Transcript DESIGN OF AN ASSEMBLY STATION FOR INTERNAL …

MIDDLE EAST TECHNICAL UNIVERSITY
AE 462 – Aerospace Structures Design
Minimum Weight Wing Design for a
Utility Type Aircraft
DESIGN TEAM : Osman Erdem ŞENOL
Hasan YILMAZ
OUTLINE
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Design Objective
Procedure
Results
Conclusion
Design Objective
Design a suitable primary structure of least
possible weight for the uniform cantilever wing
of an utility type aircraft
• Federal Aviation Regulations part 23
• An overall safety factor of 1.5
Procedure
• According to FAR 23 Appendix A with all the specified
regulations the flight envelope for the aircraft is
constructed.
• For the Dive Speed and Minimum Maneuvering
Speed points on the flight envelope spanwise Local
lift coefficent and pitching moment coefficient
variation is determined by the use of ESDU code
9510v12.
• From the output of the code, internal load variaton
along wing span is found tabulated and figured.
Procedure
Sectional Shear Force Variaton at Minimum Maneuvering
Speed Va
35000.000
30000.000
25000.000
20000.000
15000.000
10000.000
5000.000
0.000
0.0
50.0
100.0
150.0
200.0
250.0
300.0
350.0
400.0
450.0
500.0
Procedure
Bending Moment Variation Along Wing Span at Maneuvering
Speed Vd
70000.000
60000.000
50000.000
40000.000
30000.000
20000.000
10000.000
0.000
0.0
50.0
100.0
150.0
200.0
250.0
300.0
350.0
400.0
450.0
500.0
Procedure
Pitching Moment Variation Along Wing Span at Minimum
Maneuvering Speed Vd
0.00
0.0
-500.00
-1000.00
-1500.00
-2000.00
-2500.00
-3000.00
-3500.00
-4000.00
-4500.00
-5000.00
50.0
100.0
150.0
200.0
250.0
300.0
350.0
400.0
450.0
500.0
Procedure
With these calculated sectional loads wing
structure is to be formed.
The overall sizing of the internal wing
structure is done through the principle of
minimizing the margins of safety in order to
achieve minimum weight design.
Procedure
Material is
Al 2024
• Widely used in aerospace structures.
• Good strength characteristics comparing with
low density.
• Available in variable thicknesses of sheet.
Procedure
MS Excel is used for the calculations.
• Initial values for spar cap areas, web and skin
thicknesses are assigned.
• Main assumption on structural loading:
– Shear forces are carried only by skins and webs
– Axial loads are carried only by spar caps
• Bending stresses are calculated from unsymetric
bending formula.
• Spar flange areas are resized in order to achieve
positive margins of safety close to zero for axial
stresses.
Procedure
• Sectional shear flows are calculated with the
help of defined intermediate variables and
margins of safety for shear stresses are found.
• In order to aproach margins of safety to zero,
skin and web thickness values are edited.
– Thickness values becomes very small for this
requirement. Total weight was very low at the
moment.
Procedure
• After obtaining t and A values, shear center of
the airfoil at different sections are calculated.
• Additional torque due to the position of the
shear center is calculated to be small, no
further iterations are developed.
Procedure
• Spar webs, skins and ribs are to be checked for
shear buckling.
• All the skins are considered as flat plates.
– For this purpose leading edge is considered to
have a stinger and the skin is divided into two
straight lines.
– Shear flow on this two plates are considered to be
same as the curved plate.
Procedure
• The airfoil is designed with 9 ribs for semi-span
– The purpose for the selection of 9 ribs
• As the buckling is considered to be more critical, by
decreasing panel lengths, a/b ratios can be kept smaller
especially for the leading edge skins and webs. As a result
higher Ks and so higher critical buckling stresses.
• Possibility to change flange areas and skin and web
thicknesses for 10 different sections.
• Ease of calculation.
• Since there are 20 discrete sections for the
internal loads for buckling stresses of 10 sections,
the average loads are taken from the load data in
between
Procedure
Procedure
• The buckling margins of safety come out to be
negative close to 1, that shows the thicknesses
for the panels are required to be increased
drastically.
• The thicknesses are changed in order to have
margins of safety close to 0 and positive.
Procedure
• As the ribs have a very complex geometry it is
difficult to model them for buckling
• The rib is assumed to be a rectangular plate with
b = front web and a = top panel
• The shear stress acting on the wing ribs are the
difference between the neighboor hood panels
from both sides.
• The average of the maximum shear flow couple is
used for the buckling of the ribs.
• Rib thicknesses are also determined for the
minimum margins of safety.
Procedure
Selection from the standard materials:
• Throughout the iterations, skins, webs and rib
thicknesses are edited as differences of at
least 0.1mm.
• Minimum spar cap area is set to be 1 cm2.
• From Table A3.15 of Bruhn Spar flange shape
is selected as I beam model. It is available in
different areas and dimensions.
Procedure
• Now that all the margins of safety are larger than
0 considering all the fail modes, weight of the
wing is to be calculated.
• Wing Structure consists of
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Spar Caps: 4 discrete each section for 10 sections
Skins: top, front and bottom of 10 discrete sections
Spar Webs: front and rear spar 10 discrete sections
Ribs: a total of 9 between every section.
Results
Same procedure from the start is applied
for both the load cases, point A and D.
– The design weight for point A = 81.136
– The design weight for point D = 85.502 kg
Thus Point D, dive speed loading case is
more critical.
Results
All the the wing elements’ dimensions are
tabulated in the design report for each
section.
The general trend for all dimensions are as
expected; larger at wing root and becoming
smaller as advanced to wing tip.
Conclusion
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What is achieved:
How to construct a filght envelope is learned.
Difference in flight loading at different flight
conditions are seen.
The effects of the wing structural elements’
dimensions on different failure modes are
observed.
The importance of design for buckling type of
failure become evident.
Thank you for listening
Questions?