMMARY AND NEXT STEPS In summary, our sizing successfully models existing aircraft and thus also models our own aircraft with a degree of accuracy. Our results indicate that our aircraft successfully meets our design requirements, without increasing our weight nor size compared to modern aircraft. In summary, we have designed a stable aircraft that can successfully take off with the benefits of destination flexibility, fuel efficiency, and a longer range than modern aircraft of the same weight class. While it does not meet N+2 goals, it is still significantly quieter and more fuel efficient than current aircraft. From here, we will continue to refine our modeling, by first improving our drag prediction methods. We would also like to perform actual tests, running models of our aircraft through either real wind tunnels or CFD code to get better data concerning its characteristics.

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92 25. Rolls Royce. "Rolls Royce Engine Datasheets." 26. SAE International. "Aerospace Engineering Online." < 27. Saravanamuttoo, HIH, et al. Gas Turbine Theory, 6th Edition. Harlow: Pearson Education Limited, Snecma. "Research and Technology at Scecma to ensure sustained development of air transportation." 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit. Tuscon, Arizona, AIAA "The Turbojet Engine." February 2010 < 30. Thomas, Geoffrey. "Rolls-Royce Pursues Open Rotor." November March Trippensee, Gary and David Lux. "X-29A Forward Swept Wing Flight Research Program Status." Turboprop Engines. 23 December February 2010 < 33. Ullman, D.G. and B.P. Spiegel. "Trade Studies with Uncertain Information." Sixteenth Annual International Symposium of the International Council On Systems Engineering US Department of Energy; Energy Information Administration. "Voluntary Reporting of Greenhouse Gases Program." Fuel and Energy Source Codes and Emission Coefficients. 35. Vatandas, Erguven. "Geometrical and positional optimization of the foward swept lift producing surfaces in 3D flow domains." Aircraft Engineering and Aerospace Technology: An International Journal Weisstein, Eric W. "Circle-Circle Intersection." Mathworld - A Wolfram Web Resource. April 2010 < 37. What's a Scramjet? 22 November February 2010 <

93 APPENDIX A: HOUSE OF QUALITY