飞机总体方案设计例子.doc

Team V; 107=Preliminary Design ReviewAAE 451 Aircraft Design Final ReportTeam V27 April, 2006Stephen BeirneCharlie RushMiles HatemZheng WangChris KesterBrandon WeddeJim RadtkeGreg WilsonExecutive Summary The Barn Owl is Team Vs solution to the need for an alternative fuel aircraft within the single engine general aviation market. This need will arise as peak-oil approaches and petroleum-based fuels become scarce and expensive. The Barn Owl has also been designed with the near-term need for a replacement for 100 octane low-lead aviation gasoline in mind. 100 low-lead is expected to be phased out in the next 10 years. The value of this near-term market has been estimated at over $1 billion annually once 100 low-lead disappears.Previous work by Team V has focused on defining requirements for the Barn Owl and developing the aircraft concept. In this Preliminary Design Review, the feasibility of the Barn Owl is demonstrated by detailed analysis of the concept and its features.A weight fraction-based sizing code was used to produce constrained carpet plots. A light-weight design was selected from the feasible space, and used for further detailed design. This sizing approach was continually updated as the detailed analyses were refined. Detailed aerodynamic analysis consisted of several coordinated efforts. The airfoil cross-section was designed using a genetic algorithm. The wing planform was optimized to have an elliptical lift distribution at the cruise condition. A three-dimensional CMARC analysis provided detailed aerodynamic performance predictions. The team developed a structural layout and selected materials based on cost considerations. Detailed structural analysis was carried out, optimizing the wing structure and confirming the viability of the fuselage structural layout.Component weights were determined and used in calculating the static longitudinal stability of the Barn Owl. The team succeeded in demonstrating reasonable stability. Stall-spin stability was also considered and assured.For propulsion, an existing diesel engine and existing propeller were selected. The use of this off-the-shelf technology will reduce the cost of development and ease manufacturing.The Barn Owl is projected to cost $278,700. This price and the Barn Owls solid performance will make it a competitive aircraft in the general aviation market.Table of ContentsIntroduction2Design Requirements2Proposed Concept2Design Mission2Sizing22D Aerodynamics23D Aerodynamics2Performance2Structures2Weight and Balance & Stability2Propulsion2Cost2Final Design Comparisons2References2Appendix A Additional Barn Owl Pictures2Appendix B Sizing/Carpet Plot Code2Appendix C Drag Polar Data2Appendix D Wing Structure Optimization Code2Appendix E Wing Optimization Code Outputs2Appendix F Empty Weight Calculation Code2Appendix G Aluminum Equivalent Inputs for Appendix C2Appendix H Fiberglass Property Calculation Code2Appendix I V-n Diagram Code2Appendix J Flight Envelope Code2IntroductionThis teams system requirements review 14 described a plan to target the general aviation market with a single-engine, four-seat aircraft the Barn Owl. It was the goal of Team V to design a product running on an alternative fuel that will be marketable to hobbyists, fixed base operators, and training fleets in the phase-out of 100 octane low-lead (100LL) aviation gasoline and the transitional times of “peak-oil.” The system requirements review also included an estimate of the potential market. It was determined that sales of 500 aircraft per year, or more, could be expected in the wake of the 100LL phase-out.A system definition review was conducted to develop the concept of the aircraft 13. Bio-diesel was selected as the alternative fuel for the Barn Owl. The configuration was set as a low-wing, conventional tail aircraft with a piston engine powering a tractor propeller.The following report describes the method used to determine the preliminary design. Design RequirementsThe current design requirements shown below in Table 1 are still the same as presented in this teams System Definition Review 13. Through the work presented in this report, it will be shown that at the current level of detail an alternate-fueled aircraft can meet all the given requirements.Table 1 - Design RequirementsDesign Requirements Summary1500 ft takeoff distance to clear a 50 ft obstacle600 lb payload with max fuel125 kts max cruise speed, Target = 150 kts500 nm range, Target = 600 nm48x44 cabin height x widthTarget GTOW of 2800 lbs Target base price of $300,000 Proposed ConceptTeam Vs proposed concept is shown in Figure 1. The Barn Owl is a low wing, conventional tail, single engine, tractor prop, 4 seat general aviation aircraft.Figure 1 - Concept 3 ViewDesign MissionBased upon this teams QFD analysis, as presented in the system requirements review, it was determined that to be competitive in the chosen market the aircraft would need to have a range of 600 nautical miles and cruise near 8000 ft. From this the design mission was formulated.Figure 2 Design MissionTable 2 Design Mission LegendATaxi 14 minutesFClimb to divert altitude (2000 ft MSL)BTake-off roll at sea levelG45-min loiter / divertCClimb to cruise (8000 ft MSL)HDescend to sea levelDCruise at 150 KTASILanding RollEDescend to sea levelJTaxi to hangerThe mission begins with an estimated 14 minute taxi to the runway. Next, the plane will begin its takeoff roll at a runway located at sea level. The mission has been designed from sea level since it is a good benchmark from which to measure altitude. After takeoff the plane will climb at 700 fpm to its cruising altitude of 8000 ft MSL and cruise at that height at a speed of 150 KTAS. After cruising the specified 600 nmi, the aircraft will descend back down to sea level. Just before touching down, it will then climb to a divert/loiter altitude of 2000 ft MSL. It will then either spend 45 minutes loitering in pattern or diverting to an alternative airport as per FAR fuel requirements for IFR flight. Finally it will descend again to sea level, land, and taxi back to the hanger.SizingSizing and Carpet Plot CodeSizing of the Barn Owl was done through the use of carpet plots based on a weight fraction approach and its implementation in a Matlab script. The script was used to automatically generate the carpet plots with constraint lines, so that the lightest weight feasible design point could be determined. Since an existing engine was selected for the Barn Owl, as described below, fixed-engine sizing was used. Thus, carpet plots sizing the aircraft with different aspect ratios over a range of wing loadings were used to set the design.The script uses an iteration scheme which assumes an initial guess for gross takeoff weight (GTOW), then calculates GTOW based upon that guess. It then iterates the guessed weight until the two weights are within 0.01% of each other (approximately 0.27 lbs). Within the code, GTOW was calculated using the following equation: Equation 1The empty weight fraction was calculated by first calculating all of the component weights based upon chapter 15 of Raymer 10; the details of which are discussed later in this report. The weights of the individual components were then summed to give the total empty weight which is then divided by the GTOW guess. This yields the empty weight fraction.The fuel fraction was calculated using the equation: Equation 2In Equation 2, are weight fractions for the teams design mission. Specifically, represents the fuel used during the taxi segment and is calculated using Raymers equation 19.7: Equation 3where C is specific fuel consumption (SFC), d is duration in hours, and T/W is the idle (20%) thrust-to-weight ratio. Duration for this segment is assumed to be 14 minutes as specified in chapter 19 of Raymer 10. Since the engine of the aircraft was known to be 200 horsepower,the thrust-to-weight ratio was calculated using the equation: Equation 4where p is the propeller efficiency, which is assumed to be 0.86 based upon the propeller analysis discussed later. The SFC was calculated from the brake horsepower SFC (BSFC) which was calculated for the aircrafts engine to be 0.439. Engine selection and SFC calculation is discussed later in this report. The equation used to calculate SFC from BSFC is: Equation 5where V is speed and p is the propeller efficiency. corresponds to the fuel used during takeoff and is also calculated using Equation 3 with 100% thrust. Duration for the takeoff segment is assumed to be one minute as specified in Raymers chapter 19 10. represents the fuel weight used in climbing to the cruise altitude of 8000 ft. It is calculated using Raymers equation 19.8: Equation 6where C is SFC (same as previously discussed), he is the change in height energy, V is the average climb speed and D/T is the average drag divided by the average thrust during climb. It should be noted, however, that D/T is not actually calculated within the code, and is instead replaced by: Equation 7In Equation 7, L/D is calculated implicitly by calculating CL/CD. CL is calculated using the equation: Equation 8where W/S is wing loading (an input parameter) and is air density at sea level. CD is calculated using the curve fit for the planes drag polar: Equation 9he is calculated using Raymers equation 19.9: Equation 10where h is altitude, g is the gravity constant, and V is speed. Note that Vtakeoff is calculated as 1.1 times stall speed. This is conservative since it does not consider the acceleration that will occur as the plane climbs to the 50 foot altitude accounted for in the takeoff fuel weight fraction equation. The average climb speed is calculated using the Raymers equation 17.13:Equation 11where W/S is wing loading (an input parameter), is air density at sea level, CD0 is the zero lift drag coefficient, and K is the aerodynamic constant. CD0 was calculated to be 0.023 in the aerodynamic analysis. K was calculated using the equation:Equation 12Where AR is the aspect ratio and e is the Oswald Efficiency Factor. The Aspect ratio is an input parameter which was chosen through the use of carpet plots (discussed later) and the Oswald Efficiency Factor was determined as a function of aspect ratio via a curve fit of CMARC analysis data for the plane which yielded the equation: Equation 13 represents the fuel weight used during the cruise segment and is calculated using the Breguet range equation (also Raymer 19.10): Equation 14where R is range (less the distance traveled during climb), C is SFC (same as before), V is cruise speed, and L/D is the lift to drag ratio at cruise conditions. Distance traveled during the range segment is calculated by subtracting the average climb velocity multiplied by the time it will take to reach 8000 ft at a climb rate of 700 fpm from the design missions range of 600 nmi. Cruise speed is set at the design cruise speed of 150 kts. Cruise L/D was calculated in the same manner as it was for climb, except that the density and speed used were the cruise condition values, and not those of climb. represents the fuel used during descent and is assumed to be 0.9989. This was approximated by estimating the fuel usage per minute and multiplying it by an estimated time to descend. represents a missed approach and climb to a 2000 ft divert altitude. It is calculated in the same manner as the fuel used to climb to cruise altitude. represents a divert distance, however, the team opted to use a 45 minute loiter/divert segment, thus this fuel fraction is 1. represents the fuel used during the 45 minute loiter/divert segment and is calculated using the loiter equation (also Raymer 19.11): Equation 15where E is endurance time (in hours), C is SFC (same as before), and L/D is the lift to drag ratio at cruise conditions. The endurance time is specified as 45 minutes to accommodate IFR regulations and the L/D used in the loiter/divert segment is the same as that for cruise. is another descent segment and is assumed to be equal towhich should be conservative considering the aircraft is descending from a lower altitude. Finally represents the fuel used during landing and is assumed to be 0.995. This is based upon Raymers equation 6.23 which simply states it should be between 0.992 and 0.997 10.The Wcrew and Wpayload were taken from the design requirement of having a 600 lb payload including crew. Once all of these sizing equations were compiled together, it was possible to place them inside of two for loops. The first of these loops varied aspect ratio through a specified range; the other varied the wing loading. This made it possible to plot a curve of GTOW vs. wing loading for each aspect ratio. Using this for loop approach allowed the team to rapidly generate numerous data points which (with a small enough wing loading increment) could be linearly connected to adjacent points to form a smooth curve. In order to find the optimal aircraft design (i.e. the one with the minimum GTOW), Team V calculated and plotted various constraints along the aspect ratio curves. The constraints used were stall speed, cruise speed, climb rate, takeoff distance, and turn load factor (n) value.For each aspect ratio and wing loading combination, cruise speed, climb rate, takeoff distance, and the turn load factor (n) value were calculated as discussed below and placed into an individual matrix for each constraint. Then for each row in the matrix (which corresponded to a constant aspect ratio) the value above and below the desired value was found. Next, a linear interpolation was used to find exact GTOW that corresponded to the desired value for that aspect ratio. Finally, all that had to be done to create the constraint lines was plot a curve through these data points.The cruise speed was calculated by using Matlabs fsolve function to solve the following equation for V: Equation 16In Equation 16, p is the propeller efficiency, bhp is the horsepower of the engine (200), V is the cruise speed, is the air density at cruise altitude, S is the area of the wing, CD0 is the zero lift drag coefficient, K is the aerodynamic constant, W is the aircraft GTOW, and CL_min_drag is the coefficient of lift which corresponds to the minimum drag coefficient. p, V, K, and CD0 were previously discussed. CL_min_drag is based upon the drag polar generated by the aerodynamic analysis and S is calculated by dividing the GTOW by the wing loading. It can be noted that the factor of 0.75 in Equation 16 signifies that this velocity will be the cruise speed at 75% power.Climb rate is calculated using the Raymers equation 17.44: Equation 17Where p_climb is the propeller efficiency during climb, bhp is the horsepower of the engine (200), V is the cruise speed, W is the GTOW, and D is the drag force of the aircraft during climb. p_climb is assumed to be 0.76 based upon the propeller analysis (discussed later) and other than D all other parameters have been previously discussed. D is calculated using the equation: Equation 18Where is the density at sea level, V is the average speed during climb, S is the wing area, and CD0 is the zero lift drag coefficient. All of these values have been previously discussed.Takeoff distance was calculated using Raymers equation 17.112 10: Equation 19In Equation 19 W/S, , sl (same as ), CL_climb, g, and W are as previously discussed. hobstacle is the height of an obstacle to be cleared during takeoff (50 ft). G is calculated using the equation: Equation 20where T/W is the thrust to weight ratio, D is the drag force during climb, and W is the GTOW. All of these values have been previously discussed. U was calculated using the equation: Equation 21where CL_max is assumed to be 1.6 based upon the aerodynamic analysis. Tav was calculated using equation 17.144 of Raymer 10: Equation 22In Equation 22 bhp, , and sl have been previously discussed. Ne is the number of engines in the aircraft (one) and Dp is the propeller diameter of 6.166 ft (74 in.) as determined by the propeller analysis.Turn load factor was calculated using an altered version of Raymers equation 17.54 to account for a non-symmetric airfoil: Equation 23Where is the density at cruise altitude, V is the cruise speed, W/S is the wing loading, CD0 is the zero lift drag coefficient , CL is the coefficient of lift during cruise, and CL_mindrag is the minimum drag coefficient of lift. All of these values have been previously discussed.The final constraint - stall - was calculated in a different manner. Since stall speed is only a function of air density, CL_max, and wing loading, the wing loading at which the plane stalled at 57 kts at sea level could be calculated directly from the equation: Equation 2457 kts was chosen as Team Vs designed stall speed in order to leave room for error and still meet the FAR requirement of a 61 knot stall speed. Once the wing loading for stall was calculated, it was plotted as a vertical line on the carpet plot.For completeness, landing distance, best range cruise speed, and best range cruise distance are also calculated in the sizing code. Best range cruise distance is calculated because the 600 nmi range is not the best range distance, but the range for a cruise at 150 kts. Landing