1、:i. a 4 I c V a p? 1; i ii;,- ? ! f i” I I, i; 1 ;: /i; COPY 6 RM L56E2! ? ! “j;-I Y” Langley Field, Va. I. ., .,“.+ . . . . : -: , _- I * ; . k7L: C,L.L.I.e., Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . - 4 NACA FBI L56Eztg 2t 2v q P VO M S
2、% sh C namely, normal force, yawing moment, ,and root bending moment. With the horizontal tail mounted in the high position, measurements of the rolling moment of the horizontal tail about its juncture with the vertical tail also were made. A small gap existed between the juncture of the vertical ta
3、il and fuselage and was closed with a sponge seal. Some model load testi were made both with a sponge and with a solid seal to determine leakage effects. CORREETIONS Jet-boundary corrections to the angle of attack were applied in accordance with reference 8. The jet-boundary corrections to the later
4、al force, yawing moment, and rolling moment were considered negligible and therefore were not applied. From past experience, it was found that tares due to sting support were negligible; therefore, these values were not applied. Blockage corrections were applied to the data by the method outlined in
5、 reference 9. The angle of attack and angle of sideslip have been corrected for deflection of the sting support and balance system under load. No attempt has been made to correct the data for aeroelastic distortion of the model; however, based on past experience, it is believed these cor- rections a
6、re negligible. RFSULTS AND DISCUSSION Presentation of Results Basic model data: Figure CLagainsta . 5 Cz,s, Cn,s, C!y,s against p 6 to g C!zp, Cnp, CyP against a . 10 and 11 Basic tail loads data: BVagainstCNV . Cnv against Cy+ . CNVagainstp C2hagainstp cBvJ cnvJ cNv against p . cBvP, CnvP, cNvP, Cx
7、hp against a 12 13 14 15 16 17 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L56E2g c- Data related to analysis of results: ACy j ACnp) AC, P P against a . . . . . . . . aCNV - against M . . . . . . . . . . . . . . . ap acNv ap against a .
8、. . . . . . . . . . . . . . against M . . . . . . . . . . . . . . $ cv X against a . . . . . . . . . . . . . . 9 . . . . . . . . . -E 19 20 and 21 22 . i 23 t0 25 The basic model data (figs. 5 to 11) are presented about a stability system of axes as shown in figure 1 and the coefficients have been b
9、ased on the model wing area, span, and mean aerodynamic chord with moment reference at the quarter chord of the wing mean aerodynamic chord. The static lateral- and directional-stability derivatives (figs. 10 and 11) were obtained from tests at angles of sideslip of -4 and 4 through the angle-of-att
10、ack range; however, sideslip tests (in a range from -4O to approximately 12O) were made at several selected angles of attack for several model configurations (figs. 6 to 9). The basic vertical-tail loads data and the rolling-moment coeffi- cients (figs. 12 to 17) of the high horizontal tail about th
11、e point of attachment to the vertical tail are based on the area, span, and mean aerodynamic chord of the vertical tail and the area and span of the horizontal tail as given in table I. These data are about a body system of axes fixed in the model as shown in figure 1. The vertical-tail area is an a
12、pproximate exposed area and is defined as that area included above a root chord that is slightly inside the fuselage; however, it will be referred to hereafter as exposed area. The vertical-tail yawing-moment coefficients Cn V are referenced about the quarter chord of the vertical- tail mean aerodyn
13、amic chord and the vertical-tail root-bending-moment coefficients CB V are referenced about the vertical-tail root-chord line which is 0.154 foot above the fuselage center line. The vertical-tail derivatives CBV Cn p VP and CN % and the high-horizontal-tail deriv- ative C2 hP were obtained from angl
14、es of sideslip of -4 and 4 through- out the angle-of-attack range. I- Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-10 Static Lateral and Directional Stability NACA RM L56E2g Effects of seal.- The static lateral- and directional-stability derivativ
15、es presented,in figure 10 were determined with the vertical- tail loads balance installed with a sponge sealed gap; therefore, it was thought advisable to determine whether the model characteristics were influenced by the sponge rubber seal that had been installed at the base of the vertical tail. I
16、n view of this, tests were made to obtain the stability characteristics of a fuselage vertical-tail configuration in which the sponge seal was replaced by a solid seal. A coIIIparison of results with the sponge and solid seal (fig. 11) indicates that some leakage through the sponge seal may have occ
17、urred, since slight losses in the lateral- and directional-stability derivatives are noted especially at the higher angles of at-tack. (These differences at a = 20 represent approximately 8 and 4 percent of the measured CN and Cn %) 9 9 , respec- tively, and less than 1 percent of 8 These difference
18、s, however, are not expected to affect the validity of the comparisons which are made herein, since all the data (model loads and tail loads) were obtained simultaneously with the junctures sealed with sponge rubber. Tail contribution.- The vertical-tail contribution to the static lateral- and direc
19、tional-stability derivatives and the effect of horizontal-tail position (wing on and wing off) on this contribution are presented in figure 18 for Mach numbers of 0.80 and 0.92. In this fig- ure are included, for comparison, the vertical-tail contribution to the stability derivatives as determined f
20、rom the data of model breakdown tests (presented in fig. 10) and the contribution as determined from measured tail-loads data (presented in fig. 17). The measured tail-loads data, however, have been based on the model wing geometry and are pre- sented about the stability system of axes for the compa
21、rison shown in figure 18. The increment between the tail contribution as obtained from measured tail-loads data (solid curve) and that which was obtained from measured model-loads data (dashed curve) represents an interference or load induced by the vertical tail on the wing and the fuselage. This i
22、nduced load generally increases the increments of ACn B and AC+ at B least through a large part of the test angle-of-attack range; however, a decrease in the effective dihedral increment AC1 is noted which is B somewhat greater for the wing-on than for the wing-off configuration. At angles of attack
23、 below 15O, the measured vertical-tail normal force generally accounts for about 80 to 90 percent of the tail contribution to the lateral force ACY P of the complete model. From the standpoint of tail effectiveness, the vertical tail con- tributes a stabilizing increment to the directional stability
24、 Cn P Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L56E2g 11 of the model throughout the angle-of-attack range as shown in figure I-8; however, the stabilizing increment ACn B is greatly reduced at the higher angles of attack, as shown by
25、data obtained at a Mach number of 0.80 and for the present wing-on configuration in which the wing is mounted in a midfuselage position. This reduction in ACnS at the higher angles of attack is somewhat more pronounced with the horizontal tail in the high position than for the low-tail or horizontal
26、-tail-off configurations. In general the end-plate effect provided by the high horizontal tail produces considerable increases in ACn P throughout the range of test angles of attack, especially at the lowest Mach number. This end- plate effect, of course, is also shown in figure 10(a) where increase
27、s in Cn P for the complete model with the high horizontal tail are shown to exist at a Mach number of 0.80. At the higher Mach numbers (0.90 and 0.92) the favorable end-plate effect is lost and even becomes reversed at low angles of attack for the high-tail configurations tested. Results presented i
28、n reference 10 indicated that the loss in end-plate effect apparently resulted from a bad interference condition at the junc- ture of the horizontal and vertical tails. These adverse interference effects were reduced by moving the horizontal tail rearward so that its apex was approximately coinciden
29、t with the leading edge of the vertical tail. This, however, caused some reduction in end-plate effect at the lower Mach numbers. Vertical-Tail Loads Mach number effects.- The variation with Mach number of the vertical- acNv tail normal force per degree of sideslip angle - as is presented in figure
30、lg. acNv The values of - as are almost identical when determined at p = ;t4O from parameter tests or from sideslip tests at a = O“. The difference between the slopes measured near j3 = 0 and those obtained from parameter tests indicates the presence of some nonlinear- ities in the normal-force curve
31、s for sideslip angles between -4O and 4O. Figure 14 presents results that also indicate these nonlinearities espe- cially for the high horizontal-tail configuration. Since the following analysis is based on results obtained from the parameter tests (*b“ side- slip), it should be appreciated that the
32、 results may not truly represent the slopes at smaller sideslip angles. In general, the effect of Mach number on the vertical-tail normal force is small when the horizontal tail is located in the wing-chord plane extended; however, when the hor- izontal tail is located in the high position, apprecia
33、ble reductions Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-12 NACA RM L56E29 aCNV in - ai3 are evident at an angle of attack of O“ as Mach number increases. (See fig. lg.) Comparison of total-tail contribution and load on exposed vertical tail.-
34、The data presented in figure 20 include both the total-tail con- tribution and the load measured on the exposed vertical tail, both being based on exposed tail area. The total-tail contribution was determined from the differences between vertical-tail-on and tail-off tests which includes both the lo
35、ad carried on the exposed vertical tail and the load that the vertical tail induces on the fuselage. Differences are shown to exist between the total-tail contribution and the exposed tail load which indicate the load that is induced on the fuselage by the ver- tical tail and this load is referred t
36、o as the interference fuselage load in figure 20. For the wing-on configuration (fig. 20(a) the fuselage load is slightly greater throughout the test angle-of-attack range when the horizontal tail is placed in the low position than for the horizontal tail in the high position or off. It will be note
37、d that the induced fuselage load decreases more rapidly with angle of attack for the wing-off configuration than for the wing-on configuration and even changes sign at the higher angles of attack for the wing-off con- figuration. The delay in fuselage load reversal noted for the wing-on configuratio
38、n is probably a result of wing-wake effects. The total tail contribution and the load on the exposed vertical tail is considerably greater when the horizontal tail is placed in the high position for both the wing-on and wing-off configurations; however, reductions are evidenced with increased angle
39、of attack with all horizontal-tail positions for the complete model configuration. The low a%T values of - ap at the higher angles of attack for the complete model configuration (fig. 20(a) do not necessarily indicate low overall tail aC% loads, as pointed out in reference 2, because this low value
40、of - 93 is also indicative of low static-directional stability. Therefore, under the conditions of low static-directional stability (fig. 10) large angles of sideslip might be expected and, consequently, the tail loads at high angles of attack may be more critical than at the lower angles of attack.
41、 Effect of horizontal-tail position on exposed vertical-tail load.- In order to illustrate further some effects of horizontal-tail position on the exposed-vertical-tail load, some of the data from previous fig- ures have been presented in figure 21 for a more direct comparison. The horizontal tail i
42、n the wing-chord plane extended had little effect Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM 5629 13 on the exposed vertical-tail load throughout the test range of Mach number. At a Mach number of 0.80, the horizontal tail, when placed i
43、n the high position, produces an end-plate effect, or an increase in effec- tive aspect ratio of the vertical tail, and thereby increases the vertical- tail normal force through the angle-of-attack range. At a Mach number of 0.92, however, and in the vicinity of an angle of attack of O“, a decrease
44、in vertical-tail normal force exists, which is probably due to a bad interference condftion at the juncture of the horizontal and ver- tical tails as discussed previously. Results of reference 10 at a Mach number of 0.90 have indicated that significant increases in the direc- tional stability C-np,
45、at or near an angle of attack of O“, can be obtained simply by moving the high horizontal tail longitudinally with respect to the vertical tail. Therefore, these increases in cnP would be expected to be associated with increases in the exposed vertical-tail load and center of pressure when the horiz
46、ontal tail is in the high position. The results presented in figure 22 are used to illustrate further how the vertical-tail effective aspect ratio is influenced by Mach number and horizontal-tail position. The values of the ratio of effective aspect ratio of the vertical tail to geometric aspect rat
47、io A, were derived i by using the experimental a%J ai3 1 0 Av - from theoretical expressions presented in reference 11. From these results presented in figure 22 (a = O“), the end-plate effect provided by the horizontal tails is evident. The effective aspect ratio of the vertical tail for the low-ta
48、il and tail-off configurations is about constant throughout the range of Mach number; whereas, the high tail decreases the effective aspect ratio as Mach num- ber increases. Vertical-tail effective center of pressure.- The vertical-tail spanwise center of pressure will be referred to as an effective
49、 center of pressure since it was derived by dividing the vertical-tail root bending moment per degree of sideslip by the vertical-tail normal force per degree of sideslip. For the comparisons presented in figure 23, the spanwise location of the effective center of pressure I$ also includes the rolling moment and side force that the high horizontal tail imposes on the vertical tail. The effective
