NASA NACA-TR-1263-1956 Investigation of the aerodynamic characteristics of a model wing-propeller combination and of the wing and propeller separately at angles of attack up to 90 .pdf

1、*s- 0 1-s. . - , , .,I , ._. -. -. , /- z- ., i- - . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-ERRATA By Richard E. Kuhn and John W. Draper 19% Page 9: The next to last equation in column 1 should be corrected to read as follows: Page 23, fig

2、. 13: The equation in the figure legend should be corrected to read as follows: ;= cmpE IIfl 3 Tc ED The following typographical errors should also be corrected: Page 4, col. 1, line 1: “ressure“ should be “pressure“. Page 9, col. 2, line 3: “here“ should be “where“. NACA - Imgley Field, Va. Provide

3、d by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.- - REPORT 1263 INVESTIGATION OF THE AERODYNAMIC CHARACTERISTICS OF A MODEL WING-PROPELLER COMBINATION AND OF THE WING AND PROPELLER SEPARATELY AT ANGLES OF ATTACK UP TO 90” By RICHARD E. KUHN and JOHN W. DRA

4、PER Langley Aeronautical Laboratory Langley Field, Va. 1 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-National Advisory Committee for Aeronautics Headquarters, 1512 H Street NW., Washington 25, D. C. Created by act of Congress approved March 3, 19

5、15, for the supervision and direction of the scientific study of the problems of flight (U. S. Code, title 50, sec. 151). Its membership was increased from 12 to 15 by act approved March 2, 1929, and to 17 by act approved May 25, 1948. The members are appointed by the President, and serve as such wi

6、thout compensation. JEROME C. HUNSAKEB, SC. D., Massachusetts Institute of Technology, Chairman LEONARD CARMICHAEL, PH. D., Secretary, Smithsonian Institution, Vice Chairman JOSEPH P. ADAMS, LL.B., Vice Chairman, Civil Aeronautics Board. ALLPN V. ASTIN, PH. D., Dire however, deviations are assumed t

7、o be relatively small and corrections to be fairly accurat,e for most of the test condi- tions. The jet-boundary corrections applied to the angle of attack and longitudinal force were estimated by the method of reference 1. For a given model size, these corrections depend on the circulation about th

8、e wing; therefore, the corrections for a particular angle of attai however, to include these effects would needlessly complicate the presentation. For t.hc purpose of presenting coefficients, therefore, the dynamic prcssurc in the slipstream can be defined as and the thrust cocfficicnt as or T,” T (

9、2) i Dp+T a so 1 $=(7R of 20 for a thrust coefficient of 0.20. The choice of hladr angle for WC at, zero forward speed (vertical take-ofl or landing) cannot bc made, however, on the basis of the cffi :“4 1: :8 32 .94 .96 1:P - 1 - T.” 1 :2 :Z .50 2: .20 :A: .06 .04 . 02 1 ,949 ,894 ,837 :a: ,632 ,54

10、8 :K ,283 :%? 0 -. 051 -. 106 -. 163 -. 226 -.293 -. 368 -.452 -. 553 -. 694 -. 717 -. i55 -. 800 _ I 141 I -.taJ 0 0 -1.m TABLE I FUNCTIONS OF T,” ,h - T,“-1 / 1+.,/1-T,” 2 1.949 1.894 1.837 1.774 1.707 1.632 1.545 1.447 1.316 1.263 1.245 1.200 1.141 1.000 The maximum static-thrust efficiency of 0.

11、7 shown in figure 7 (diamond symbol) was obtained with a blade angle of 8”. With the propeller disks overlapped, the static-thrust effi- ciency was reduced to 0.65 as indicated in figure 8. A corre- sponding reduction in efficiency at forward speeds is indi- cated in figure 8. This loss in static-th

12、rust eRicicncy with the propeller disks overlapped does not necessarily mean t,hat overlapping is undesirable, however, because, for an airplane of a given size and with a given number of propellers, overlapping permits the use of larger diameter propellers, which can result in an incareasc in stati

13、c thrust. for a given horsepower, even though the efficiency is reduced somewhat by overlapping. The variations of the propeller thrust coefficient C, and power coefficient Cp with angle of attack are presented in figure 9. It shoulcl be remembered when use is made of these data that the thrust was

14、held constant throughout. the angle-of-attacli range ancl the rotational speecl and power were allowed to decrease with increasing angle of attack. In general, the data for the isolated propeller show somewhat lower values of CIT and C, than the data, for the propeller 01 propellers operating in the

15、 presence of the wing. The biggest differences, however, occur under conditions that are not likely to be of practical interest (high forward speed (T,“=O.2) at high angles of attack). The corresponding variations of v cos a are presented in figure 10. nD The power required for a constant, thrust co

16、ndition through the angle-of-attack range is presented in figure 11. In general, the power decreased as the angle of attack increased. The normal-force and pitching-moment coefficients of the. outboard propeller are presented in figure 12. Similar data for the inboard propeller are not presented bec

17、ause of difi- culties experienced with the instrumentation for the inboard propeller that resulted in excessive scatter arid large shifts in the zero readings. The general trend of the data, however, was similar to that for the outboard propeller. The problems of obtaining reliable data were conside

18、rably increased be- cause the strain-gage beams, which measured the normal force and pitching-moment loads, were also required to sup- port, the relat,ively heavy motor and carr- the high thrust and torque loacls. Also presented in figure 12 are the tdleoretical variations of normal force obtained b

19、y the method of reference 3. ThC tlieoretical variation of propeller normal force with angle of - 1 - - 0 ,051 106 ,163 226 .293 365 ,452 553 : 654 : 2 ,800 ,859 1.000 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.i-. AERODYNAMIC CHARACTERISTICS O

20、F A WING-PROPELLER COMBINATION AT ANGLES OF ATTACK UP TO 90 5 attack of reference 3 is intended to be applicable only at Wing characteristics.-Figure 16 presents the variation angles of attack near zero. The additional factors which with angle of attack and thrust coefficient of the lift co- contrib

21、ute to the normal force at high angles of attack can- efficient based on the dynamic pressure in the free stream. not readily be included in the theoretical treatment. It is Eigure 17 (a) presents the same data based on the dynamic interesting to note, however, that, for the configuration of pressur

22、e in the slipstream. The lift variation for a thrust this investigation, calculations of the normal-force coefficient coefficient of 1.0 (dashed line of fig. 17) cannot be presented using the q-factor (which accounts for the inflow to the pro- in figure 16 because, if the free-stream dynamic pressur

23、e were peller) based on the component of thrust in the free-stream used to obtain the coefficient, the lift coefficient at all angles of direction rather than in the thrust direction, as assumed in attack would be infinite. The disadvantage of basing the reference 3, show relatively good agreement w

24、ith the meas- coefficients on the free-stream dynamic pressure is thus ured data. readily apparent. The operation of both propellers in the presence of the wing is seen almost to double the pitching moment of the outboard propeller as compared with that of the propeller alone. .This magnitude of inc

25、rease cannot be attributed to an increase in wing-induced upwash at the propeller disk, because such an increase should produce corresponding in- creases in propeller normal force. It is probable that these increases in pitching moment are due to a change in the velocity through the upper and lower

26、portions of the propeller disk (as referenced t,o the wing-chord plane). An increase in velocity over the wing (upper part of the propeller disk) 1voultl tend t,o decrcnse the thrust from the top part of tllc disk. Conversely, a dccrcase in velocity through tllc lowcl half would increase the thrust

27、of this part of t;le propeller; thus an increase would occur in the nose-up pitching moment of the propeller with increasing angle of attack. Increasing the thrust coefficient, with either one propeller (fig. 18) or two propellers (fig. 17), results in an increase in the angle of atttick at which ma

28、ximum lift is reached and a more gradual variation of the lift with angle of attack above maximum lift. It should be remembered that these results are for constant thrust throughout the angle-of-attack range. If the power were held constant as the angle of attack was increased, the thrust would incr

29、ease with angle of attack ancl an even more gradual variation of lift above maximum lift would be indicated. The propeller pitchingrnoment can be regarded as being clue to the fact that thrust of the propeller is appliecl at some dis- tance from the center of rotation. The effective raclial lo- cati

30、on of the thrust vector is presented in figure 13 and was dctcrmined from the pitching-moment data of figure 12 by the following relation The data at T,“=O (figs. 17 and 18) were obtained with the propellers removed. Data are compared in figure 19 fol conditions of propeller removctl, of zero thrust

31、 with the pro- peller on, and of propeller windmilling. Removing the pro- peller results in a small reduction in lift in the region of maxi- mum lift,. As woulcl be expected from the propeller data discussed previously, removing the propellers appreciably decreases the unstable variation of pitching

32、 moment with angle of attack at the low angles of attack. These effects should be kept in mind when the propeller-of? pitching- moment clata of figures 17 and 18 are used. The data of figure 20 indicate, as might be expected, that the narcllrs disturb the flow over the wing so that the wing with nac

33、elles stalls at a lower angle of attack and lower lift coefficient than the wing alone. As would be expected, the nacelles also cause a marked decrease in the static longi- For the most extreme condition, the effective location of the thrust vector is seen to move downwarcl more than 20 percent of t

34、he propeller radius. (See sketch, fig. 13). The significance of these propeller pitching moments can be juclgecl from figure 14, which presents the total contribu- tion of both propellers to the total model pitching moment. Because the data on the inboard propeller was unreliable, the data obtained

35、for the outboard propeller was used for both the inboarcl and the outboard propellers in the summa- tion represented by ACm”. The calculatecl variation was obtained by using the calculated values of normal-force coefficients (fig. 12) and letting the pitching-moment co- efficient be zero. It can be

36、seen that the usual procedure of basing the propeller contribution only on the propeller nor- mal force accounts for less than half of the total contribution for this configuration at these thrust coefficients. Aclditional data on the normal force and pitching moments of isolatecl propellers are pre

37、sented in references 4 and 5. The contributions of the spinner (when not rotating) to the propeller normal force and pitching moment are presented in figure 15. tuclinal stability ( ,C,” increase in bar below the stall. The destabilizing effects of the propellers and nacelles at zero angle of attack

38、 are summarized in figure 21. These data show that, near zero angle of attack, there is only a small ac Ir variation of stability, as indicated by the curve of + L against thrust coefficient. Estimation of the lift-curve slope.-Reference 6 presents relationships for estimating the increase of lift d

39、ue to a slipstream flowing over the wing. Rearranging the relation- ship to provide an espression for the lift-curve slope of the wing with slipstream and using the notation of the present report gives the following equation: (1-T;) l+$ (F-1) ) For the present configuration, X can be taken as 1.0 (r

40、ef. S), Provided by IHS Not for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 REPORT 1263-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS is obtained from equation (B5) of appendix B, 8 8 -=- c-t 4 is obtained from equation (B7) of appendix B, and Oaf3 3% ( ( (acL,L),/., =

41、 The lift-curve slope can be expressed as MI?,” bCL -= _ aff ( act Tc,=o(l-Tc ) 1+- ” (;g)u+li, (6) where dl can be obtained from equation (B4) of appendix B. Calculation of the lift-curve slope by this equation under- estimates the measured lift-curve slope (fig. 22). If it is assumed that the incl

42、ination of the slipstream is 7 equation (5) reduces to (7) and much better agreement with the experimental data is obtained (fig. 22). This equation, in effect, is based on the assumption that the circulation around the wing is un- changed by the presence of the slipstream and the increase in lift i

43、s directly proportional to the increase in velocity across the circulation. If it is further assumed that the wing is far enough behind the propeller so that the full slipstream velocity is developed (K= 1) and that the wing is fully immersed in the slipstream 9 the equation reduces to the simple fo

44、rm T,“=O (8) According to equation 7, the fact that the lift-curvt slope obtained for the configuration with two propellers is higher than that obtained for the configuration with one propeller is primarily clue to t.he greater percent,age of wing area that is immersed in the slipst,ream. PERFORMANC

45、E ESTIMATES The proceclure outlined in appendix C was used to esti- mate the performance of a hypothetical four-propeller air- plane as it traversed the regime of flight represented in figure 6. The hypothetical airplane was assumed to have linear climensions 6 times those of the model. Calculations

46、 were made for standard sra-level conclitions, and the fuselage and other parts of the airplane not represented by the model were assumecl to have a drag coefficient of 0.01. The variation of the thrust coefficient required ancl the forward velocity reached in constant-altitude transition, as the wi

47、ng attitude is lowered from 90 for take-off to convcn- tional flight attitude, is shown in figure 23. The corre- sponding variation of thrust power requirecl with forward velocity for an assumed wing loading of 40 pounds per square foot is shown in figure 24. The thrust power required is seen to dec

48、rease quite rapidly in the low speed range. The mini- mum thrust power required occurs in the normal flight range at a speed of 160 mph and a wing attitude of 8.5. It will be noted that (with the assumption that =0.75 at high speed and $=0.65 for static-thrust-take-off), if this airplane were design

49、ed for a high speed of the order of 350 to 400 mph, sufficient power would be available for vertical take-off with the 12-foot-diameter propellers represented by the ones used on this model. The untrimmed pitching moment and corresponding eff ec- tive aerodynamic-center location, as indicated by g for the conditions of this analysis, are indicated in figute 25. No allowance for the effects of trimming