The attached file is the questions for the assignment.1
Exercise 3: Lift and Airfoils
The first part of this week’s assignment is to choose and research a reciprocating engine
powered (i.e. propeller type) aircraft. You will further use your selected aircraft in subsequent
assignments, so be specific and make sure to stay relatively conventional with your choice in
order to prevent having trouble finding the required data during your later research. Also, if you
find multiple numbers (e.g. for different aircraft series, different configurations, and/or different
operating conditions), please pick only one for your further work, but make sure to detail your
choice in your answer (i.e. comment on the condition) and stay consistent with that choice
throughout subsequent work.
In contrast to formal research for other work in your academic program at ERAU,
Wikipedia may be used as a starting point for this assignment. However, DO NOT USE
PROPRIETARY OR CLASSIFIED INFORMATION even if you happen to have access in
your line of work.
1. Selected Aircraft:
For the following part of your research, you can utilize David Lednicer’s (2010) Incomplete
Guide to Airfoil Usage at http://m-selig.ae.illinois.edu/ads/aircraft.html or any other reliable
source for research on your aircraft.
2. Main Wing Airfoil (if more than one airfoil is used in the wing design, e.g. different between
root and tip, pick the predominant profile and, as always, stay consistent):
Please note also the database designator in the following on-line tool (see picture below):
Find the appropriate lift curve for your Airfoil from 4. You can utilize any officially published airfoil
diagram for your selected airfoil or use the Airfoil Tool at http://airfoiltools.com/search and text
search for NACA or other designations, search your aircraft, or use the library links to the left of
the screen. Once the proper airfoil is displayed and identified, select the “Airfoil details” link to
the right, which will bring up detailed plots for your airfoil similar to the ones in your textbook.
Text search input
Library links
Search result display
Airfoil details tab
This document was developed for online learning in ASCI 309.
File name: Ex_3_Lift&Airfoils
Updated: 06/23/2015
Please note the airfoil
database designator (in
parenthesis) in your
answer to 2 above.
2
Concentrate for this exercise on the Cl/alpha (coefficient of lift vs angle of attack) plot. Start by
de-cluttering the plot and leaving only the curve for the highest Reynolds-number (Re) selected
(i.e. remove all checkmarks, except the second to last, and press the “Update plots” tab).
Details Link
“Update plots”
tab
3. From the plot, find the CLmax for your airfoil (Tip: for a numerical breakdown of the plotted
curve, you can select the “Details” link and directly read the highest CL value, i.e. the highest
number within the second column, and associated AOA in the table, i.e. the associated number
in the first column):
4. Find the Stall AOA of your airfoil (i.e. the AOA associated with CLmax in 3.):
5. Find the CL value for an AOA of 5 for your selected airfoil:
6. Find the Zero-Lift AOA for your airfoil (again, the numerical table values can be used to more
precisely interpolate Zero-Lift AOA, i.e. the AOA value for which CL in the second column
becomes exactly 0):
This document was developed for online learning in ASCI 309.
File name: Ex_3_Lift&Airfoils
Updated: 06/23/2015
3
7. Compare your researched airfoil plot to the given plot of NACA 4412
(http://airfoiltools.com/airfoil/details?airfoil=naca4412-il).
a) How do the two CLmax compare to each other? Describe the differences in airfoil
characteristics (i.e. camber & thickness) between your airfoil and the given NACA 4412,
and how those differences affect CLmax. (Use your knowledge about airfoil designation
together with the airfoil drawings and details in the on-line tool to make conclusions
about characteristics.)
b) How do the two Stall AOA compare to each other? Explain how the differences in
airfoil characteristics (i.e. camber & thickness) between your airfoil and the given NACA
4412 affect Stall AOA.
in
c) How do the two Zero-Lift AOA compare to each other? Evaluate how the differences
airfoil characteristics between your airfoil and the given NACA 4412 affect Zero-Lift AOA.
8. Compare your researched airfoil plot to the NACA 0012 plot.
in
a) How do the two Zero-Lift AOA compare to each other? Evaluate how the differences
airfoil characteristics between your airfoil and the given NACA 0012 affect Zero-Lift AOA.
b) What is special about the design characteristics of NACA 0012? How and where
could this airfoil design type be utilized on your selected aircraft? Describe possible additional
uses of such airfoil in aviation.
For the second part of this assignment use your knowledge of the atmosphere and the
Density Ratio, (sigma), together with Table 2.1 and the Lift Equation, Equation 4.1, in
your textbook (remember that the presented equation already contains a conversion
factor, the 295, and speeds should be directly entered in knots; results for lift will be in
lbs):
L = CL *
* S * V2 / 295
Additionally, for your selected aircraft use the following data when applying Equation 4.1:
9. Research the Wing Span [ft]:
Use 36’ (Cessna 180k) Given
in Class
This document was developed for online learning in ASCI 309.
File name: Ex_3_Lift&Airfoils
Updated: 06/23/2015
4
10. Find the Average Chord Length [ft]:
Note: Average Chord = (Root Chord + Tip Chord) / 2
found
Use 6 ft; Given in Class
(if no Average Chord is directly
in your research)
2800 lbs; Given in Class
11. Find the Maximum Gross Weight [lbs] for your selected aircraft:
A. Calculate the Wing Area ‘S’ [ft2] based on your aircraft’s Wing Span (from 9.) and Average
Chord Length (from 10.):
S=216 sqft; Given in Class
12. Use the CL value for an AOA of 5 for your airfoil found in 5. above to simulate cruise
conditions in the following exercise B. (Note it here for easier reference):
CL=.75; Given in Class
B. Prepare and complete a table of Lift vs. Airspeed at different Pressure Altitudes utilizing the
given Lift Equation and your previous data. (For the calculation of Density Ratio ‘ ’ you can
assume standard temperatures and neglect humidity.)
You can utilize MS® Excel (ideal for repetitive application of the same formula) to populate table
fields and examine additional speeds and altitudes, but as a minimum, include six speeds (0,
40, 80, 120, 160, & 200 KTAS) at three different altitudes (Sea Level, 10000, 40000 ft), as
shown below:
Calculate LIFT (lb)
Airspeed:
0 KTAS
40 KTAS
80 KTAS
120 KTAS
160 KTAS
200 KTAS
0
Pressure Altitude (PA) ft
10,000
Example Excel Chart, version 2
in Canvas. Sigma from Table 2.1
40,000
= 1, .7385,
.2462; and using EQ
4.1
I) What is the relationship between Airspeed and Lift at a constant Pressure Altitude?
Evaluate each Altitude column of your table individually and describe how changes in
Airspeed affect the resulting Lift. Be specific and mathematically precise, and support
your answer with the relationships expressed in the Lift Equation.
II) What is the relationship between Altitude and Lift at a constant Airspeed?
Evaluate each Airspeed row of your table individually and describe how changes in
Altitude affect the resulting Lift. Be specific and mathematically precise, and support
your answer with the relationships expressed in the Lift Equation.
III) Estimate the Airspeed required to support the Maximum Gross Weight of your
selected airplane (from 11. above) at an Altitude of 10000 ft and flying at the given AOA
of 5. (As initially indicated, a more detailed table/Excel worksheet is beneficial to
precision for this task. To support the Weight of any aircraft in level flight, an equal
amount of Lift has to be generated – therefore, you can also algebraically develop the lift
This document was developed for online learning in ASCI 309.
File name: Ex_3_Lift&Airfoils
Updated: 06/23/2015
5
equation to yield a precise Airspeed result, i.e. substituting L=W and solving for V in the
lift equation. Remember that conditions in this question are not at sea level.)
C. Fill in the Example spreadsheet using the CL and AOA values from the previously used
NACA 2412 (an example is listed below).
Airspeed (KTAS)
Required Lift = Weight
Required CL
Corresponding AOA
for your airfoil
0
40
80
120
160
200
Finally, use your researched airfoil Cl/alpha plot (from 3. through 8.) to find corresponding AOA
to your calculated CL values (enter the plot in the left scale with each calculated CL value, trace
horizontally to intercept the graph for that CL value, then move down vertically to find the
corresponding AOA and note it in your table (alternatively, you can also look up values in the
detailed table):
THIS PLOT IS AN
EXAMPLE ONLY
AND NOT
APPLICABLE FOR
YOUR AIRFOIL –
PLEASE USE YOUR
RESEARCHED LIFT
CURVE FROM 3.
THROUGH 8.
ABOVE.
Enter
with CL
in the
vertical,
left
scale
Read corresponding AOA on the bottom scale
I) What is the Takeoff speed at 2800 lbs?
II) What is the stall speed at 17 degrees AOA?
(change the A/S in the 17 AOA row until L is >2800 lbs)
This document was developed for online learning in ASCI 309.
File name: Ex_3_Lift&Airfoils
Updated: 06/23/2015
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