Need Ki84 HAYATE's data!

[Steven Que]

Hello guys.
I need some data to analyse the aircraft's ability.
Actually, that is my final exam report I have to prepare.
Then I choose Ki84 Hayate to do my report.

data in detail.
about the gravity center's location. Aerodynamic center's location
The airfoil's shape "NACA XXXX" or "NACA XXXXX"...



Thanks. I hope someone will share me these data!

[peril]

Airfoil was NN-21, work I have done in finding a NACA equivalent is based on data for NN series from Ki-27 maru book. Whilst not definitive it is the best I have been able to achieve given the few clues in text.

Root was 16.5%, tip 8% NN-21.

GoG / MAC location unknown at this time.

Estimated shape is:-

NN-21 16.5%
1.00000 0.00172
0.99679 0.00253
0.99071 0.00401
0.98314 0.00584
0.97449 0.00790
0.96538 0.01005
0.95615 0.01220
0.94683 0.01435
0.93747 0.01648
0.92812 0.01858
0.91877 0.02066
0.90939 0.02272
0.89998 0.02475
0.89057 0.02677
0.88117 0.02876
0.87178 0.03073
0.86238 0.03267
0.85295 0.03459
0.84352 0.03649
0.83409 0.03836
0.82465 0.04021
0.81520 0.04205
0.80577 0.04386
0.79638 0.04564
0.78698 0.04738
0.77755 0.04912
0.76814 0.05084
0.75871 0.05252
0.74925 0.05419
0.73981 0.05585
0.73040 0.05746
0.72089 0.05907
0.71141 0.06066
0.70198 0.06221
0.69247 0.06373
0.68294 0.06526
0.67352 0.06675
0.66416 0.06818
0.65477 0.06961
0.64537 0.07101
0.63590 0.07240
0.62641 0.07376
0.61691 0.07510
0.60749 0.07640
0.59798 0.07768
0.58850 0.07895
0.57905 0.08017
0.56952 0.08137
0.56003 0.08254
0.55052 0.08368
0.54098 0.08481
0.53154 0.08588
0.52202 0.08693
0.51254 0.08795
0.50303 0.08893
0.49348 0.08990
0.48403 0.09082
0.47453 0.09170
0.46506 0.09255
0.45559 0.09336
0.44611 0.09412
0.43660 0.09486
0.42711 0.09556
0.41767 0.09621
0.40821 0.09682
0.39873 0.09739
0.38925 0.09792
0.37985 0.09840
0.37043 0.09882
0.36097 0.09920
0.35149 0.09954
0.34213 0.09983
0.33269 0.10006
0.32330 0.10024
0.31393 0.10037
0.30457 0.10043
0.29518 0.10043
0.28585 0.10038
0.27655 0.10026
0.26722 0.10007
0.25793 0.09982
0.24866 0.09950
0.23943 0.09911
0.23019 0.09863
0.22099 0.09808
0.21177 0.09746
0.20266 0.09677
0.19369 0.09600
0.18496 0.09513
0.17634 0.09411
0.16779 0.09294
0.15925 0.09162
0.15082 0.09016
0.14246 0.08853
0.13411 0.08674
0.12585 0.08480
0.11770 0.08271
0.10964 0.08045
0.10170 0.07804
0.09386 0.07545
0.08616 0.07270
0.07858 0.06977
0.07119 0.06669
0.06401 0.06345
0.05701 0.06004
0.05030 0.05650
0.04386 0.05283
0.03780 0.04904
0.03209 0.04517
0.02688 0.04127
0.02217 0.03740
0.01806 0.03366
0.01456 0.03013
0.01169 0.02688
0.00935 0.02395
0.00749 0.02136
0.00599 0.01908
0.00481 0.01708
0.00393 0.01525
0.00319 0.01361
0.00245 0.01219
0.00187 0.01087
0.00152 0.00957
0.00125 0.00834
0.00099 0.00719
0.00051 0.00621
0.00031 0.00513
0.00023 0.00400
0.00014 0.00289
0.00008 0.00178
0.00003 0.00069
0.00001 -0.00036
0.00006 -0.00142
0.00012 -0.00250
0.00020 -0.00358
0.00027 -0.00466
0.00036 -0.00573
0.00075 -0.00667
0.00113 -0.00765
0.00137 -0.00877
0.00166 -0.00995
0.00207 -0.01114
0.00274 -0.01231
0.00345 -0.01361
0.00419 -0.01507
0.00511 -0.01662
0.00631 -0.01824
0.00780 -0.02002
0.00959 -0.02193
0.01179 -0.02397
0.01444 -0.02613
0.01762 -0.02837
0.02141 -0.03069
0.02581 -0.03303
0.03088 -0.03536
0.03657 -0.03762
0.04287 -0.03979
0.04970 -0.04185
0.05698 -0.04377
0.06462 -0.04553
0.07264 -0.04716
0.08091 -0.04866
0.08946 -0.05004
0.09817 -0.05130
0.10707 -0.05248
0.11605 -0.05356
0.12525 -0.05457
0.13453 -0.05553
0.14382 -0.05643
0.15325 -0.05727
0.16278 -0.05810
0.17235 -0.05891
0.18195 -0.05970
0.19149 -0.06047
0.20091 -0.06120
0.21018 -0.06186
0.21938 -0.06242
0.22863 -0.06291
0.23794 -0.06333
0.24722 -0.06368
0.25661 -0.06397
0.26596 -0.06420
0.27534 -0.06438
0.28471 -0.06449
0.29412 -0.06455
0.30359 -0.06457
0.31302 -0.06454
0.32245 -0.06445
0.33199 -0.06432
0.34146 -0.06416
0.35091 -0.06395
0.36044 -0.06369
0.36994 -0.06341
0.37946 -0.06308
0.38898 -0.06272
0.39854 -0.06232
0.40809 -0.06189
0.41763 -0.06143
0.42719 -0.06094
0.43676 -0.06042
0.44635 -0.05987
0.45590 -0.05930
0.46550 -0.05869
0.47511 -0.05807
0.48468 -0.05743
0.49427 -0.05676
0.50379 -0.05607
0.51337 -0.05535
0.52296 -0.05462
0.53256 -0.05387
0.54218 -0.05310
0.55179 -0.05231
0.56139 -0.05150
0.57100 -0.05068
0.58055 -0.04985
0.59010 -0.04900
0.59964 -0.04814
0.60925 -0.04725
0.61886 -0.04635
0.62849 -0.04544
0.63810 -0.04451
0.64774 -0.04357
0.65740 -0.04262
0.66704 -0.04166
0.67666 -0.04068
0.68629 -0.03970
0.69587 -0.03870
0.70543 -0.03769
0.71499 -0.03669
0.72460 -0.03566
0.73424 -0.03461
0.74388 -0.03356
0.75349 -0.03250
0.76310 -0.03142
0.77272 -0.03035
0.78226 -0.02927
0.79176 -0.02817
0.80124 -0.02709
0.81079 -0.02597
0.82042 -0.02484
0.83003 -0.02371
0.83962 -0.02255
0.84921 -0.02140
0.85877 -0.02023
0.86836 -0.01905
0.87795 -0.01787
0.88755 -0.01666
0.89718 -0.01545
0.90681 -0.01422
0.91641 -0.01299
0.92602 -0.01174
0.93563 -0.01048
0.94524 -0.00921
0.95478 -0.00794
0.96423 -0.00667
0.97360 -0.00539
0.98257 -0.00415
0.99042 -0.00307
0.99670 -0.00220
1.00000 -0.00174

[Shinpachi]

Hi peril!
The coordinates you have provided are applied on my CAD software.
Thanks for such a fantastic data.

[HoHun]

Hi Peril,

Have you managed to genrate a polar from the coordinates? I've been toying around with Javafoil, but without results :-/

Regards,

Henning (HoHun)

[peril]

I set Javafoil polars to transition model 'xfoil after 1991'.

I usually have no problems getting results, ensure closed trailing edge.

If still having issues play with reducing max number of plotted points.

[Steven Que]

HI Peril,
I wonder the coefficient of lift with NN-21 when AOA is 0 degree
I can not find out the data from my program
which have more than 300 airfoil's data
But have no data of NN-21...

Thanks