NASA NACA-RM-A9B17-1949 Chordwise and spanwise loadings measured at low speed on large triangular wings《在低速时测量大三角形机翼弦向和顺翼展方向的荷载》.pdf

《NASA NACA-RM-A9B17-1949 Chordwise and spanwise loadings measured at low speed on large triangular wings《在低速时测量大三角形机翼弦向和顺翼展方向的荷载》.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-RM-A9B17-1949 Chordwise and spanwise loadings measured at low speed on large triangular wings《在低速时测量大三角形机翼弦向和顺翼展方向的荷载》.pdf（79页珍藏版）》请在麦多课文档分享上搜索。

1、F 1 ,- CHORDWISE AND SPANWISE LOADINGS MEASURED AT LOW SPEED ON LARGE TRlANGULAR WINGS a - By Adrien E. Anderson Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-“ _c Pressure distributiom have been obtained *om three triangular wing modela : a -lone

2、model having 89 aspect mtio of 2.04 and a modified hubme airfoil eection, the same wing caibined ulth a body of fineness ratio 12.5, and a mockiup of a tri“ sirplane which had an aspect ratio of 2-31 and then, in conjunction with reference 1, makes available a comparison of the loading on two triang

3、ular wings, one having a relatively thick subsoni-type ernd the other a thin super- sonio“type airfoil section. In addition, the data presented in this report indicate the effect of 8 body on the wing load distribution as well as prodding a qualitative comparison of the loadings on two wing- body cm

4、binationa, one having a thin subs0ni”type and the other a thin sugerecmic-type airfoil eection. The symbols and coefficients ueed in this report are defined as follows : A aspect ratio a free-etresm angle of attack. of wing chord plane, degrees Provided by IHSNot for ResaleNo reproduction or network

5、ing permitted without license from IHS-,-,-RACA RM NO. 917 3 a b wing span, feet A C wing chord, measured parallel to air stream, feet average wing chord (S/b) , feet - C mean aeroaynamic chord, measured parallel to sir / section lift coefficient wing lift coefficient (9) P Pt P -Q S X Xf Y fre-trea

6、m static pressure, pounds per square foot local static pressure, pounds per square foot freestream dynamic pressure, pounds per square foot wing area, square feet distance along chord Fram leading edge, feet distance along Melage center line from nose, feet distance dong wing semispan to chord locat

7、ion, feet Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 RACA RM Eo. AgBl7 The three models used in this investigation were: 1. A wing model, hereafter referbed to as the mdifieb-wedge ving, which had a triangular plan form and an aspect ratio of

8、2.04. Thle airfoil section parallel to the wing center line was derived frum 8 synrmetrical dauble-wedge secttan having a maximum thicknesa of %percent chord at 2O”percent chord. The modification conrriated of- a 0.0025 =Be radiue and a rounded maxfmum thicknee6 formed from an arc (0.62 radius) whic

9、h uas tangent to the eurface of the double wedge at 1% and *percent chord. Both the tog and bottom of the section were rounded. In term of the resultant chord, the airfoil section had a nose radlue of 0.00254 and a thickness of 4.83 percent at 21.6-percent chord. This was the same wing for which for

10、ce data were presented in reference e. 2. A modified.Kedge winpbody model, which consisted of the abovmtioned wing combined with a slender and polnted body of revolution. The radius r of the body at any ststion xf was obtained from the following eqpation: 3. A mock-up of a triangula3Ywing airplane w

11、hich had an aspect ratio 2.31 and an HACA 65-006.5 airfoil section parallel to the King center line. This model, which will be referred to as the NMA - series wing-body model, wa8 equipped with conatanhhord trailing- edge controls. The control had a horn balance, a nose radius, esd a smell but unsea

12、led gap. The ducting sgetem was gpen and the power plant removed. We-view drawings of the two K.Lng-baciy modele appear in figure 1, while figure 2 contains photographe of the models as mounted for telting in the Ames 4” by 8 they may have been weaker. .c During the investigation reported herein, va

13、por trails did not appear on any of the models. In the case of the modifieihredge wfng model, however, their prelaence was detected by means of a survey of the flow above the wing. Their presence is also apprent in the chordwise pressure plots for the modifiedwedge wing and -body models. There is so

14、me evidence in the pressure plot8 for the HCLCA 6eries -body model to indicate that separation vortices may have existed on this model also. An understanding of the pattern of this vortex type of flow will aid in the interpretation of the resulta of this investigation. $ The following description of

15、 the fomtion of the separation vortices is based on the visual observation of vapor trails on the wedge wing (referenoe 2). and the survey of the flar above the mod3- fied-wedge-wing model. The vortex, oonaidering one-half of the wing at a time, originated firet on the lead- edge near the wing tip,

16、following separation of the flow Fn this region. Line A of figure 6(b) represents the vortex pattern for a = 30. hcreming the angle of etttaok caurred the point of origin of the vortex to move Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM N

17、o. AgB17 7 . J . forward along the lea- edge toward the apex of the wing until the point of origin reached the apex at about a = 8O. The angle of sweepback of the yoetex increased with angle of attack, the rate of increase betag largest at the lower angles of attack of the wing. Also, at the lower a

18、ngles, that portion of the vortex lying over the outer 30 percent of the wing spen uas bent back elightly toward the free stream, me latter condition is represented by line B of fw ure 6(b). At an angle of attack of about 18*, the angle of sweep back of the vortex seemed fairly well established and

19、apparent- the strength of the vortex begas to dlmfnish toward the apex (the yapor had stalled, all evidence of the vortex had disappeared. , trails of fig. 6(a) or line C of fig. 6(b). By the time the wing General Comments Catparison of the results of this investigation with those for the wing with

20、MCA 0012 airfoil section ,(reference 1) indicated that some generalities exist concerning the wing characteristics which should be kept in mind throughout the discussion to follow. Similar pressure distributions occurred over the wings only at the lower angles of attack where esaentially fnviscid po

21、tential flow existed on the wings. The preesure dietributione became dissimilar in the middle anglf-attack range follaring flow separation from the leading edge. The occurrence of lealllng-edge separatfon and the intensity of the resultant vortex-type flow appeared to be dependent upon the airfoil s

22、ection. The modifiemdge wing and winody models showed signs of leadlng-edge separation first and a stronger effect of the vortex-type flow. In the case of the HACA 65series -body model, separation was delayed and the effect of the vortex lese strong. It is d2fficult to conclude fram the pressure dia

23、grma for the model with NAca OOl2 sections (reference I) if the vortex- type flow was present. The nrinor bumps in the pressure diagrams make it appear likely that this model had a very weak vortex-type flow. At angles of attack near wing stall, where the Fnfluence of the vortex-type flow had became

24、 negligible, the characteristics of the three models were nearly similar. In general, the characteristics of the NACA 6Fseries wimody model more closely reedled those for the model with WA 0012 eection than those for the two models with modif ie-dge sections. Chordwise Pressure Distribution The chor

25、dwise pressure diagrams for the modifiei+wedge-wing model are presented in figure 7. Whereas data were obtained at very Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 NACA RM No. with the negative pressure values increasing toward the root. The di

26、agrams for the inboard stap tione give indication of various degrees of separated flow, with the most inboard station having the- least si- of separated flow. The preasure-dietribution curv-es for the modifM+wedge xinep body model are preeented in figure 9. The prlrmarg effect of adding the body wa

27、to shift the origin of the separation vortices outboard. The outward shift of the vortices is mre clearly seen by comgaring the pressure curves for the -ody conibinstion (fig. 10) with the Wing-alone pressure curves at the same angle of attack. (See fig. 8. ) Although the bumps on the inboard statio

28、ns had a more negative pres- sure peak, they extended over a shorter chord distance, the net result being a reduction in lift on the moard stations as compared to the wingglone model. The chordwise pressmstribution curves for the RACA 6Meries “body model are presented in figure ll. In general, the p

29、ressure diagrams for this model closely resable those for the WA 0012 wing (reference I). The chordwise distribution of pressure Etnd tie value of the wing lift coefficient at xhich separation occurs on a given section are slaibr. At the 0.60-semispas station, for -le, the pressure diagram agree clo

30、sely at the law angles of attack and probably would show closer agreement in the middle rmge of de of attack had it been p0ssib;le to record pressures at the very leading edge of the RACA 6j-series King. At maximum lift, the corresponding pressure diagrams for the two winga are again very similar. T

31、he data outboard of the 0,hemiapas statdon on the IUCA 65- series -body model are open to qwation bemuse of the previously mentioned scarcity of Treasure orifices. For example, the presaure diagram for the O.gLsemispan station give the hrpressian that separated flaw existed even at the vei-y lowest

32、angles, which ia proh ably not the cme. Following complete separation, however, the pree sure diagrams for the O.emispsm station reedle thoae for the modifieo higher angles of attack by use of a subsonic-type airfoil section. The section lift curve for this model at the 0.- semispan station bear8 no

33、 reseniblance to the curve for the modifieb wedge wing-body model over the first 12 angle of attack, as would be expected from the previous discussion of the pressure diagrams for this station. At the higher angles, where it is re860n8ble that complete separation has occurred, the tX0 curves for the

34、 0.9Cbsemispan station agree more closely. The section lift curves, particularly those for the modifieL.ge Scale Triangular Wing of Aapect Ratio %.- II. The Effect of Airfoil Section Modifications and the Determination of the Wake Downwash. RACA RM No. A7H28, 1947. Provided by IHSNot for ResaleNo re

35、production or networking permitted without license from IHS-,-,-16 4 I Item =de wing Span, feet Area exposed outaide of - 307 Area, square feet 25 .oo fwlage, square feet 307 Mean aerodynamic chord, feet 16.37 Dihedral, dsme 0 Angle of incidence, degrees 2 .Oh Aspect rat io -“ IRngth, feet “- Maximu

36、m diaaaster, feet “- Finernee ratio Maximum diameter wing- “- span ratio “- Fuselage Trafling-dge control Area (total aft of hinge Area (total wfth horn ) , SqUere f 00% “- balance), square feet “- 1 I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-

37、Station 0 1.25 .!X . 2.5 5 -0 7.5 10 00 15.0 20 .O 25.0 30 .O 40.0 50 .O 60.0 70 .O 80.0 go .o 95 90 100.0 Ordinates 0 +- .5l2 +- .771 f 1.032 z 1.415 k 1.719 f 1:gn f 2.379 f 2.685 * 2.920 * 3.087 f 3.244 f 3.133 f 2.719 f 2.095 f 1.327 * 548 f .210 0 L. E. radius: 0.2s percent chord Provided by IH

38、SNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-c r ! I 1- 0 - - u fa) Mod;fied-wedge Figure L- General arrangemenf of fhe wing-body mode/s investigafed + Provided by

39、 IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-20 1 f3.03 I 4 I! “ t- “37 (b) NACA 65-series wing-body mode/. Figure 1- Conchded Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-CA BM NO. 9.17 21 (a) Modifi-wedg

40、e Kfng model. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-t Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. - . . . . . . . . . . 4 Provided by IHSNot for ResaleNo reproduction or networking per

41、mitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . . . . . . . . . . . * . . . . . . . . . . 4 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. Provided by IHSN

42、ot for ResaleNo reproduction or networking permitted without license from IHS-,-,-i .I . . . D I 1 I (b)NACA *se&s whg-bot model. Figure 3.- Approximate pressure &ifice distribution. . I 3 . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I I1 Figu

43、re 4- Method of deriving chwdwjse pressure diagrums on NACA 65-serjes wing-body moobl. 3 .- - . . . . . . . . . . . . . . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. - . . . . . . . . I I c . , 4 Figwe 5.- Force test lift and pitching-mnt was. ro D I . . . . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-