Post on 19-Jan-2016
AAE 451
AERODYNAMICSPDR 2
TEAM 4Jared Hutter, Andrew Faust, Matt Bagg, Tony Bradford,Arun Padmanabhan, Gerald Lo, Kelvin Seah
November 18, 2003
TEAM4OVERVIEW
Concept Review
Aircraft CL and CM
Updated Wing Size
Aircraft Plots
Follow-Up Actions
TEAM4CONCEPT REVIEW
High WingS = 47.8 ft2
b = 15.5 ft, c = 3.1 ftAR = 5
Twin Booms3 ft apart;
7.3 ft from Wing MAC to HT MAC
Twin Engine1.8 HP each
Avionics Pod20 lb; can be positioned front or aft depending on
requirements
EmpennageHorizontal and Vertical
Tails sized using modified Class 1 Approach(per D & C QDR 1)
TEAM4AIRCRAFT LIFT COEFFICIENT
Lift Coefficient
CL = CL* + CLe*elevator + CL0
Matlab script based on Roskam Vol VI Ch8:
CL = 5.41(rad-1)* + 0.4675(rad-1)*elevator + 0.3086
Predator Codes from AAE 565
CL = 5.473(rad-1)* + 0.454(rad-1)*elevator + 0.3113
TEAM4AIRCRAFT PITCHING MOMENT
Moment Coefficient
CM = CM* + CMe*elevator + CM0
Matlab Script based on Roskam
CM = -2.0496 (rad-1)* + (-0.1771)(rad-1)* elevator + 0.0425
Predator Codes
CM = -2.2682 (rad-1)* + (-1.058)(rad-1)* elevator - 0.2785
TEAM4AIRCRAFT CL AND CM
AAE 565 Predator code similar to Roskam
Roskam uses graphs in his book
Predator has the graphs hard coded into the program
Predator will be more accurate
Update Constraint Diagram
Need Maximum CL for Constraint Diagram
Roskam Code solves for Maximum CL
.06 difference between Roskam Code and Predator for CL
TEAM4
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 10
0.05
0.1
0.15
0.2
0.25
0.3
0.35Distribution of section cl along span of the wing @ AOA=0
Percent Span (y/(b/2))
Sec
tion
cl
MAXIMUM LIFT COEFFICIENT
Need section cl along wing span
Increase Angle of Attack and find new section cls
Repeat until the wing begins to Stall
That is the stall angle Integrate section cl’s to find
Maximum CL
TEAM4
AIRCRAFT CL AND CM
Three Major Codes: Predator, CL max, and Constraint Diagram Predator:
Input: Aircraft Geometry Output: CL and CM equations
Maximum Lift Coefficient Input: Main Wing and Horizontal Tail Geometry Output: CL Max and Alpha at CL Max
Constraint Diagram Input: Flight Conditions, CL at 0 Alpha, CL max, Engine Info Output: Wing Area and Required Power
TEAM4
AIRCRAFT CL AND CM
Iterative loop can be used
Used old constraint diagram values for initial guess
Used Wing Area as the Control variable
Constraint Code
Predator Code
Max CL Code
TEAM4
AIRCRAFT CL AND CM
Lift Coefficient
CL = CL* + CLe*elevator + CL0
CL = 5.931(rad-1)* + 0.59(rad-1)*elevator + 0.2809
Moment Coefficient
CM = CM* + CMe*elevator + CM0
CM = -3.6947(rad-1)* + (-1.058)(rad-1)* elevator -.3956
Reduce CM0 for clean flight
TEAM4
AIRCRAFT CL AND CM
CM0 main contribution is from the Incidence angle of Horizontal Tail (-2.51 degrees)
Using the Iterative Loop, ran over a range of Horizontal Tail Incident angles
Found Incident Angle that reduced CM0 the most
TEAM4
AIRCRAFT CL AND CM Lift Coefficient
CL = CL* + CLe*elevator + CL0
CL = 6.0339(rad-1)* + 0.6201(rad-1)*elevator + 0.4237
Moment Coefficient
CM = CM* + CMe*elevator + CM0
CM = -4.0421(rad-1)* + (-1.058)(rad-1)* elevator + 0.00
Incident Angle=.23 degrees
Does not seem right, may be caused by Downwash from the Main Wing
TEAM4
AIRCRAFT PARAMETERS
Wing Area= 34.5 ft^2 Wing Span= 13.1 ft Max CL= 1.8034 @ 12.88 Degree Angle of Attack CD= .0339 @ 0 Angle of Attack
TEAM4
-10 -5 0 5 10-1
-0.5
0
0.5
1
1.5
2
Angle of Attack (degrees)
CL
-1 -0.5 0 0.5 1-1
-0.5
0
0.5
1
1.5
2
CM
CL
Dele 10d
Dele 0d
Dele -10d
TRIM DIAGRAM AT CRUISE
CL=.4327
TEAM4
0.02 0.04 0.06 0.08 0.1 0.12 0.14-0.5
0
0.5
1
1.5Drag Polar at Cruise
CD
CL
Drag Polar
Based on Roskam Vol VI Ch 4
TEAM4FOLLOW-UP ACTIONS
Verify CD calculations
Triple Check CM0 and Incident Angle of the
Horizontal Tail
React to changes from D+C, Propulsion,
and Structures
AAE 451
Questions?
TEAM4
AppendixLift Curve Slope CL = f(CLW, CLHT, HT, w) HT = Ratio of dynamic pressure. Mostly caused by
propeller wash and velocity Downwash, w = Caused by main wing’s vortex flow on tail. Changes effective angle of attack for the tail.
elevatorPositive
Negative
TEAM4AIRCRAFT PARAMETERS Lift Curve Slope for Elevator Deflection
CLe = f(elevator size, horizontal tail planform)
Zero Angle of Attack Lift Coefficient
CL0 = f(CL0W, CL0HT, HT, incident angles)
HT = Ratio of dynamic pressure. Mostly caused by
propeller wash and velocity
Incident angles are for both main wing and
horizontal tail
TEAM4AIRCRAFT PARAMETERS Moment Coefficient
CM = CM* + CMe*elevator + CM0
CM = -0. 0225(deg-1)* + (-0.0027)(deg-1)* elevator + 0.0280
Moment Curve Slope
CM = f(dCM/dCL, CL)
dCM/dCL = f(CG, Aerodynamic Center of Aircraft)
TEAM4AIRCRAFT PARAMETERS Zero Angle of Attack Moment Coefficient
CM0 = f(CM0_W, CM0_HT [both about the CG])
LIFT
WEIGHT
Aerodynamic Center