NASA-TN-D-7331-1973 Effect of blockage ratio on drag and pressure distributions for bodies of revolution at transonic speeds《在跨音速下 阻塞比对回转体阻力和压力分布的影响》.pdf

《NASA-TN-D-7331-1973 Effect of blockage ratio on drag and pressure distributions for bodies of revolution at transonic speeds《在跨音速下 阻塞比对回转体阻力和压力分布的影响》.pdf》由会员分享，可在线阅读，更多相关《NASA-TN-D-7331-1973 Effect of blockage ratio on drag and pressure distributions for bodies of revolution at transonic speeds《在跨音速下 阻塞比对回转体阻力和压力分布的影响》.pdf（114页珍藏版）》请在麦多课文档分享上搜索。

1、11r NASA TECHNICAL NOTE!ZI-.,:!Z; !kNASA TN D-7331EFFECT OF BLOCKAGE RATIO ON DRAGAND PRESSURE DISTRIBUTIONS FOR BODIESOF REVOLUTION AT TRANSONIC SPEEDSby Lana M. Couch and Cuyler IV. Brooks, Jr.Langley Research CenterHampton, Va. 23665NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D.C. N

2、OVEMBER 1973Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-, , /_ i _1. Report No. 2. Government Accession No. 3. Recipients Catalog No.NASA TN D-73315

3、. Report DateNovember 19734. Title and SubtitleEFFECT OF BLOCKAGE RATIO ON DRAG AND PRESSUREDISTRIBUTIONS FOR BODIES OF REVOLUTION ATTRANSONIC SPEEDS7. Author(s)Lana M. Couch and Cuyler W. Brooks, Jr.9. Performing Organization Name and AddressNASA Langley Research CenterHampton, Va. 2366512, Sponsor

4、ing Agency Name and AddressNational Aeronautics and Space AdministrationWashington, D.C. 205466. Performing Organization Code8. Performing Organization Report No.L-844910. Work Unit No.501-06-09-0111. Contract or Grant No.13. Type of Report and Period CoveredTechnical Note14. Sponsoring Agency Code1

5、5. Supplementary Notes16. AbstractExperimental data were obtained in two wind tunnels for 13 models over a Machnumber range from 0.70 to 1.02. Effects of increasing test-section blockage ratio in thetransonic region near a Mach number of 1.0 included change in the shape of the drag curves,premature

6、drag creep (i.e., transonic creep), delayed drag divergence, and a positive incre-ment of pressures on the model afterbodies. Effects of wall interference were apparent inthe data even for a change in blockage ratio from a very low 0.000343 to an even lower0.000170. Therefore, models having values o

7、f blockage ratio of 0.0003 - an order ofmagnitude below the previously considered “safe“ value of 0.0050 - had significant errorsin the drag-coefficient values obtained at speeds near a Mach number of 1.0. Furthermore,the flow relief afforded by slots or perforations in test-section walls - designed

8、 accordingto previously accepted criteria for interference-free subsonic flow - does not appear to besufficient to avoid significant interference of the walls with the model flow field for Machnumbers very close to 1.0.t7. Key Words (Suggested by Authoris)Transonic wall interferenceDragFacilities19.

9、 Security Classif. (of this report)Unclassified18. Distribution StatementUnclassified - Unlimited20. Security Classif. (of this page)Unclassified21. No. of Pages109*For saleby the National Technical Information Service,Springfield, Virginia 2215122. Price*Domestic, $4.2_Foreign, $6.7!Provided by IHS

10、Not for ResaleNo reproduction or networking permitted without license from IHS-,-,-iProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-EFFECT OF BLOCKAGE RATIO ON DRAG AND PRESSURE DISTRIBUTIONSFOR BODIES OF REVOLUTION AT TRANSONIC SPEEDSBy Lana M. Couc

11、h and Cuyler W. Brooks, Jr,Langley Research CenterSUMMARYIn an attempt to determine the severity of wind-tunnel wall interference under near-sonic test conditions, aerodynamic force and pressure measurements were obtained atzero normal force for 13 bodies of revolution over a Mach number range from

12、0.70to 1.02.Three specific effects on the drag data occurred near a Mach number of 1.0 as aresult of increasing the test-section blockage ratio for a given model profile:i. The shape of the drag curves changed from a relatively rapid increase to a verygradual increase in drag with increasing Mach nu

13、mber. This change occurred forincreases in test-section blockage ratio above approximately 0.0010. The shape ofthe drag-coefficient curves, obtained at values of test-section blockage ratio less than0.0010, is relatively insensitive to changes in blockage ratio.2. Increasing the blockage ratio above

14、 approximately 0.0003 produced a premature,positive deviation, or transonic creep, of the drag curve from the trend of the subsonicdata. Since the Mach number at the initiation of transonic creep agrees with the calcu-lated Mach number for choked flow in a solid-wall tunnel, transonic creep may be t

15、hefirst indication of significant wind-tunnel wall interference near a Mach number of 1.0.3. The occurrence of drag divergence was delayed by approximately 0.013 in Machnumber because of an increase in blockage ratio from 0.0002 to 0.0010. For blockageratios greater than 0.0010, the drag-divergence

16、Mach number was essentially constant.Therefore, near a Mach number of 1.0 increasing the blockage ratio delays the occurrenceof drag divergence - the result that would be expected because of an effective decrease inthe free-stream Mach number at the model.There was only one obvious effect of wall in

17、terference on the model surface-pressure distributions obtained for a given model shape at different values of blockageratio. For Mach numbers greater than approximately 0.96, a region of supersonic flowexisted around the models. An increase in the value of blockage ratio for Mach numbersgreater tha

18、n 0.96 caused a positive increment of pressure to occur on the model. TheProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-effect on the drag data of this pressure-drag increment coincided quite well with thechange in shape of the drag curves.The resul

19、ts of this investigation indicated that models having values of test-sectionblockage ratio of 0.0003 - an order of magnitude below the previously considered “safe“value of 0.0050 - had significant errors due to wall interference in the drag-coefficientvalues obtained near a Mach number of 1.0. Furth

20、ermore, the flow relief afforded byslots or perforations in test-section walls - designed according to previously acceptedcriteria for interference-free subsonic flow - does not appear to be sufficient to avoidsignificant interference of the walls with the model flow field for Mach numbers veryclose

21、 to 1.0.INTRODUCTIONIn the past, numerous experimental investigations have been conducted in transonicwind tunnels to determine the drag of both complete aircraft configurations and variousaircraft components at high subsonic and transonic Mach numbers. For these investi-gations, data obtained at te

22、st-section blockage ratios of 0.0050 or less were generallyconsidered to be free of wall-induced blockage effects. However, the advent of the super-critical design concept and the effort to develop a near-sonic transport have placedrenewed emphasis on accurate drag measurements in the near-sonic spe

23、ed range (Machnumbers of 0.95 to 1.0). As a result of this increased emphasis, it was believed that theseverity of wind-tunnel wall interference should be reexamined at near-sonic Mach num-bers. Therefore, geometrically similar bodies of revolution were tested in two windtunnels and in flight under

24、the same conditions. The model-to-wind-tunnel blockage ratioswere 0.0028 and 0.000684. A comparison of the results obtained in the flight test andwind-tunnel test, at a blockage ratio of 0.0028 (ref. 1), indicated differences in drag char-acteristics of sufficient magnitude to cause concern about th

25、e validity of drag data obtainedin the wind tunnel near a Mach number of 1.0. This discrepancy was believed to be theresult of wall interference due to test-section blockage in the wind tunnel. To investigatethis result further, bodies of revolution which provided a systematic variation in blockager

26、atio from 0.00017 to 0.0043 were tested in the Langley 16-foot transonic tunnel and theLangley 8-foot transonic pressure tunnel to assess the extent of wall interference effectsneara Mach number of 1.0. The significant variables included in the investigation wereMach number and model diameter and le

27、ngth.SYMBOLSValues are given in both SI and U.S. Customary Units. The measurements and calcu-lations were made in U.S. Customary Units.2Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-A local cross-sectional area of modeld2Adx 2AmaxA sC DCD ,bCpCp,bC

28、p,sonicdKirefMNRePq_reference area, maximum cross-sectional area of modelcross-sectional area of stingdrag coefficient at zero normal force, adjusted to a condition of free-streamstatic pressure at model base, Dragq_Amaxbase-drag coefficientpressure coefficientbase-pressure coefficientpressure coeff

29、icient corresponding tomaximum diameter of modelM = 1.0 local flowcurvature constant in model profile equationslength of modelreference length of model (length to body closure)free-stream Mach numberReynolds number based on model lengthmodel surface pressuremodel base pressurefree-stream static pres

30、surefree-stream dynamic pressureProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-R radius of modellongitudinal center-line distance from model noseX m longitudinal center-line distance from model nose to maximum diameterof model(7 standard deviation o

31、f variableAPPARATUSModel DescriptionThe five model shapes investigated in the wind-tunnel tests were (I) the blunt-nosesupercritical body of type A, (2) the blunt-nose, finned body of type B, (3) the blunt-nose,supercritical body of type C, (4) the blunt-nose, supercritical body of type D, and (5) t

32、hepointed-nose body of type E. Sketches of all models are shown in figure I; the lettersindicate the profile type and the numbers are coded to the diameters - as indicated inthe following table:Configurationnumbercode123456cm6.358.9912.7015.7022.2527.08Diameterin.2.503.545.006.188.7610.66Pertinent d

33、imensions of all models and stings are listed in table I, and model coordinatesare listed in table II.Generally, the models had machined metal noses and were otherwise constructed ofwood with fiber-glass-covered surfaces, or they were made totally of aluminum. Themodels were turned on a lathe to obt

34、ain the required profiles within a tolerance of+0.0025 cm (+0.001 in.) in radius. Although the models were generally smooth, deviationsin the radius of the models may have been as much as 0.0127 cm (+0.005 in.) due towarping or shrinking of the wood. These deviations were well within the limits for

35、hydrau-lic smoothness for the test conditions and model sizes of this investigation. (See ref. 2.)4Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-The surface-pressure orifices were constructed of metal tubing with 0.076-cm(0.030-in.) inside diameter

36、, installed flush with andperpendicular to the model surface.Onelongitudinal row of pressure orifices wasinstalled in each model, and eachmodelwas tested with the orifices in the sameorientation in roll. The locations of all surface-pressure orifices are listed in table III.Type A.- The profile of t

37、he blunt-nose, supercritical body of type A is defined bythe following equations:For 0 x Xm=lref =lre fA,/ x _ = K xiref ) - A (l_ef_Xm x 0.990and for lref = lrefXm 1where = -. The profile is symmetrical fore and aft of the point of maximum diam-lre f 2eter, except for truncation at the base to acce

38、pt the sting. Three models of this profile,having maximum diameters of 8.99, 12.70, and 15.70 cm (3.54, 5.00, and 6.18 in., respec-tively) and a fineness ratio (i/d) of approximately 9.0, were investigated.Type B.- The profile of the blunt-nose, finned body of type B was identical to thetype A profi

39、le, except that the aft 12.7 cm (5.0 in.) of the body were modified. The cross-sectional area distribution of the body including the four 3-percent-thick delta biconvexfins was equivalent to the cross-sectional area distribution of the A-3 body. One modelof this profile having a 12.70-cm diameter (5

40、.00 in.) was investigated in the wind tunnelsand in a flight test (ref. I).Type C.- The profile of the blunt-nose, supercritical body of type C is defined bythe following equations:X XmFor 0 =lref - lref,K x lref)5Provided by IHSNot for ResaleNo reproduction or networking permitted without license f

41、rom IHS-,-,-and for Xm x 0.990,iref - iref -A,fx _= Kire f AmaxXm 1.0 Xmwhere - The profile for 0 _- x _ is identical to that of type A;lref 1.0 + (2/_)“ lref lrefhowever, aft of the location of the maximum diameter, the profile has a constant secondderivative in the area distribution. The model is

42、truncated at the base to accept the stingand to give a ratio of base diameter to maximum diameter equivalent to that of the type Amodels. Consequently, the type C models have a fineness ratio (l/d) of approximately9.5 - a value slightly larger than the fineness ratio of the type A models. Five model

43、sof this type, having maximum diameters of 8.99, 12.70, 15.70, 22.25, and 27.08 cm (3.54,5.00, 6.18, 8.76, and 10.66 in., respectively), were investigated in the wind tunnels.Type D.- The profile of the blunt-nose, supercritical body of type D is defined by thefollowing equations:X (: XmFor 0 =lre f

44、 - lrefXm xand for _ - 1.052,Iref iref/l_e f/ (v/2)(x - Xm_A“ x =Ccos_ l_efLX_ _where Xm_ 1.0 and CIref 12( 1 A-_a;) +1“0( lm;C = (A s - Amax) lref - xis a function of the base area and is defined byTherefore, the actual base is slightly larger than As, and the actual length of the modelis slightly

45、less that that predicted by the equations in order to provide clearance for thesting. Three models of this type, having maximum diameters of 6.35, 8.99, and 15.70 cm(2.50, 3.54, and 6.18 in., respectively), were investigated in the wind tunnels.Type E.- The pointed-nose model of type E has a modifie

46、d Sears-Haack profile, asgiven in reference 3. This model has a maximum diameter of 12.70 cm (5.00 in.), aProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-i / H / _ fineness ratio I/d of approximately 9, and Iref/d of 12.0. The free-flight model ofref

47、erence 3 also had a fineness ratio iref/d of 12.0, but the maximum diameter was25.4 cm (i0.0 in.) and there was no simulated sting-type support.StingsAll tests were made with models sting supported through an access hole at themodel base, except for the model which was tested in flight. The model su

48、pport stings,except for configuration C-6, were scaled to approximately the model size and weredesigned in accordance with the criteria of reference 4 in order to minimize the effectsof sting interference on the flow over the aft end of the model. This reference specifiesthe minimum length of straight_-sting section aft of the model base and the maximum angleof the sting flare. The ratio of cylindrical sting length to sting diameter for configura-tion C-6 was smaller than desirable. Pertinent dimensions and flare angles of the stingsare presented in table I.Wind TunnelsTh