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Aviation History
1927
1927 - 0830.PDF
SUPPLEMENT TO FLIGHT THE AIRCRAFT ENGINEER OCTOBER 27, 1927 Therefore V nia.\. = 123 m.p.h. This is close enough to the estimated Vm;ix. and Yues. that it is unnecessary to repeat the calculation. Absolute Ceiling. 'FA1'4FA =14(y) =47-6 0095 X 0-62 x 0-251 0-63P > * = -0092 x 0-745 = 0-343 From Fig. 4. Absolute ceiling = 21.500. Hate of climb at Ground. F wtY f. = 18-4 i - i = 62-0F,.' K,,= 62-6 = 0-722 0-745 x 0-722 . 0-251 =: 1.635 - 559 120 R.C. = 33.000 x 0 092 > — 3,588 X 0-62 R.C. = 1.07(5 ft. min. These two examples illustrate the procedure to be u»ed in the calculation of performance. Like all other methods, the ttft'Uracy of the final result depends upon thf estimate of parasite. The computed performance of the XCO — t> is probably somewhat high, due to the fact that Case 1 has been used. However, a slight reduction in A,, over that estimated might allow these figures to be equalled or even exceeded. Again, the percentage of motor power at ceiling may be lower than the average used. APPENDIX I PARASITE COEFFICIENTS OF VARIOUS AIRFOILS From M.I.T. Test 40 m. hr. 36 • 6 Airfoils Airfoil R.A.F. 15 U.S.A. 27 Goth 387 Goth 430 Goth 436 U.S.A. 35B ... U.S.A. 35A ... Clark W Clark X Clark Y Clark Z U.S.A. 16 N.A.C.A. 61 ... N.A.C.A. 77 ... N.A.C.A. 81 ... N.A.C.A. 58 ... Curtiss C-27 ... CurtissC-62 ... U.S.A. 17 U.S.A. 35 U.S.A. 45 Turm No. 1 ... N.A.C.A. 73 ... Goth 429 Sloane 105 Min. Prof.Drag ... 0-000025 0-0000345 ... 0-000041 0-000033 ... 0-0000313 ... 0-0000325 0-000044 ... 0-0000294 0-00002895 0-0000269 ... 0-000030 ... 0-0000229 0-0000449 ... 0-000048 ... 0-000(1348 ... 0-0000344 ... 0-0000249 — ... 0-0000278 0-0000334 ... 0-0000276 ... 0-000028 ... 0-000034 ... 0-000035 ... 0-0000232 Kv. Max.Min. Prof. Drag 104 99-8 851-3 102-5 98 102-5 85-5 99 99-8 118-2 107-4 119-5 72-9 61-3 77-3 81-1 94-8 — 95 114-5 120 90 75 64 102-5 APPENDIX II TOTAL PARASITE AREAS OF VARIOUS AIRPLANES (A,,) W A,, (A,,),,. Struc- ture 1L-1 ... 0-00428 5,686 24-35 4-6 19-75 XB1-A ... 0-00376 3,590 13-5 31 10-4 Airplane Fr Airplane PW-1 ... MB-3 ... XB-1 ... JL-6 VE-9 ... D.W.C. ... CO-5 CO-1 CO-4 Y-40 D-7 MB-2 ... PW-8* ... XB1-A ... Messenger PS-1 ... T-2 D-8 PW1-A ... PW-2 ... PW-2A ... P\Y 2B ... PW-9 ... PW-6 ... USA (11 DeH-4 ... TA-1 TA-2 TA-3 TW-1 ... PW-7 ... US-Mail ... T-3 TA-5 ... TA-6 SE-5A ... Orenco 1) TW-4 ... PW-8+ ... Morane ... F, 0 0031 0-0035 0-00387 0-00375 0 0037 0-00336 0-00306 0-0039 0-00324 0-0035 0-0035 0-00567 0-00292 0-00333 0-0064 0-00344 0-00292 0-00619 0-00375 0-00414 0 • < (038 0-00354 0-OO323 0-00374 0•00397 0-00408 0-00566 O-(H)431 0-O0636 0-00565 0-0O324 0-00384 o < 1M »44 O-00665 0-00546 0-00453 0-00536 0-OOK21 0-0025 0-00547 W 3,005 2.548 3,679 3.605 2,269 7.216 4.193 4.751 4.493 2.686 2.462 10.363 2.784 3.988 862 1.673 7.993 1.238 3.075 2.78S 2.799 2.976 2.1171 2.763 3.746 4.297 2.062 1.761 1.693 3.225 3.176 4.712 K.95O 2.215 1.954 2.060 2.820 1.967 3.151 1,458 A, 9-3 8-9 14-2 13-5 8-4 24-25 12-8 18-5 14-5 9-4 8-6 58 • 8 81 13-3 5-5 5-8 23-4 t ' i 11 • 54 11-54 10-63 10-54 9-6 10-25 14-K5 17-55 117 7-6 10-78 17-.-> 10 3 18-1 30-K 14-75 lot; !) • 33 10 12-2 7-9 7-98 (A,,), 2-05 1-9 31 (4-5) 21 7-32 4-7 (5-1) (4-5) (2-6) (2-5) (10-6) (2-0) 31 (1-8) (10-3) (1-1) (3-1) 2-4 3-35 (3-1) 1-8 2-48 3 2 3 S 8-0 3-05 2- 2 1-88 21 (2-0) (A,,) Struc- ture 7-25 700 11-1 (9-0) 6-3 16-93 8-1 (13-4) (10-0) (6-8) (6-1) (48-2) (6-1) 10-2 (3-7) (131) (6-6) (8-44) 7-20 14-20 1.8-6) 5-8 8-30 14-3 14 3 22-8 11-7 8-4 7-4.". 7-9 (5-9j Values of (A.,)„. in parentheses are estimated. * Low-compression motor. + High compression motor. THE REDUCTION OF FLIGHT TEST DATA TO STANDARD ATMOSPHERE CONDITIONS. By LIEUT-COL. .1. D. BI.YTH. O.B.E.. M.I.Ae.E. The Reduction Chart accompanying this article has been prepared with the object of providing a rapid and simple method of reducing data obtained on flight tests to conditions of standard atmosphere. As is generally known, atmospheric conditions are con- tinually varying, with the result that the observed |>erform- ances of the same aeroplane at the same aneroid height may differ considerably on different occasions. To enable com- parisons to be made on a constant basis a standard atmosphere has been adopted, to the conditions of which flight test data are reduced. There are at present two standard atmospheres : the first being that given in Chapter IX of Bairstow"s "' Applied Aerodynamics." in which relative density and pressure are taken as unity at a height of 800 ft. ; and the second one in which the relative densitv and pressure are taken as unity at a height of 0 ft., at which altitude t he temperature and pressure are 15° C. and 29-92 ins. of mercury respectively. The second standard atmosphere has l)een universally adopted in America, and has superseded, to a great extent, the first one in this country, and is therefore the one to which this article refers. The reductions given by the chart correspond to those given by the method of calculation described in N.A.C.A. Report No. 216, which may \x> des- cribed briefly. 7506
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