Post on 02-Jun-2018
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w0 74442.28823
Warmup and takeoff w1/w0 0.97
Climb w2/w1 0.985
Cruise R (ft) 7898950
C (/ft) 2.52525E-07 *coefficienL/D 15 *fraction is
Propeller effieci 0.85 *pg 45 sha
V (ft/s) 654.735 *using repr
w3/w2 0.855177841
Loiter(30 minute) E(s) 1800
C (/ft) 3.03E-07
L/D 15
Propeller effieci 0.85
V (ft/s) 286.3 *1.4vstall
w4/w3 0.985956212
Land w5/w4 0.995
mission weight ratio w5/w0 0.80157675
fuel weight ratio wf/w0 0.210328645
empty weight ratio we/w0 0.54787344 *raymer p
wpax (80*225lb) 18000
w0 iterate (lb) 74442.32927
Wing loadingconstraint by vstall
Clmax 1.924*initial
row 0.002377
Vstall 204.5
W/S 95.62928
constraint by landing variable *PG 411 P.D.A
vf 251.535 a 22.45151067
R 982.4512 b 72.14678256
hf 1.346415 c -3020.216776
sa 928.3657
sf 51.41752 W/S 102.0582528
sg 3020.217 W/S 177.3115252
wing area
w/s 95.62928 *lowest
S 778.447
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Wf 15657.35
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in english unit Overall length
4000 ft 4000 Height (to top of horizontal tai
0.65M 387.92 kts 654.735 Fuselage diameter
3500 4000 * Maximum cabin width
1600 ft/min 26.66667 Cabin length1300 nm 7898950 Wingspan (geometric)
80 Wing area (reference)
27000 ft 2700 Basic operating data
18225 lb 18225 Engines
Typical passenger seating
121.162937 kts 204.5*ade link d 6 Passenger seating range
Typical cruise speed
it: /ft) 2.02E-07 Maximum operating altitude
0.85 Range (w/typical pax)
Range (w/LR tanks)
70 feet Takeoff run at MTOW
1600 f/min Design weights
Maximum takeoff weight
Maximum landing weight
Maximum zero fuel weight
Maximum fuel capacity
Typical operating weight empt
Typical volumetric payload
HIGH LIFT DEVICE
leading edge check
TAKE OFF WING LOADING
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877.1486
sweep
1.2
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q400
107 ft 8 in (32.81 m)
27 ft 3 in (8.3 m)
61 ft 8 in (18.8 m)93 ft 2 in (28.4 m)
679.20 ft (63.1 m)
2 PW150A
78 (Single Class)
6886[55]
414 mph (667 km/h) 360 knots
27,000 ft (8,230 m)
1,567 miles (2,522 km)
n/a
4,600 ft (1,402 m)
64,500 lb (29,260 kg)
61,750 lb (28,010 kg)
57,000 lb (25,850 kg)
1,748 US gal (6,616 L)
37,886 lb (17,185 kg)
19,11
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Stall Takeoff
density at sea level (sl 0.002377 flight path
Clmax 1.924 flight path
Vstall (ft/s) 204.5 airborne di
Sg (ft)
S TO (ft)
W/S 95.62928 density at 4
friction coeW/S W/P Vto
95.62928 0 CLc
95.62928 2.5 del CL flap
95.62928 5 CL to
95.62928 7.5 CDoLG
95.62928 10 CDoHLD
95.62928 12.5 CDoTO
95.62928 15 Cdto
95.62928 17.5 CLR
95.62928 20 CDg
95.62928 22.595.62928 25 W/S
0
5
Cruise 10
propeller efficiency 0.85 15
density at 27000ft (slu 9.93E-04 20
CD0 0.025 for twin engine prop (m. sadraey table 4.12) 25
AR 11 m sadreay table 5.8 30
e 0.85 35
relative density 4.18E-01 40
K 0.034044 45
Vcruise (ft/s) 654.735 5055
W/S W/P 60
0 #DIV/0! 65
5 0.280085 70
10 0.55891 75
15 0.835237 80
20 1.10786 85
25 1.375631 90
30 1.637471 95
35 1.892381 100
40 2.1394645 2.377908
50 2.60703
55 2.826248
60 3.035091
65 3.233202
70 3.420329
75 3.596323
80 3.761129 25
26
27
28
29
30
31
32
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85 3.914777
90 4.05738
95 4.189115
100 4.310225
design point
W/S 94W/P 4
0
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
17
18
19
20
21
22
23
0 2 4 6 8 10 12 14 16
W/P(lb/hp)
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Rate of climb
adius R 9039.408075 ROC 26.66667
ngle 0.088027679 L/D 15
tance Sa (f 794.6908614
4000 W/S W/P
4794.690861 0 17.53125
795 ft 1.96E-03 5 15.19531
fficient 0.04 10 14.40052224.95 15 13.84486
0.3 20 13.40868
0.6 25 13.04656
0.9 30 12.73561
0.009 35 12.46246
0.005 40 12.21854
0.039 45 11.99799
0.066575509 50 11.79659
1.590082645 55 11.6112
0.030575509 60 11.43943
65 11.27939W/P 70 11.12957
#DIV/0! 75 10.98873
30.27504 80 10.85587
24.43005 85 10.73012
20.36313 90 10.61078
17.43562 95 10.49723
15.23747 100 10.38895
13.52892
12.16368
11.04811
10.11962
9.3348898.662974
8.081207
7.572603
7.124188
6.725882
6.369735
6.049393
5.759717
5.496507
5.256295
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8 20 22 24 26 28 30 32 34 36 38 40 42 44 46 48 50 52 54 56 58 60 62 64 66 68 70 72 74 76 78 80 82 84 86 88 90 92 94 96 981001021
W/S (lb/ft^2)
Design Point:
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04106
Stall
Cruise
Takeoff
Rate of climb
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Vstall (ft/s) 204.5
Dihedral 2 deg
sweep (LE) 1.2
Sweep (TE) 10.9
Wing setting anlge 0 from naca 4418 Cla curve
sweep (0.25c) -1.858
S (ft^2) 791.9396731
S (m^2) 73.5736
b (m) 28
AR 10.65599617
c root (m) 4
c tip (m) 1.3
taper ratio 0.325
mean aero chord (m) 2.879245283 9.446228 (ft)
y mac (m) 5.811320755 19.06578 (ft)
rho at 27000 ft (slug/ft^ 9.93E-04
cruise speed (ft/s) 654.735
viscosity (lb/ft s) 1.06E-05
Reynold at cruise 5.78E+05
viscosity at sea level (lb/ 1.23E-05
density at sea level 2.38E-03
takeoff speed (ft/s) 224.95
Reynold during takeoff 1.25E+05
cla (/deg) 0.0994
cla (/rad) 5.695200484
CL alpha (/rad) 4.867177693
h
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W/s 95.62928
Vto takeoff vel
CL to #DIV/0!
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W0/S 94
W3/W2 0.855178
W2/W1 0.985
W1/W0 0.97
W2/W0 0.95545
Waverage/W2 0.927589
Waverage/S 83.30889
CLc 0.391418 ideal cruise lift coefficient
CLcw 0.412019 wing cruise lift coeffcient
Cli 0.457799 airfoil ideal lift coeffcient
Clmax 1.924
Clmax w 2.025263 wing maximum lift coeffcient
Clmax gross 2.250292 airfoil gross maximum lift coeffcient
use plain flap
delta ClHLD 0.8 lift coeffcient increment for plain flap deflecteClmax 1.450292 net maximum lift coeffcient
airfoil (based on sadreay fig 5.23) naca 4415 use as wingtip (smaller thickness to prevent st
naca 4418 root
naca 4421
naca 4412
Naca 4418
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naca 4418
d 60 deg
all at tip)
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XFLR5 v6.09.01 beta
Calculated polar for: NACA 4418
1 1 Reynolds number fixed Mach number fixed
xtrf = 1 (top) 1 (bottom)Mach = 0 Re = 0.578 e 6 Ncrit =
alpha CL CD CDp Cm Top Xtr Bot Xtr Cpmin
------- -------- --------- --------- -------- ------- ------- -------- --------- ---------
-16 -0.8917 0.05671 0.05283 -0.0979 1 0.0359 -6.5071 0 0.1088
-15 -0.9661 0.04065 0.03634 -0.101 1 0.0362 -6.6691 0 0.1184
-14 -0.9909 0.02827 0.02335 -0.1051 0.9914 0.0387 -6.5971 0 0.12
-13 -0.8952 0.0239 0.01874 -0.1105 0.9801 0.0423 -5.9586 0 0.1036
-12 -0.7851 0.02116 0.01576 -0.1139 0.9669 0.0459 -5.2765 0 0.0824
-10 -0.5946 0.01625 0.01042 -0.1112 0.9253 0.0566 -4.1357 0 0.041
-9 -0.4998 0.01449 0.00843 -0.1084 0.8982 0.0632 -3.6475 0 0.0107
-8 -0.3958 0.01346 0.00727 -0.107 0.8712 0.072 -3.1414 0 -0.044
-7 -0.2925 0.01224 0.00593 -0.1054 0.8394 0.085 -2.6681 0 -0.1369
-6 -0.1854 0.01138 0.00494 -0.1043 0.8059 0.1012 -2.2368 0 -0.3469
-5 -0.0777 0.0107 0.00414 -0.1032 0.7687 0.1246 -1.8518 0 -1.143
-4 0.0306 0.0102 0.00362 -0.1023 0.7291 0.1554 -1.4892 0 3.7179
-3 0.1377 0.00988 0.00321 -0.101 0.6848 0.1965 -1.1868 0 1.0026
-2 0.2453 0.0096 0.003 -0.1 0.6436 0.2518 -0.9154 0 0.6633
-1 0.3511 0.00949 0.00291 -0.0987 0.6016 0.3246 -0.8994 0 0.5314
0 0.4549 0.00923 0.00296 -0.0972 0.5646 0.4653 -0.9877 0 0.4609
1 0.5564 0.00892 0.00323 -0.095 0.5349 0.6904 -1.0814 0 0.4163 0.7801 0.00941 0.00419 -0.0937 0.4881 0.9808 -1.3079 0 0.3623
4 0.9308 0.01004 0.00469 -0.1024 0.4661 1 -1.4663 0 0.351
5 1.0077 0.01056 0.00509 -0.0956 0.446 1 -1.5855 0 0.3345
6 1.0915 0.01125 0.00567 -0.0904 0.4235 1 -1.7276 0 0.3209
7 1.1759 0.01184 0.00628 -0.0853 0.4034 1 -1.8988 0 0.3093
8 1.259 0.01275 0.00717 -0.0806 0.3783 1 -2.1112 0 0.2992
9 1.337 0.01403 0.00841 -0.0755 0.3492 1 -2.3654 0 0.2901
10 1.4058 0.0159 0.01019 -0.0696 0.3159 1 -2.6479 0 0.2815
11 1.4727 0.0181 0.01241 -0.0643 0.283 1 -2.9616 0 0.2738
12 1.5209 0.02159 0.01581 -0.0576 0.2456 1 -3.2765 0 0.2661
13 1.5611 0.02608 0.02031 -0.0514 0.2108 1 -3.5886 0 0.259214 1.5791 0.03283 0.027 -0.0451 0.1747 1 -3.8626 0 0.2526
15 1.5889 0.0413 0.0355 -0.0404 0.1454 1 -4.1415 0 0.2471
16 1.5779 0.05273 0.04702 -0.037 0.1208 1 -4.3584 0 0.2427
17 1.5622 0.06562 0.06006 -0.0355 0.1016 1 -4.5538 0 0.2394
18 1.5376 0.08026 0.07487 -0.0356 0.0863 1 -4.7113 0 0.2372
19 1.5122 0.09557 0.09037 -0.0371 0.075 1 -4.8538 0 0.236
20 1.4974 0.10984 0.10491 -0.0396 0.0666 1 -5.0505 0 0.2353
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cl/cd max
cl/cd max 62.3767
-1.2-1.1
-1-0.9-0.8-0.7
-0.6-0.5-0.4-0.3-0.2-0.1
00.10.20.30.40.50.60.70.8
0.91
1.11.21.31.41.51.61.71.8
0 0.00250.0050.00750.01 0.01250.0150.0175 0.02 0.02250.0250.0275 0.03 0.03250.0350.0375 0.04 0.04250.0450.0475 0.05 0.0525
-20 -
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cm0 -0.0975
-10
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9
Chinge XCp
stall angle 15 deg
stall lift clmax 1.6
zero lift alpha -4.5
cla (/deg) 0.0994
cla (/rad) 5.6952
-1.2-1.1
-1-0.9-0.8-0.7-0.6-0.5-0.4-0.3-0.2-0.1
00.10.20.30.40.50.60.70.80.9
11.11.21.31.41.51.61.71.8
-18 -17 -16 -15 -14 -13 -12 -11 -10 -9 -8 -7 -6 -5 -4 -3 -2 -1 0 1 2 3 4 5 6 7 8 9 10
Cl alpha
0.60.70.80.9
11.11.21.31.41.51.61.71.81.9
Cl alpha
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-1.1-1
-0.9-0.8-0.7-0.6
-0.5-0.4-0.3-0.2-0.1
00.10.20.30.4
.
-15 -14 -13 -12 -11 -10 -9 -8 -7 -6 -5 -4 -3 -2 -1 0 1 2 3 4
.0550.0575 0.06 0.06250.0650.0675 0.07 0.07250.0750.07750.08 0.08250.0850.0875 0.09 0.09250.0950.0975 0.1 0.10250.1050.1075
Cl vs cd
y = 1E-04x2+ 0.0018x - 0.0987
y = 0.0022x - 0.0881
-0.095
-0.09
-0.085
-0.08
-0.075
-0.07
-0.065
-0.06
-0.055
-0.05-0.045
-0.04
-0.035
-0.03
-0.025
-0.02
-0.015
-0.01
-0.005
0
15 -10 -5 0 5 10 15 20
Cm vs alpha
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-0.13
-0.125
-0.12
-0.115
-0.11
-0.105
- .
-8 -6 -4
Cl vs cd
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1 12 13 14 15 16 17 18 19 20 21 22
y = 0.0994x + 0.4279
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5 6 7 8 9 10 11 12 13 14 15
0.11 0.11250.1150.1175
25
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cma (/deg) 0.0012
cma (/rad) 0.068755
y = 0.0012x - 0.0972
-0.108
-0.106
-0.104
-0.102
-0.1
-0.098
-0.096
-0.094
-0.092
-2 0 2
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4
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XFLR5 v6.09.01 beta
Calculated polar for: NACA 4418
1 1 Reynolds number fixed Mach number fixed
xtrf = 1 (top) 1 (bottom)Mach = 0 Re = 0.578 e 6 Ncrit =
alpha CL CD CDp Cm Top Xtr Bot Xtr Cpmin
------- -------- --------- --------- -------- ------- ------- -------- --------- ---------
-16 -0.8917 0.05671 0.05283 -0.0979 1 0.0359 -6.5071 0 0.1088
-15 -0.9661 0.04065 0.03634 -0.101 1 0.0362 -6.6691 0 0.1184
-14 -0.9909 0.02827 0.02335 -0.1051 0.9914 0.0387 -6.5971 0 0.12
-13 -0.8952 0.0239 0.01874 -0.1105 0.9801 0.0423 -5.9586 0 0.1036
-12 -0.7851 0.02116 0.01576 -0.1139 0.9669 0.0459 -5.2765 0 0.0824
-10 -0.5946 0.01625 0.01042 -0.1112 0.9253 0.0566 -4.1357 0 0.041
-9 -0.4998 0.01449 0.00843 -0.1084 0.8982 0.0632 -3.6475 0 0.0107
-8 -0.3958 0.01346 0.00727 -0.107 0.8712 0.072 -3.1414 0 -0.044
-7 -0.2925 0.01224 0.00593 -0.1054 0.8394 0.085 -2.6681 0 -0.1369
-6 -0.1854 0.01138 0.00494 -0.1043 0.8059 0.1012 -2.2368 0 -0.3469
-5 -0.0777 0.0107 0.00414 -0.1032 0.7687 0.1246 -1.8518 0 -1.143
-4 0.0306 0.0102 0.00362 -0.1023 0.7291 0.1554 -1.4892 0 3.7179
-3 0.1377 0.00988 0.00321 -0.101 0.6848 0.1965 -1.1868 0 1.0026
-2 0.2453 0.0096 0.003 -0.1 0.6436 0.2518 -0.9154 0 0.6633
-1 0.3511 0.00949 0.00291 -0.0987 0.6016 0.3246 -0.8994 0 0.5314
0 0.4549 0.00923 0.00296 -0.0972 0.5646 0.4653 -0.9877 0 0.4609
1 0.5564 0.00892 0.00323 -0.095 0.5349 0.6904 -1.0814 0 0.4163 0.7801 0.00941 0.00419 -0.0937 0.4881 0.9808 -1.3079 0 0.3623
4 0.9308 0.01004 0.00469 -0.1024 0.4661 1 -1.4663 0 0.351
5 1.0077 0.01056 0.00509 -0.0956 0.446 1 -1.5855 0 0.3345
6 1.0915 0.01125 0.00567 -0.0904 0.4235 1 -1.7276 0 0.3209
7 1.1759 0.01184 0.00628 -0.0853 0.4034 1 -1.8988 0 0.3093
8 1.259 0.01275 0.00717 -0.0806 0.3783 1 -2.1112 0 0.2992
9 1.337 0.01403 0.00841 -0.0755 0.3492 1 -2.3654 0 0.2901
10 1.4058 0.0159 0.01019 -0.0696 0.3159 1 -2.6479 0 0.2815
11 1.4727 0.0181 0.01241 -0.0643 0.283 1 -2.9616 0 0.2738
12 1.5209 0.02159 0.01581 -0.0576 0.2456 1 -3.2765 0 0.2661
13 1.5611 0.02608 0.02031 -0.0514 0.2108 1 -3.5886 0 0.259214 1.5791 0.03283 0.027 -0.0451 0.1747 1 -3.8626 0 0.2526
15 1.5889 0.0413 0.0355 -0.0404 0.1454 1 -4.1415 0 0.2471
16 1.5779 0.05273 0.04702 -0.037 0.1208 1 -4.3584 0 0.2427
17 1.5622 0.06562 0.06006 -0.0355 0.1016 1 -4.5538 0 0.2394
18 1.5376 0.08026 0.07487 -0.0356 0.0863 1 -4.7113 0 0.2372
19 1.5122 0.09557 0.09037 -0.0371 0.075 1 -4.8538 0 0.236
20 1.4974 0.10984 0.10491 -0.0396 0.0666 1 -5.0505 0 0.2353
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
0 0.2 0.4
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9
Chinge XCp
0.6 0.8 1 1.2
FLR5 v6.09.01
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wing
canard S (m^2) 73.5736
c root (m) 1.07 b (m) 28
c tip (m) 1 AR 10.65599617
sweep le 1.2 sweep (0.25c) -1.858
span (m) 6.56 c root (m) 4
taper 0.934579 c tip (m) 1.3mac (m) 1.035395 taper ratio 0.325
y mac (m) 1.621514 mean aero chor 2.879245283
area (m2) 6.7896 y mac (m) 5.811320755
AR 6.338164 cla (/deg) 0.0994
t/c max 0.18
CL alpha 4.428549 cg from LE(m) 1.519
sweep (0.25c) 0.9 weight wing (N) 3780.01
cg froom LE (m) 0.552
weight canard (kg) 173.55
horizontal tail cla (/rad) 5.695200484c root (m) 2.27 CL alpha w(/rad 4.867177693
c tip (m) 1.42 alpha zero lift (d -4.5
sweep le 30 CL0 w -0.382267242
span (m) 8.36 x from .25ac to -1.420424312
taper 0.625551 Cmo w 0.188584727
mac (m) 1.877633 depsilon/dalpha 0.290779154
y mac (m) 1.929521 epsilon 0 -0.022837741
area (m2) 15.4242
AR 4.531165 xacw 15.71975
cg with wing (m) 11.63 xcg 15.5
CL alpha 4.067761
sweep (0.25c) 26.767cg from LE (m) 2.053
weght h tail (kg) 2110.35
l canard 10.8712
V canard 0.348434
ltail 10.83975
V tail 0.789262
Cm0 0.080023cm alpha -1.55408
cm alpha
0.080023497 0
-1.474054964 1
1.634101958 -1
-1.5 -1Cm
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xcg/c 0.866927
xnp/c 0.688773
SM -0.17815
xnp 12.31474
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-0.5
0
0.5
1
1.5
2
-0.5 0 0.5 1
cm vs alpha
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-2
-1.5
-
alpha (deg)
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1.5
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vertical tail wing
S (m^2) 73.5736
c root (m) 4 b (m) 28
c tip (m) 4 AR 10.656
sweep le 40 sweep (0.2 -1.858
span (m) 4.7 c root (m) 4taper 1 c tip (m) 1.3
mac (m) 4 taper ratio 0.325
y mac (m) 1.175 mean aero 2.879245
area (m2) 18.8 y mac (m) 5.811321
AR 1.175 cla (/deg) 0.0994
t/c max 0.18
CL alpha #VALUE! cg from LE( 1.519
sweep (0.2 40 weight win 3780.01
cg from LE 3.3
weght h tai 1744.44
l vtail 4.28
V tail 0.039059
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W (lb) W(kg) cg moment ar moment
passenger 6240 11.5 71760
pilot 160 0.25 40
landing gear 0
wing 3780.01 16.519 62441.9852
canard 173.55 1.052 182.5746
horizontal tail 2110.35 23.053 48649.8986vertical tail 1744.44 25.3 44134.332
fuel 15657.35 7102 17 120734
engine 965.245 17 16409.165
baggage 1764.8 8 14118.4
0
fuselage 1446.014851 11.5 16629.1708
sum 25486.40985 395099.526
mac 2.879245283
c 17.87924528 xnp/c 0.68877288
Column1 Column2 Column3 Column4
SM Weight (kg)xcg full (m) 15.50 -0.18 25486.41
xcg without fuel (m) 14.92 -0.15 18384.41
xcg without payload (with fu 17.85 -0.31 17321.61
xcg empty (m) 18.44 -0.34 10219.61
xcg mid cruise (m) 15.28 -0.17 22148.47
xcg end of cruise (m) 15.48 -0.18 25060.29
0.00
5000.00
10000.00
15000.00
20000.00
25000.00
30000.00
0.00 0.20 0.40 0.60
weight(kg)
% mac
CG variation
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Column5
xcg/c 0.87
0.83
1.00
1.03
0.85
0.87
0.80 1.00 1.20
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Lf (m) 23
Dfmax (m) 3.18
rho materi 2711
Krhof 0.0028n 0.45
g 9.81
weight fus 1446.015
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