PFR - Aerospace Engineering Sciences Senior Design Projects ...

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Project Final Report – CUDBF April 30 th , 2009 ASEN 4028: Aerospace Senior Projects = ∙ ∙ 1 + ∙ ∙ ∙ ∙ ∙ ∙ ∙ Equation 11: Weight-to-Power ratio for cruise speed. In Equation 11, is the dynamic pressure based on the ambient density, and cruise velocity, is the parasitic drag, is the takeoff thrust to cruise thrust ratio, is the Oswald efficiency factor of the wing, is the aspect ratio of the wing, is the ratio of the cruise weight to the takeoff weight, is the wing loading and is the cruise speed. A detailed list of the assumptions and values used to size this aircraft can be seen in Appendix F. The equations used to do the performance sizing were obtained from Roskam [11] . The design point chosen from this design area is also marked in the plot. This design point gave a wing loading, W/S of 2.1 lb/ft 2 and a weight to power ratio, W/P of 0.047 lb/lb ft/s. Using an estimated gross takeoff weight of 15lb, the wing area necessary from the wing loading was determined to be 7.14 ft 2 . The power required from the weight to power ratio was determined to be 530 W. The performance sizing shown demonstrates that a number of designs that meet all the performance requirements for stall, takeoff and cruise are feasible, and an optimized design can be created that would maximize the cruise speed while keeping the weight to power ratio small. 7.1.4 Propulsion Feasibility To ensure that the propulsion system can meet all the design to specifications, a feasibility study was conducted. The two major requirements to verify are the 100 ft takeoff distance and the maximum 4 lb battery weight. The rest of the requirements can be verified by inspection. In order to make the 100 ft takeoff, the thrust required from the motors was calculated. The equation for takeoff distance is in Equation 12. S LO = g * ρ * S * C L 2 1.44 * W *{ T − [ D + µ * ( W r − L)} Equation 12: Takeoff Distance Calculation To simplify the equation, it is assumed that drag, D is very small compared to thrust and that the coefficient of friction, µ of the landing gear wheels is negligible. This yields the simplified version solved for thrust in Equation 13. r 54
Project Final Report – CUDBF April 30 th , 2009 ASEN 4028: Aerospace Senior Projects T = 2 1.44* W g * ρ * S * C L * S LO Equation 13: Simplified Takeoff Distance Calculation Where W is the fully loaded weight of the aircraft (15 lbs), ρ is the density for a 5,000 ft density altitude, S is wing area, g is gravity, C L is takeoff lift coefficient, and S LO is takeoff distance. The aircraft stalls at a C L of 1.38 and an angle of attack of 12.6°. Therefore, assuming a takeoff rotation of 10°, takeoff C L was assumed to be 1.1. To account for the small drag and rolling friction assumption, a 10% margin was added to the takeoff distance. Therefore, S LO was chosen to be 90ft. With these variables, the thrust required is 6.0lb, or 3.0lb per motor. To ensure that this thrust could be met, a static thrust stand was created shown in Figure 31. Figure 31: Static Thrust Stand The thrust stand was designed so that as the motor generated thrust, it would pull against a load cell to record the static thrust. The load cell was connected to a LabView VI to record the thrust data [12] . During static thrust tests, an Eagle Tree telemetry system was used to record voltage, current, power, and motor RPM [13] . The motor set up utilized to test the thrust available was the Neu [14] 1107 2Y motor with 3300 rpm/V at 19.0 V, 40 amps, and an 8,000 ft density altitude. To make up for the extra rpm/V of the Neu 1107 motor compared to the Neu 1110 2Y motor, the voltage was lowered below that of the estimated battery pack voltage to compensate. Several different propeller sizes were tested. The amount of thrust produced by this single motor varied between 4.30 lb and 5.87 lb of thrust depending upon the propeller size and pitch. This confirmed that the amount of thrust required to make the 100ft takeoff requirement is feasible. The other major design to specification of the propulsion system is the maximum battery weight of 4 lbs. To check the feasibility of this requirement, the amount of power required to fly four laps at maximum weight was estimated. The average current draw was assumed to be 20 amps at 55
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University of Colorado Design/Build
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PFR - Aerospace Engineering Sciences Senior Design Projects ...

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