NASA NACA-RM-L54C16-1954 Experimental effects of propulsive jets and afterbody configurations on the zero-lift drag of bodies of revolution at a Mach number of 1 59《在马赫数为1 59时 推进式喷.pdf

《NASA NACA-RM-L54C16-1954 Experimental effects of propulsive jets and afterbody configurations on the zero-lift drag of bodies of revolution at a Mach number of 1 59《在马赫数为1 59时 推进式喷.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-RM-L54C16-1954 Experimental effects of propulsive jets and afterbody configurations on the zero-lift drag of bodies of revolution at a Mach number of 1 59《在马赫数为1 59时 推进式喷.pdf（34页珍藏版）》请在麦多课文档分享上搜索。

1、 - Copy ZUJRM L54C16EXPERIMENTAL EFFECTS OF PROPULSIVE JETS AND AFTERBODYCONFIGURATIONS ON THE ZERO-LIFT DRAG OF BODIES OFREVOLUTION AT A MACH NUMBER OF 1.59By Carlos A. de Moraes and Albin M. NowitzLangley Aeronautical LaboratoryLangley Field, V for nozzle 4, thewas 2.16.A sketch of the assembled n

2、mdel, prior to testing, isure 3.Mach numbershown in flg-TESTS AND INSTRUMENTATIONA detailed description of the preflight jet used in this investiga-tion is given in reference 6. The present-tests were conducted in the27- by 2T-inch jet at a Mach number of 1.59. The stagnation temperature.was approxi

3、mately 70 R and the free-stresm static pressure was standardsea level. The Reynolds nurriherwas 1.7.8x 106, based on model length.,.A photograph of a typical setup priorto a test is shown as figure 4.In order to have the model completely within the Mach dismond of thefreejet and to meet the interfer

4、ence criteria presented in references 7and 8, the nose of the model wa6 placed 8_inches-upstre of the Jetexit. Pressure measurements on the model and of the tunnel conditionswere obtained with electrical pressure pikups of the strain-gage type.Free-stream stagnation temperature was measured with-m i

5、ron-constarhn -thermocouple. All data were recorded by oscil.lographs.Shadowgraphswere made of all tests and were time correlatedwith the pressure data.Estimated accuracies of the test psmmeters are given in the fol- -lowing table:Free-streamMach nuniber, . . . . . . , .Pressure coefficient,Cp . . .

6、 . . . . . .Jet pressure ratio, pj p. . . . . . . . . ./. 0 a71 * . M.03. . . . . . . . . . (b) the boattail angle, P; (c) the jet nozzle half-angle, A; and(d) the jet-exitMach number, Mj.The results of the present tests sre presented as pressure distri-butions and pressure drag. No attempt has been

7、 made to include the fric- .tion drag because it would vsry with the Reynolds number and heating con-ditions of a particular flight plan.Power OffBoattail pressures.- Boattail power-off pressure distributions weredetermined theoreticallyby the method of characteristics (ref. 9) andare presented in f

8、igure p(a) as pressure coefficient plotted againstsxial distance from the model nose. Experimentally determined pressuredistributions, which were obtained over the afterbody sections only,are also shown for purposes of comparison.The pressures measured on the afterbody of rrmdel1 show a trenddissimi

9、lar to theory. Although positive pressures on cylindrical after-bodies have been reported before which seem to substantiate the measure-ment at station 0.947, the measurements on the afterbody of model 1 weretoo few to either substantiate or reject the theoretical pressure distri-bution even though

10、the large drop-off of pressure at station 0.992 wasnot predicted by theory. This sudden decrease in presswe is due to thelocation of the orifice in the expansion field of the flow as it turnsthe corner of the base.The theoretical pressure distributions for models 2 md 3 correctlypredict the increase

11、 in expansion and in the boattail pressure gradientwith increasing boattafi angle. However, for both models the predictedexpansion was too large. The measured pressure distribution over theboattail of model 2 ( = “) was psrallel to, but less negative than,the theoretical pressure distribution. The p

12、ressure measurement atstation 0.997 was not made in the present tests but was obtained on anidentical model tested at the same Mach number. Here again a pressureorifice, located within the expansion field at the base, measured a pres-sure that was considerably lower than that which would be expected

13、 froman extrapolation of the measurements in the present tests.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-The three pressure orifices on the boattail of model 3 (p = loo)were not sufficient to give a good pressure distribution. AE b thecase of t

14、he rearmost orifices of models 1 and 2, the orifice at sta-tion 0.992 read considerably lower than the theoretical value at thatstation. In view of the fact that the measured distribution over theboattail of model 2 was parallel to the theoretical distribution, a curvewas drawn through the measured

15、pressures, at-stations 0.924 and 0.950parallel to the theoreticalboattail pressure distribution. uIntegrating the pressure distributions results in the curve of theboattail drag coefficient shown in figure (b). The method of character-istics yielded drag coefficients that were consistentlyhigh, 15 p

16、ercentfor model 2 and 16 percent for model 3.Base pressures.- Measured base pressure-coefficientsare presentedin figure 6 as a function of boattail angle. Base pressure coefficientsdetermined by the methods of references 1 10 exe also shown for pur-poses of comparison. The method of reference 1 gave

17、 excellent agreement(within percent) with the present test results, whereas the method of- - -reference 10 indicated correctly the increase in base pressure withincreasing boattail angle but predicted base pressures considerably higherthan the measured values. .The base pressures measured in the pre

18、sent tests were lower thanmost of the available data. The present tests were conducted at a rel- *-atively high Reynolds number, however, with a turbulent boundary layerobtained from natural transition; whereas most other estigations havebeen conducted at a lower Reynolds number tith either natural

19、or artifi-cial transition. Either natural transition at a lower Reynolds numberor an artificially induced transition would tend to produce a thickerturbulent boundary layer, at the base, with an accompanying increase inbase pressure.Several investigations (for example, ref. l-l)have sho that tif-cia

20、l transition produces base pressures 5 to 10 percent higher than thatfor natural transition, the larger differences being at the lower Machnumbers. It has also been shown many times (for example, ref. 7) thatthere is a decrease in base pressure with increasingReynolds number,when the boundsry layer

21、just ahead of the base is turbulent. Applicationof these corrections,where applicable, results good %reement betweenthe present data and existing data. ,-Another factor which might affect the base pressure i.s.thepresenceof the supporting strut. This strut is 6.2-percent thick in the stresm-wise dir

22、ection and is tapered from a J-inch chord at the model to a10.5-inch chord at the base. At the model, the trailing edge is 1: chordsforward of the base. Although not strictly applicable, because of the .Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,

23、-NACA RM L51+c16 7taper snd sweep of the strut, the analysis and data of reference E indi-cate that the effect of the strut on the base pressure would be verysmall. This is in agreement with the tests of reference 13 in which therearwsrd position of the side strut closely approximates the conditions

24、of the present tests. At the higher Reynolds numbers used in the refer-ence tests, the curves of measured and interference-freebase drags con-verge. The side support strut is therefore believed to have had only asmall effect, it any, on the results of the present tests.Afterbody drag.- Combining the

25、 measured boattail snd base dragsyields the power-off afterbody drag coefficients shown in figure 7 as afunction of afterbody fineness ratio. Increasing the afterbody finenessratio from O to 1.92 results in a 50-percent reduction in afterbody dreg,and further increases in afterbody fineness ratio wi

26、ll result in furtherdecreases in the afterbody pressure ,whereas, on model 2-,thereduction was only 1 percent of its power-off drag. Inasmuch, as thepower-off boattail drag on model 3 is more than twice that on model 2,however, the mme favorable interference effects of the jet still do not wsrrantit

27、s choice-from a drag standpoint.In anffort to gain a feel for the effect of the ratio of jet tofree-stream Mach number, a fourth nozzle (A = 11) was tested in model 3.This nozzle had the same exit =ea as the other nozzles, but had a larger “throat so that the exit Mach number was 2.16 compared With

28、2.65 fornozzles 1, 2, and 3. Thus, for a given jet pressure ratio, the mass flowwas less from this nozzle than from the other three.The interference effects of-the M = 2.16 Jet on the boattall pres-sures and drag of model 3 (I3= 10o), may be seen in figures 8(c) and 9,respectively. As in the other c

29、ases where the boattail pressure distri-bution was disturbed by the propulsive jet, only the orifice closest tothe base registered any change from its power-off reading. In thiscase, _,this orifice registered an increase when the Jet pressure ratioexceeded 0.8. When nozzle 2 (A = l.lo)was tested in

30、this model, therewere no interferenceswith the boattail pressure distribution even atthe highest jet pressure ratio.Apparently then, there is an increasinginterference from the Jet as the ratio of jet-to free-stream-Machnumberis decreased. This trend was also noted in reference 2 at a free-stream -”

31、 -Mach number of 1.92”.A comparison of the interference effects.fromnozzle 3 ( = 22)and nozzle 4 (X = no) is given in figure 9. At jet pressure ratiosnesr the ideal pressure ratio of 1.0, a greater drag reduction is avail-able from the jet of lower Mach number even though its divergence singleis but

32、 half that of the jet of higher Mach number. Above a jet presseratio of 1.6, the greater Jet expansion from nozzle 3 results in greaterJet interference on the boattail and consequefitlya greater pressure dragreduction. However, one might surmise that an even larger boattail dragreduction might be av

33、ailable if the divergence angle of nozzle 4 were 22instead of no.A comparison of nozzles 3 (h= 220) and 4 (A = ll.”)on the basisof the ratio of the total pressure to the free-stream static pressurein figure.10shows that the nozzle of lower Mach number (nozzle 4) alwaysproduced t lsrger boattail drag

34、 reduction._ This lsrger boattil bagreduction was accomplished despite the fact that the divergence angleof the nozzle was but half that of the nozzle with the higher Mach number.Base pressure.- Base-pressure variationswith jet pressure ratio -.are shown in figure.11. Power-off base-pressure coeffic

35、ients are shown,for purposes of compariscaq:-.at%h”io of power-off base pressure to .Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L5M16.9.free-stresm static pressure. Inasmuch as the jet-exit pressure was com-puted from the measured combus

36、tion-chsmberpressure, no attempt was madeto correlate the data while the nozzle flow was h the separated condi-tion. Thus, the curves of figure 11 begin at the petit where the nozzleflow attaches.At this flow-attachingpressure, the base pressures of all of themodels are higher than their resective p

37、ower-off values. For model 1,the increase in base-pressure coefficient was 0.025, whereas for models 2and 3 the increase was approximately 0.04. As the jet pressure increasedfrom the flow-attaching condition, the base pressure decreased until for”model 1 it was 0.175 less than the power-off value. F

38、or models 2 and 3,the base pressures had decreased to approximately 0.125 less than theirrespective power-off values. These reductions in base pressure amountto increases in the base annulus drag of 192 percent, 209 percent, and237 percent on models 1, 2, and 3, respectively. Further ticreases inthe

39、 jet pressure ratio result in an increase in base pressure.For model 1, the jet pressure ratios of the present tests were nothigh enough to result in the base pressure ever returning to its power-off value regardless of the nozzle half-angle. With models 2 and 3,however, the base-pressure increase w

40、ith jet pressure ratio was suffi-cient to raise the base pressure to equal or exceed its power-off valueby a pressure ratio of 2 for all nozzle half-angles. When nozzle 3(h = 22) was used in models 2 and 3, the base pressure returned to itspower-off value at a jet pressure ratio slightly above the i

41、deal pressureratio of 1.0.As shown in figure 12, nozzle 4 produced the S= base-pressuretrends with et pressure ratio as had the other nozzles. Shown also, forcomparison purposes, are two curves from figure n(c). At a given jetpressure ratio, the highest base-pressure coefficient was obtained withthe

42、 highest nozzle divergence angle at the higher jet Mach nuniber. How-ever, because of this difference in jet Mach numbers, the total pressuresof the two jets would be very different. Figure 13 illustrates the morepractical case where an engine produces a given jet total pressure and thechoice of an

43、exhaust nozzle must be made. Viewed in this manner the lowerMach number et induces considerably less drag than the higher Wch ntierjet with the same divergence. It is also superior (from a drag standpoint)to the higher ch number jet with twice the nozzle divergence.The physical phenomenon which resu

44、lts in these large pressure changesmay be seen in the shadowgraphs presented as figure 14. The lsrge dropin base pressure between the nozzle starting pressure ratio and 0.8 isdue to the aspiration or ejection effect of the propulsive and externalflows on the low-energy boundary-layer air which flows

45、 into the “deadair” region around the annulus. Increasing the jet pressure increasesthe ejection of the air from the dead-air region. Because the flow intothis region is not incre the external and the jet flows must turnEEProvided by IHSNot for ResaleNo reproduction or networking permitted without l

46、icense from IHS-,-,-10 IIMNumL NACA RM L134C16.more shsrply towards the dead-air region with the result that the wakeshock moves closer to the base. This increase in turning angle of theexternal flow increases its expansion and hence results in a decrease a71In the base pressure. At the ssme time, t

47、he wake shock becomes strongeras it moves towards the base. The increasing presstie gradient from thewake shock slows down the ejection action and the expanding jet beginsto compress the dead-air region. In order to equalize the pressure inthe external stream, the expansion from the boattail is redu

48、ced. When,because of the increase h jet pressure, the external flow over thedead-air region has the ssme inclination to the kiy center line as theboattail (that is, no expansion at the base), a further increase in jetpressure will result in a compression of the external flow at the baseand the appea

49、rance of a lambda leg ahead of the main wake shock, as shown in figure 14(b). Further increases in the jet pressure result in thestrengthening of these shocks and the continuation of their forwardmovements.Afterbody drag.- Combining the measmed power-on boattail and basedrags results in the curves of afterbody drag coefficientspresented infigure 15. Each set of curves is for a constant jet pressure ratio.Also included in this