FUSELAGE-MOUNTED FINS ON THE STATIC DIRECTIONAL STABILITY. By M. Leroy Spearman, Ross B. Robinson, and Cornelius Driver

Transcription

1 RESEARCH MEMORANDUM THE EFFECTS OF THE ADDITION OF SMALL FUSELAGE-MOUNTED FINS ON THE STATIC DIRECTIONAL STABILITY CHARACTERISTICS OF A MODEL OF A 45' SWEPT-WING AIRPLANE AT ANGLES OF ATTACK UP TO 15.3? AT A MACH NUMBER OF 2.1 By M. Leroy Spearman, Ross B. Robinson, and Cornelius Driver Langley Aeronautical Laboratory Langley Field, Va. 1 f NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS m WASHINGTON October 12, 1956

2 ~ 1B rl NACA RM L56Dl6a NATIONAL ADVISORY COMI- FOR AERONAUTICS 'J RESMCH MEMORANDUM THE EFFM='lS OF TRE ADDITION OF SMALL FUSELAGE-MOUNTED FINS ON THE STATIC DmTIONAL STABILITY CHARACTERISTICS OF A MODEL OF A 45' SWEPT-WING AIFPLCWE AT ANGLES OF ATTACK UP To 15.3' AT A MACH NUMBER OF 2.1 By M. Leroy Spearman, Ross B. Robinson, and Cornelius Driver SUMMARY k 1 d Tests have been made in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 2.1to determine the effects of the addition of four small'fuselage-mounted cruciform fins on the directional characteristics of a 45' swept-wing airplane model at angles of attack up to 15.3' and angles of sideslip up to about 16'. The results showed that the addition of the four cruciform fins to the model increased the directional stability substantially at the highest angle of attack and, at the same time, caused relatively small changes in drag. I"I!RODUCTION 1. d Among the problems of supersonic aircraft is that of maintaining sufficient directional stability, particularly at high angles of attack. For this reason, tests have been made in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 2.1 of a 45' swept-wing fighter-type airplane model to which four small cruciform fins were added for the purpose of augmenting the directional stability. The fin arrangement investigated was suggested by the Republic Aviation Corporation. The fins, which extended over the rear 31percent of the fuselage, were about the size and shape of a conventional ventral fin and were located in a cruciform arrangement in planes at 45O to the horizontal

3 and verticalplanes. The fins had a combined area of 76.6 percent of the vertical-tail area. The fins were tested in conjunction with a normal vertical tail and ventral fin. Ln addition, various cambinations of the cruciform fins, the ventral fin, and the vertical tail were made. All tests were made with the wing on and with air flow through the wingroot inlets. SYMBOLS The results are presented as coefficients of forces and moments with the moment reference point located at 25 percent of the wing mean geometric chord. The yawing moment, rolling moment, and side force are referred to the body-axis system and the lift, drag, and pitching mament are referred to the stability-axis system. The symbols are defined as follows : Lift lift coefficient, - qs rag (approximate) drag coefficient (approximate), qs equivalent to true drag at p = Pitching mament pitching-mament coefficient, Yawing moment yawing=mament coefficient, qsb Rolling moment rolling -moment coefficient, q=.p I side-force coefficient, Side force qs free-stream dynamic pressure wing area,.795 sq ft wing chord wing mean geametric chord,.522 ft wing span, 1.59 ft 4 V

4 P a C C 2P cyp F V U angle of sideslip, deg angle of attack, deg rate of change of yawing-mament coefficient with angle of sideslip (p = ), &n/ap rate of change of rolling-moment coefficient with angle of sideslip (p ), &ZPP rate of change of side-force coefficient with angle of sideslip (P = 1, acy/ap cruciform fins vertical tail ventral fin MODEL AND AF'PARADJS?. d I. d A three-view drawing of the model is shown in figure 1. Details of the four cruciform fins are shown in figure 2. Photographs of the model showing the cruciform fins installed are presented in figure 3. Some geometric characteristics for the model are presented in table I. The model had a wing with 45' sweepback at the quarter-chord line, an aspect ratio of 3.2, and a taper ratio of.468. The wing was composed of NACA 6%-series sections having thickness ratios of 5.5 percent just outboard of the inlet (.38bj2) and 3.7 percent at the tip. The wing was located slightly above the fuselage center line and the horizontal tail was mounted below the fuselage center line. The stabilizer incidence angle was fixed at -3'. The model was equipped with twin wingroot supersonic inlets ducted to a single exit at the base of the body. All tests were made with air flow through the ducts with a mass-flow ratio of about.8. The four cruciform fins were attached in planes at 45' to the horizontal and vertical planes and were arbitrarily shaped and positioned so that their span would cause no increase in the projected side area of the afterbody. As indicated in figure 2(a), the lower fins were slightly asymmetric at the model base. Forces and moments were measured by means of a six-component internal strain-gage balance.

5 4 NACA RM L56Dl6a TESTS Test Conditions and Procedure It t The tests were made at a Mach number of 2.1, a stagnation pressure of 5 pounds per square inch absolute, and a stagnation temperature of LOOo F. The dew point was maintained sufficiently low (below -25' F) so that no condensation effects were encountered in the test section. The Reynolds number based on a mean aerodynamic chord of.522 foot was about 644,. Tests were made through an angle-of-sideslip range fram about -4' to about 16 at angles of attack, primarily, of Oo and 15.3O. CORRECTIONS AND ACCURACY The angles of attack and sideslip have been corrected for deflections of the balance and sting under load.. The drag coefficients have been adjusted to correspond to the condition of free-stream static pressure at the model base. The measured drag was corrected to account for the internal drag and for an internal buoyant force on the balance so that a net external drag was obtained. The estimated errors in the individual measured quantities are as follows : CL... ia.73 CD'... f.11 c,.... fo.23 c,.... m.3 c p.... ~.2 cy..... ~.2 u,deg... to. 1 p,deg * PRESENTATION OF RESULTS The results are presented in the following manner: tail... Figure 4 L!! V

6 Effects of top pair and bottom pair of fins at a = 15.3', vertical tail and ventral fin on... Effects of top left and bottom right fins at a = 15.3O, vertical tail and ventral fin on... Effects of ventral fin at a = ' and 13.3', vertical tail and four cruciform fins on... Effects of ventral fin and four cruciform fins with the vertical tail off, a = Oo and 15.3'... Effects of four cruciform fins on lift, drag and pitchingmoment variations with sideslip at a = ' and 1>.3', vertical tail and ventral fin on... Summary of sideslip derivatives for various fin and tail arrangements... Summary of directional stability increments provided by vertical tail, cruciform fins, and ventral fin... DISCUSSION Directional Characteristics Figure g Effects of cruciform fins.- The effects of the addition of the four cruciform fins to the model are presented in figure 4 and summarized in figure 1. The addition of the four fins to the cmplete model (vertical tail and ventral fin on) caused only a slight increase in directional s t abi lity at a = '. However, the directional stability was ( ".e) increased significantly at a = 15.3' where a reflex in the variation of C, with p near p = Oo was removed by the addition of the four fins. (See fig. J+(c).) The addition of either the top pair or bottam pair of fins separately at a = 15.3' provides about half the increase in directional stability as that provided by all four fins. (See figs. 5 and 1.) "he nature of the cruciform fin effect on the directional stability at a = 15.3O may be seen by comparing the configuration without fins with the configuration equipped with the top left and bottom right fins (fig. 6). At p = ", the inclination of these fins is such that a component of force is directed in the positive side-force direction and provides a negative yawing moment. Throughout the negative sideslip range, these fins remain in a plane inclined to the stream direction and provide an increased stabilizing increment of yawing moment. Through the positive sideslip range, however, these fins tend to move into the plane of the relative stream direction and thus provide essentially no

7 change in Cn above about p = 2'. Thus, it is apparent that the upper and lower fins on both sides are required to provide the maximum effectiveness in increasing the restoring yawing moments throughout the sideslip range. 1 c The variations with angle of attack of the incremental cnp contri- butions for the various configurations (fig. 11) indicate a rather rapid decrease in effectiveness of the vertical tail. The ventral fin effectiveness also decreases with angle of attack but to a lesser extent. The effectiveness of the four cruciform fins, on the other hand, increases with increasing angle of attack. All tests were made with the wing-root inlets open. Any effects of these inlets on the directional characteristics are believed to be small, however, since no significant disturbance from the inlets in the region of the tail was visible in schlieren or vapor screen views of the model. Effects of ventral fin.- The complete-model results thus far have included a conventional ventral fin on the body center line as a part of the vertical stabilizing surfaces. Since ventral fins may cause some restriction to the ground clearance angle, it would be desirable, if practical, to eliminate such surfaces. Accordingly, some tests were made at a = Oo and 15.3' with the ventral fin removed but with the four cruciform fins and the vertical tail on. These results (fig. 7) indicate some reduction in directional stability at a = Oo, but the level of stability is sufficiently high at this angle so as not to be critical even without the ventral. At a = 15.3', the removal of the ventral fin had very little effect on the directional stability, particularly at small sideslip angles where there is essentially no change in the. Thus, it appears that it would be practical to remove the CnP ventral fin without any significant effect on the directional stability. The effects of the ventral fin and of the four cruciform fins were also determined at a = ' and 13.3' for the configuration with the vertical tail removed. These results (fig. 8) indicate that the increment in C, provided by the ventral alone is about half that provided by the ventral fin and the four cruciform fins combined. Lift, Drag, and Pitchingdment Characteristics There is no appreciable change in the variations of CL, CD', or C, with p resulting from the addition of the four fins (fig. 9). The drag-coefficient increment due to the fins at a = ' is about.2. At a = l5.3', the increment indicated in CD is probably a result of 4 Y

8 .. *.*. NACA RM L36Dl6a 7 the increment in CL change in drag. so that for a constant lift there would be little Lateral Force and Effective Dihedral The effect of the cruciform fins and the ventral fin was to increase the lateral-force derivative Cy in a manner consistent with the attend- P ant directional-stability changes (fig. 1). There was no significant change in the effective dihedral at a = ' (fig. lo), although the value obtained at a = l5.3o was somewhat greater with the fins on than with the fins off. CONCUTSIONS Tests made to determine the effects of the addition of four small fuselage-mounted cruciform fins on the directional characteristics of a 45' swept-wing airplane model at a Mach number of 2.1 indicated the following conclusions: 1. "he addition of the four cruciform fins to the model, although it caused only a slight increase in directional stability at zero angle of attack, provided a substantial increase in the directional stability at the highest angle of attack (15.3') primarily because of direct forces on the fins. 2. The effectiveness of either the upper pair or lower pair of fins alone was about half that for all four fins together. 3. Removing the conventional ventral fin had no significant effect on the directional stability of the configuration with the four cruciform fins. 4. The addition of the four cruciform fins caused an increase in the drag coefficient at zero angle of attack of about.2. Langley Aeronautical Laboratory, National Advisory Committee for Aeronautics, Langley Field, Va., April 3, 1956.

9 NACA RM L56D16a a TABU 1.- GEx)ME'IRIC CHARACTERISTICS OF MODEL Y Area. including body intercept. sq ft span. ft Mean geanetric chord. ft Aspect ratio Taper ratio Sweep. quarter-chord line. deg Dihedral. deg Theoretical root chord at body center line. in Theoretical tip chord. in Airfoil sections: 13.8-percent semispan station NACA 65~5.5 Theoretical tip... NACA 65~3.7 wing:... span. ft... Aspect ratio... Taper ratio line. in Airfoil sections: Root... Tip... Horizontal t ai 1: Area including body intercept. sq ft Sweep. quarter-chord line. deg... Theoretical root chord at body center Theoretical tip chord. in NACA 65~6 NACA 65AOO4 Vertical tail: Area to body center line. sq ft span. ft Aspect ratio Taper ratio Sweep of leading edge. deg Theoretical root chord at body center line. in Theoretical tip chord. in Ventral: Area (exposed). sq ft Fuselage fin: Area one. exposed) sq ft Area t four. exposedj. sq ft r,

10 IU 9 y! rn

11 1 NACA RM L56D16a c

12 c *' 4,

13 12 NACA REI L56D16a c I

14 iyi A '

15 14 NACA RM L56D16a.8.m.4.a c, O 4.I CY ' -;I A (a) a 5 QO. Figure 4.- Aerodynamic characteristics in sideslip for various fin arrangements at several angles of attack.

16 .w -.2 c, - I -.2.I CY -.I c J P, deg (b) u = 1'. Figure 4.- Continued.

17 16.6 Cn SM m On On On On A Off Off U Cl I CY -.I P,W (c) u = 15.3'. Figure 4.- Concluded.

18 3B a? Cn -2 1 c, -.o I -2-3.I CY -.I P, deg Figure 5.- Effect of top pair and bottan pair of fins on sideslip characteristics with vertical tail and ventral fin on. a = 15.3O.

19 18 NACA RM L36D16a.O 4.o 2 8 J Cn o c, 3 -.o I.2-2.I -3 CY -.I , deg Figure 6.- Effect of top left and bottom right fins on sideslip characteristics with vertical tail and ventral fin on. a = 15.3O.

20 I.6.w Cn c, - I -.2.I CY -.I -.2 Y a -4 4 a 12 is 2 P, deg (a) a = '. Figure 7.- Effect of ventral fin on sideslip characteristics at two angles of attack with the vertical tail and four cruciform fins on. \ i

21 2.a.4 Cn.w -2 DI c, - I -2.I CY O -.I a P, deg (b) u = 15.3'. Figure - 7, Concluded.

22 21 '.2 V c,o -.o I c.i c cv -.I P, deg (a) a = '. Figure 8.- Effect of ventral fin and four cruciform fins on sideslip characteristics at two angles of attack with the vertical tail off.

23 22 NACA RM ~36D16a.6 r'.2 Cn -.2 Cl -.o I -.2.I CY -.I a -4 4 a P, deg (b) u = 13.3'. Figure 8.- Concluded. _._

24 NACA RM L36D16a moo P, deg (a) a = '. Figure 9.- Effect of four cruciform fins on lift, drag, and pitchingmoment variations with sideslip at two angles of attack with the vertical tail and ventral fin on. d

25 24 (b) u = 13.3'. Figure 9.- Concluded. L

26 4B. W Vertiwl on on Off Off On On m Off tail c , deg Figure 1.- Summary of sideslip derivatives for various fin and tail arrangements.

27 26 b.4 (AC 2.2 F U on on I off on -- J n- KI -- (ACn )U P.2 I ' Ktfftftttf Ventral fin contribution I U v on off -- Figure ll.- Incremental values of Cn for various configurations. B NACA - Langley Field, VJ