A201_PM6.docx
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School
Purdue University *
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Course
251
Subject
English
Date
Apr 3, 2024
Type
Pages
10
Uploaded by CaptainComputerEel9
Team Name
Team Number
RAAAW
A201
AAE 251: Introduction to Aerospace Design
Project Milestone 6
Due Thursday October 26
th
, 11:59 PM ET on Brightspace
Instructions
Answer each question in this Project Milestone assignment as a team and record your team’s response in the “Team Response” box
under each question. Complete each question in full sentences.
Leave the “Peer Review Comments” box empty when your team submits your Project Milestone assignment. You will use these
boxes for the three Peer Reviews in the semester.
Upload your google doc link on Brightspace by the due date specified in this assignment. Only one team member needs to submit
on behalf of the entire team.
Refueling Requirement Update (Important)
After a discussion among the teaching team, we've decided to relax one of the requirements for the aircraft project. As of now,
we will no longer require that the refueling aircraft refuel all 4 scanning aircraft in the same flight. You can now choose to have
it only refuel 1, 2, or 3 scanning aircraft per flight (for example, you could have it refuel scanners 1 and 2, land/refuel itself,
then fly out to refuel scanners 3 and 4), as long as you take precautions to prevent all 4 scanners from running out of fuel at the
1
same time (such as by staggering launch times). This should help you reduce the size of your scanning aircraft if you choose to
take this approach.
2
Wing Planform
Wing loading helps you size your wing. Based on your wing loading estimation from PM5, you should have an idea of how large
your wing should be. Estimate the wingspan and chord your wing should have. Research different wing planforms and select
one that is appropriate for your design. Insert a sketch or image of the selected planform. Explain your reasoning. Refer to
Nicolai Chapter 7 for some helpful guidance.
Team Response
Peer Review Comments
3
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Wing loading is 26.9 pounds per square feet. Aircraft weight is the maximum of refueling and
scanning vehicles, in our case 515 lb (subject to change based on fine tuning of sizing code).
So surface area is weight divided by loading, or 19.14 square feet.
According to Nicolai Ch 7, a taper ratio around 0.35 yields a nearly elliptical lift distribution,
meaning the lift induced drag is minimized. To reduce build complexity, however, we will go
with rectangular wings.
Since ours is a relatively low speed aircraft, a low aspect ratio will suffice. The AR of the
Cessna 172 is around 7, while for the Cirrus SR22 it is 10, just for reference. Too high an
aspect ratio comes with considerable extra weight, while choosing one that is too low will
affect drag as well as lateral stability and aileron performance. For now, we are going with 7.
With this value, the wingspan becomes:
? =
𝐴? × ?
=
7
×
19. 14 ??
2
=
11. 57 ??
Thus the chord is:
1.65 ft
? = ?/? =
Calculating coefficient of lift for later reference. Surface area S is 19.14 square feet. Cruise
velocity is 100 ktas or 168.78 ft/s. Lift equals weight, which is 515 pounds. Density depends
on cruise altitude, which we have not determined yet. For now, let us take the standard
atmosphere value for density at 5000 ft (our aircraft will probably cruise much lower than
this, but high temperature will lower density anyways). So, take density as .0020482
slug/ft^3. Thus the minimum lift coefficient to maintain level flight is:
0.92
𝐶
𝐿
=
(2)(515 𝑙?)
(0.0020482 ?𝑙??/??
3
)(168.78 ??/?)
2
(19.14 ??
2
)
=
Planform sketch:
4
Airfoil Characteristics
To determine which NACA series of airfoils is appropriate for your aircraft, you need to estimate airfoil characteristics such as
the maximum thickness to chord ratio, location of this ratio, and camber of the airfoil. Nicolai, Chapter 7 lays out the effects of
these characteristics on aircraft performance. Based on your requirements, estimate these airfoil characteristics. You will likely
need to make a couple of educated guesses, which is reasonable at this stage of the design process.
Team Response
Peer Review Comments
Ideally, we want an aircraft with the greatest amount of lift for the lowest amount of a drag at
a given velocity. To achieve the closest to ideal scenario the characteristics that need to be
considered are the maximum thickness to chord ratio (t/c), location of this ratio, and the
camber of the airfoil. From Nicolai Chapter 7, we know that a higher t/c ratio leads to a higher
maximum lift coefficient. Additionally, positive camber gives an increase in section maximum
lift coefficient;it can also lead to an increase in coefficient of drag. The location of the t/c ratio
is what determines the end of the favorable, decreasing, pressure gradient and the adverse
pressure gradient. Furthermore, if the location is far forward, then this can lead to a nose-up
pitching moment (the opposite results in a nose down moment). The ideal maximum thickness
to chord ratio 0.273 and at a location of 0.3. By locating the maximum thickness to chord ratio
a little towards the aft it will allow for more stability and control in flight without pointing the
nose too far down. Ideally, the camber would be more towards the front as well.
5
Airfoil Selection
Research existing airfoils. The table of airfoil families in the Drag Polar lecture slides provides a good starting point. Nicolai
Appendix F also has some useful information on airfoil selection. Using the airfoil characteristics you identified above, select an
appropriate airfoil for your mission. Explain your reasoning.
Team Response
Peer Review Comments
6
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ROHAN
Efficient airfoils typically have the maximum thickness around one-third of the way back from
the leading edge of the wing. Thus, the team decided to select an airfoil based on a
combination of good max thickness positioning and high coefficient of lift.
Two options quickly presented themselves: the NACA 63
4
-021 and the NACA 65(421)-420.
The thickness positioning ratios are .21 and .20 respectively, some of the highest of all the
NACA profiles. However, the 65(421)-420 has a higher maximum C
L
of 1.552, while the
63
4
-021 has a C
L
of 1.38. Thus, it was judged that the former would be the selection for the
mission.
Airfoil Geometry
Insert an image of the airfoil you selected. On your airfoil, show the chord length, thickness, and camber line. Plot the airfoil
profile using MATLAB,
http://airfoiltools.com/
, xfoil, or other computational tool of your choice.
Team Response
Peer Review Comments
7
NACA 65-421
Lift and Drag
Insert airfoil charts (plots of
vs
and
vs
) for your airfoil. Indicate the maximum lift coefficient, stall angle, coefficient of
𝐶
𝑙
α
𝐶
𝑙
𝐶
?
parasitic drag, and maximum L/D ratio for the wing. Use the tools suggested in the previous section.
8
Team Response
Peer Review Comments
Clmax = 1.25
Stall angle = 19 degrees
Cd = 0.10
L/D = 120
9
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Questions
If you have questions about the project, please include them below. As always, for questions that likely require longer
discussions, ask during class sessions or use the online help sessions. We’re here to help! If you are waiting for answers from
previous PMs, please repeat them here and also email your instructor.
N/A
10