NASA-CR-2334-1974 Correlation of full-scale drag predictions with flight measurements on the C-141A aircraft Phase 2 Wind tunnel tests analysis and prediction techniques Volume 2 .pdf

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1、Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-FLIGHT MEASUREME TEST, ANALYSIS ANI) BASIC DATA la G. MacWilkinson, W. T. Blackerby and J. H. Paterson tion Name and Address Lockheed-Georgia Company flaiiariettrz, Georgia Vational Aeronautics and Spac

2、e Administration dashington, D. C. 8. Performing Organization Report No. LG73ER0058 10. Work Unit No. 501-06-09-01 11. Contract or Grant No. NAS1-10045 13. Type of Report and Period Covered Contractor Report 14. Sponsoring Agency Code I This is one of two final reports. A research program has been c

3、onducted to determine the degree of cruise drag correlation on the Volume C-lhlA aircraft between predictions based on wind-tunnel test data, and flight test results. 2 contains information on the wind-tunnel test program and basic aerodynamic data on the C-14U wind- tunnel model used in the correla

4、tion studies described in Volume 1. The model was tested in the NASA Langley 8-foot transonic wind tunnel. 4 18. Distribution Statement I 17. C-14l-h Wind-Tunnel Test Langley 8-Foot Transonic Tunnel Unclassified-Unlimited 20. Security Classif. (of this page) Domestic, $5.2: * For sale by the Nationa

5、l Technical Information Service, Springfield, Virginia 22151 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. T . . 1 s . . 3 . 4 . 5 . 5 5 5 a . Test617 . 6 ata . Test617 . 6 . Test591 6 617 6 7 . 7 . 8 . III Provided by IHSNot for ResaleNo reprodu

6、ction or networking permitted without license from IHS-,-,-IN TRODUCTI 0 N s information on the wind tunnel test program and basic aerodynamic configuration used in the correlation studies described in Volume 1. The ny was contracted to prepare an existing C-141A high-speed wind on and test in the N

7、ASA wo principal phases, and e 1971 to March 1973. g I ey 8-f oot t ra to scheduling acquisition and analysis of data from both tests was WlND TUNNEL TEST Facility 141A model was made in the NASA Langley 8-f ngle-return ciosed-circuit tunnel. The test sec e sides measure 7.1 feet (2.16m). The upper

8、wer wal Is are it a variable test section Mach number from a ,20 to 1.30, with locka e. The total pressure can be varied from a minimum .2 x 18 N/m2) at all test Mach numbers, to a maximum IO5 N/mq at transonic Mach numbers. The stagnation cally controlled and is usually held constant at720“F ntil t

9、he dew point temperature in the test section is reduced tion effects. The resulting maximum Reynolds number available t (19.7 x lo6 per meter) for this program. Model and Instrumentation 275-scale C-141A model was used for the wind tunnel investigation. accommodate a ne pport system, in which the mo

10、del was supported ng blade from the I e and attached to a support sting below ed and fabricated to locate the NASA e, and two fuselage afterbody fairings of identical geometries were system and the the same as development program adaptor blocks were quirements for testi ctured and tested during iIs

11、of the model dimen Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-ch Center provided a 2.0 inch (5.08 c om total model loads. In addition, a sec balance was used to measure blade loads. T standard fuse lag scanivalves for measuring pressures from 67

12、 s ic orifices on th e fuselage. The layout of these orifices is she ive configuration with a dorsal s atic pressures on the afterbody, and 18 internal pressures e afterbody and one inside a wheelwell fairing. e model angie-of-attack was fott servo acceler tion from which ndevco servo-ac el Support

13、Configurations etails of the principal model configurations are given in figures 2 t 2 shows the model mounted on the live blade and sting combination model. Configurations 2, but with the blade removed. In order to measure the load on the blade separately on configur R e model was attached directly

14、 to the dorsal strut and the blade attac r balance. In addition to the blade lo included in this measurement. e interference of the model on t tioned in a cavit Tests were conducted with different s (f igure 6), was offset fairing o rogram. in config 2 Provided by IHSNot for ResaleNo reproduction or

15、 networking permitted without license from IHS-,-,-ine, and the bullet fairing was therefore not required. toial configuration was included during the test program, and designated CIS obtained from configuration 8 by removing the blade so that measurements the lower sting only in position. Test Cond

16、itions and s fixed on all model surfaces by 0.05-inch (0.125 cm) bands of ballotini in layer of lacquer. The choice of bead size was based on previous eaience on transition fixing for high speed drag evaluation during the C-5A m. ASA Langley has shown that it is possible to obtain ue to the transiti

17、on strip providin narrow, sparsely d istri- . The roughness size for a given eynolds number is deter- 2 600, the ctical roughness ynolds number based on t the top of the roughness, uk, and the kinematic visco- e 2, showed that for a eynolds number 2 3 x lo6 range of Rx values from 0.2 x 106 to to 0.

18、0045 inches (0.0114 cm) no measurable as obtained with transition of the boundary layer occuring immediately nditions. These results suggested that for the C-141A from 3 x 106 per foot (TO x 106 per meter) to 5 x 106 per chosen value of Rx based on ten percent wing MAC, a m) was required. Further, i

19、t was ncluded that no ing was necessary as a result of t C-5A data. Transi- ode1 components, except the fuselage and wheel well e inches (1.855 cm) back from the leading-edges. The he nose of those components. well fairings were placed 2 * (5,08 cm) and tests were completed with transition located f

20、urther aft on the wing upper ly ten percent local chord ahead of the wing shock wave at cruise ess size was 0.0054 inches (0.0137 cm). This technique was included simulation of the interaction of shock and boundary layer at high Mach ions, where forward transition is known to induce a premature rear

21、 Ids numbers for this wing design. Flow visualization tests, wing upper were conducted to locate the main shock position for this investigation. eral, covered the mber was 5.0 x 106/foot (16.4 x 106/mete ach number range from M = 0.600 to 0.825. In most or 3.05 x 106/MAC. 3 Provided by IHSNot for Re

22、saleNo reproduction or networking permitted without license from IHS-,-,-Scale effects were obtained on some configurations by testing also at 3. foot (10 x lo6 and 13.12 x 106/meter). Six-component force m at zero yaw over an angle-of-attack range from -4O (-0.0619 Selected configurations also incl

23、uded measurements of the fuselage afte cavity pressures. A summary of the test program is given in tab1 tions, Nos. 4 and 8, showing the installation in the photographs of figures 7 and 8. Data Reduction and Corrections Force balance data was reduced to coefficient form in the stabili a reference wi

24、ng area of 2.44 square feet (0.227 square meters), and a chord of 7,328 inches (18.62 cm). Pitching moment was referred to a cog. The accuracy of the strain gaged nce urements was b ers quoted figure of = 0,0010 and 0 A006, respectively, he 591 data. As a final check on the validity of the initial 6

25、17 data, the retest il-off configuration during March 1973 confirmed these results. It should be lasic data given in this volume indicate these discrepancies from both test the subsequent analysis used only some limited component drag and trim m the 591 test, and the main analysis was concentrated o

26、n the test 617 Flow ins moment and drag Angularity. Test 617 data for model upright and inverted runs are shown in and (c). These data have been corrected for the flow angularity and indi- ent over the full ach number ran: . C ;he tests. Repeatability. Test 617 es of the degree of test data repeatab

27、ility during the 617 series are given in In general, values of drag are repeated within ACD =0.0001 to 0.0002. ows the repeatability of the afterbody pressures by a comparison of integrated 1 tests. 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-C

28、omplete Model Baseline ata. Test 697 Figures 11 (a-d) present some of the complete model baseline data w the analysis and drag correlation studies in volume 1. These data were obt wing surface finish was prepared to a high degree of smoothness and a new applied prior to this main series of runs. Com

29、parison of these data indicate! change it-, drag level of the order of ACD = 0.0001 for this configration. results for the model baseline is given in figures 20 (a-d). Support Tare and Interference r configuration 7, the blade load measurements, are 12 (a) and (b). Results from the tests with the do

30、rsal sting arrangement, con 5 and 10 are given in figures 13 through 16. The sting interference eval are given in figures 17 through 20, for configurations 1, 2, 3 and 4 respective for the central sting configuration, No. 9 are contained in figures 21 thro Basic Data. Test 5 Selected results from th

31、e 591 series are presented in figures 25 t three horizontal tail settings are contained in figures 25, 26 Reynolds number range from 1.83 x 106/MAC to 3.05 x 10 through 30 for the complete model configuration, and those for tail-of figures 31 and 32. The pylon-nacelle effects can be derived from the

32、 33 and 34. Aft Located Transition. Test 617 Results for the tests with wing upper surface transition moved back fror location to a position ten percent local chord ahead of the shock at cruise cion given in figures 35, 36 and 37. 6 Provided by IHSNot for ResaleNo reproduction or networking permitte

33、d without license from IHS-,-,-Miscellaneous Data. Test 617 ta from which pylon-n ment are derived are given in cts of sting ff, are given in figures Afterbody Pressure Data lar plots giving static pressure variation with fuselage station are shown in 54. These, in general, have been selected to sup

34、port the discussion of re drag effects, for the different model support configurations, given in 7 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-REFERENCES 1. Braslow, A. L.; Hicks, R. M.; and Harris, R. V. Jr.: Use of Grit- Layer-Transition Trips

35、on Wind-Tunnel Models. NASA TN D-3579, 2. Palfery, J. G.: C-5A igh Speed Drag Reduction Study and lnvestigatio Boundary Layer Transition on a 0.0226 Scale Model in the NASA-Ame Tunnel. High Speed Series 4C and 40. Lockheed- Georgia Company LGlT6-1-40, April 1968. 8 Provided by IHSNot for ResaleNo re

36、production or networking permitted without license from IHS-,-,-axirnurn diameter, inches uiselage reference line 8 airing, b TABLE I MODEL DIMENSIONAL DATA 0.0275 Scale 43.656 4.675 9.362 WL 5.500 FS 6,336 eness ratio ( I/D) rea, square feet clot chord Peading edge izontal Stabi I izer , H 8 n aero

37、dynamic chord, inches weep of 25% chord, degrees ail length, inches ail volume coefficient oil section NACA 64A(010)010.5 FS FS 54.148 8.246 0.017 FS 37.998 0.365 16,554 0 370 25.000 23.885 0.488 9 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MODE

38、L DIMENSIONAL DATA (CONT.1 19 Nace.lle Pvlon, K 0.027 e Inboard area, square feet inboard span, inches Inboard chord, inches Outboard area, square inches Outboard span! inches Outboard chord, inches Sweep of leading edge, degrees 7 Nacelles. Length, inches Inlet diameter, inches Exit diameter, inche

39、s Inboard Leading edge location Outboard leading edge location nboard toe-in, degrees oard toe-in, degrees 6 Vertical Stabilizer, V Area, square feet ean aerodynamic chord, inches FS 5L WL FS 51 WL 1 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-OD

40、EL DIMENSIONAL DATA (CONT.) Stabilizer (Cont. _._I_ ica I tai 1 volume coefficient :an aerodynamic chord, inches of 25% chord degrees (inboard) ?h mean aerodynamic chord location eep of 25% chord, degrees (outboard) eral of 25% chord, degrees (inboard ord degrees (outboard) mk station chord, inches

41、(inboard station I oca tion (outboard) 0.0275 Scale 1.208 0.617 35.000 21.065 0.049 2.440 52.689 7.328 7.900 0.373 23.734 FS 25.378 WL 7.248 B 10.818 25.025 4.891 -5 * 584 0.941 1.195 10.91 1 6.419 6.487 WS 11.138 ws 11.734 1 Provided by IHSNot for ResaleNo reproduction or networking permitted witho

42、ut license from IHS-,-,-Wing (Cbnt.) - MODEL DIMENSIONAL DATA (CONT.) 0.0275 Scale Airfoi I section Root NACA 0013.0-1.10-40/1,575 (MOD) a0=0.8 (MOD) C1. =0.153 Inboard break 1 NACA 001 1.2-1.10-40/1.575 (MOD) a. = 0.8 (MOD) C1. = 0.194 I Outboard break NACA 001 1 .O-1.10-40/1.575 (MOD) a. = 0.8 (MO

43、D) c1. = 0.201 I Tip NACA 0010.0-2.20-40/1.575 (MOD) mean line 1/2 (NACA 66 at C1. = 1 .O-NACA 230 at C1 = 1 .O C1. =0.452) I I a1 Antenna Fairing, Z Leading edge location Trailing edge location G2 I Wheelwe I I Fairing ,Z Length, inches Maximum frontal area per side, square inches Fineness ratio Le

44、ading edge location w7 Wing-Fuselage Fillet, Z Leading edge location Trailing edge location FS 26.208 FS 33 . -It- / h c 17 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-a3 I= 0 .- c U L 2 0) f 0 .- cc v L 0 Q Q 3 m c 18 Provided by IHSNot for Resa

45、leNo reproduction or networking permitted without license from IHS-,-,-19 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TABLE 2 WIND TUNNEL TEST SCHEDULES (A) TEST591 Description 9 12 W7 19 7 a1 G21 BW Z K NZ Z 01 nu IO IS to I! no (Inverted) 9 12

46、W7 19 7 a1 G21 4 6 8 8 BW Z K NZ Z DVH.b -I II 11 II so 11 in Io nI 11 nt I! nu II on in 11 IU it II nn OB I! IO nu IO II II 81 !I II IO II II 9 12 W7 19 7 a1 G21 4 6 8 8 BW Z K NZ Z DVHb 0 4-1 to ii ti It it ii ii i! in H8 ii 9 12 W7 19 7 a1 G21 4 6 8 8 BW Z K NZ Z DVH,b - II ni ni in II ii it in i

47、n nt ti 01 II !I 01 00 I1 (I It #I I! iI 9 12 W7 19 7 a1 G21 BW Z K NZ Z nu nn it OB II II st 9 12 W7 a1 G21 BWZ ZZ B9z G2 1 2 2 1 4 4 4 4 4 4 4 4 4 4 5 .O 5 .O 5 .O 5 .O 5 .O 5 .O 5 .O 5 .O 5.0 4.0 3.0 3.0 5.0 5.0 5.0 1/8 7 16 15 55 54 53 62 61 t 69 68 67 -s 73 72 - - 74 75 - 81 80 - - - 85 84 - 86

48、 87 - - 93 92 I- m- 97 96 I- - 20 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-(B) TEST 617 S u pport Config. escri ption RdFt x Zw 7( I nverted) 2 2 Zal ZG2 Zw7B lade Sea I 7Za 1 ZG21 ZW7D4V6H8,b8 - 7 7 8 6 5 5 10 4 4 4 4 5 .O 5 .O 5 .O 5 .O 5 .O 5 .O 5 .O 5 .O 5 .O 5 .O 5 .O 3 .O 3 .O 3 .O 5 .O 5 .O 5 .O 5 .O 5 .O - 45 44