Document
LOW-SPEED W I N D - T U N N E L STUDIES
RELATING TO PITCH-UP O N A
SUPERSONIC TRANSPORT MODEL
W I T H A HIGH-ASPECT-RATIO
VARIABLE-SWEEP WING
by Vernurd E. Lockwood
LczngZey Reseurch Center
LungZey Station, Humpton, Vu.
. . ,l N A T I O N A L A E R O N A U T I C S A N D S P A C E A D M I N I S T R A T I O N W A S H I N G T O N , D. C . A U G U S T 1 9 6 6 TECH LIBRARY KAFB, NM
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013028L NASA ‘I” U-3342 LOW-SPEED WIND-TUNNEL STUDIES RELATING TO PITCH-UP ON A SUPERSONIC TRANSPORT MODEL WITH A HIGH-ASPECT-UTI0 VARIABLE-SWEEP WING By Vernard E. Lockwood Langley Research Center Va.
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LOW-SPEED WIND-TUNNEL STUDIES RELATING TO PITCH-UP ON A SUPERSONIC TRANSPORT MODEL WITH A HIGH-ASPE CT -RATIO VARIABLE -SWEEP WING By Vernard E. Lockwood Langley Research Center SUMMARY A low-speed investigation has been made t o determine methods of reducing the pitch- up tendency of a high-aspect-ratio variable-sweep wing model. The model represented a three-engine arrangement of a supersonic transport configuration where two of the nacelles were mounted below the wing on pylons supported from the fuselage and a third nacelle was mounted in the vertical tail just above the fuselage. The model had a variable- sweep cambered and twisted wing and an outboard pivot location. Among the variables studied were movable wing sweep and incidence; sweep, size, and deflection of the fore- Also included wing; canard size and deflection; and horizontal-tail size and deflection.
are the effects of fuselage nose cross-sectional shape on the lateral directional stability characteristics.
The investigation showed the basic model to have a pitch-up tendency which increased as the movable wing leading-edge sweep was increased from 0 ' t o 2 5 O . At the wing-sweep angle of 160, for which most of the investigation w a s made, the pitch-up started at an angle of attack as low as 80. The results of the study showed the pitch-up tendency of the model could be reduced materially either by deflection of a forewing flap o r by reduction in sweep and size of the forewing. Addition of deflected canards also reduced the pitch-up tendencies. Improvement in the lift-drag ratios was noted for the model with deflected forewings o r forewings with reduced sweep and for the canards with small deflections.
Positive incidence changes in the movable wing also increased lift-drag ratios. The modifications which improved the stability and lift-drag ratios generally resulted in improved horizontal-tail effectiveness.
A change from the basic c r o s s section of the fuselage nose to a circular section reduced the directional stability at an angle of attack above 1 0 ' by reducing the favorable side-force variation with angle of attack.
The investigation was made in the Langley high-speed 7- by 10-foot tunnel at a Mach
number of 0.20, which corresponds to a Reynolds number of 10.7 X lo6 based on the fuse-
lage length.
INTRODUCTION Recent wind-tunnel investigations of variable-sweep supersonic transport configu- rations have indicated pitch-up problems in the intermediate to high angle-of -attack range.
One investigation (ref. 1) which employed high-aspect-ratio wing panels with outboard pivots indicated that the pitch-up began at an angle of attack as low as 5" for the 16O movable-wing-sweep position. It was thought that the pitch-up was caused by a combina- tion of nonlinear pitching moments of the forewing-fuselage combination reinforced by flow separation on the outer movable wing. The possibility of reducing this pitch-up by raising the horizontal tail from its position in the wing-chord plane was investigated and the results (ref. 1) indicated that the tail gave g r e a t e r stability at low angles of attack, but contributed destabilizing moments at the high angles of attack because it became immersed in the wing wake. Lower horizontal-tail positions were considered, but were of the tail to the engine exhaust.
assumed to be impractical because of the proximity Since it w a s believed that the forewing-fuselage combination w a s responsible, to a large degree, for the pitch-up of the configuration, the possibility of variations in forewing geometry was considered as a means of reducing the pitch-up problem.
The present paper is an extension of the investigation of reference 1 t o include studies of variables associated with the forewing geometry. The variables included changes in sweep, size, shape, and deflection of a forewing flap. Canards also were studied as a means of reducing the pitch-up of the basic model; in this respect, it was thought that the stalling of the canard might tend to compensate for the nonlinear lift on the forewing.
Most of the investigation was made with the movable-wing leading-edge sweep angle at 16O and the movable-wing incidence at 00. A limited amount of data, however, w a s also obtained at other sweep angles and incidence angles. The basic horizontal tail of the model (ref. 1) which had 20° anhedral w a s used for most of the study, but data also were obtained with the anhedral eliminated and with a flat-plate tail having an a r e a about 51 p e r - cent greater than the basic configuration. The investigation included experiments to determine the effect on directional stability of changing the basic fuselage nose cross- sectional shape to one having a circular shape.
The investigation was made in the Langley high-speed 7- by 10-foot tunnel at a Mach number of 0.2 and a Reynolds number of 10.7 X lo6 based on the fuselage length.
SYMBOLS The f o r c e and moment data contained herein are referred t o the axis system shown in figure 1. The reference dimensions used in reducing the data are given in figure 2 and table I and a r e the same as those used in reference 1.
Measurements for this investigation were taken in the U.S. Customary System of Units. Equivalent values a r e indicated herein parenthetically in the International System (SI) in the interest of promoting use of this system in future NASA reports. Details con- cerning the u s e of SI, together with physical constants and conversion factors, are given in reference 2.
The moment reference for most of the configurations investigated is at the wing pivot station o r 5.715 reference chords behind fuselage station 0. For model configura- tions with the canards installed or with forewing-sweep angles other than 76O, a moment reference was chosen which would give the same degree of stability as that of the basic model with a movable-wing-sweep angle of 1 6 O .
b reference wing span, 85.84 inches (2.180 meters) bC span of canard, inches (meters) drag coefficient, CD Lift
lift coefficient, -
CL q s increment in lift coefficient ACL C lift-curve slope at CL = 0 L a rolling-moment coefficient, Rolling moment cZ qSb Pitching moment pitching-moment coefficient, ---- Cm qsc
increment in pitching - mo ment coefficient
ACm II I 111 I 1 1 1 1 1 1111111111111111 111111111111111111111.1111111111.1111 I 1 1 1 1 II 111 II I . I 11.11l.1 .11111111 111 1 1 1 - . . . - . - 1 1 . 1 1 . - . 1 1 - - - 1 - - 1 1 - a C m / a C ~ pitching-moment-curve slope at CL = 0 Yawing moment yawing - moment coefficient , Cn qSb Side force side-force coefficient, CY q s
c = (CY, p=50 - CY, p =- 54
yP
reference wing chord, 9.93 inches (0.2522 meter) root chord of canard, inches (meters) tip chord of canard, inches (meters) lift-drag ratio maximum lift-drag ratio dynamic pressure, pounds/feet2 (newtons/meter2) reference wing area, 3.94 feet2 (0.3660 metera) canard area, feet2 (meter2) moment reference location from nose of model, reference wing chords angle of attack of fuselage reference line, degrees angle of sideslip, degrees increment in sideslip between p = *5O, corrected for balance and strut deflection, degrees horizontal-tail deflection (positive when trailing edge is down), degrees forewing flap deflection (positive when leading edge is down), degrees 6f movable-wing incidence (positive when leading edge is up), degrees
iw
wing leading-edge sweep angle, degrees A canard incidence (positive when leading edge is up), degrees 6, Configuration de signations: B fuselage specific fuselage (see fig. 5) B1 ,B2 C canard specific canard (see fig. 7(b)) C1,C2,C3
F1 . . . F7 forewing o r forewing flaps (see fig. 6)
H horizontal tail specific horizontal tail (see fig. 7(a)) H ~ , H ~ , H Q N fuselage-mounted engine nacelles vertical tail with engine nacelle VN W movable wing MODEL The basic model investigated features a high-aspect-ratio variable- sweep wing with an outboard pivot and a three-engine nacelle arrangement as shown in figure 2. A photograph of the model in the Langley high-speed 7- by 10-foot tunnel with the movable wing at 16O leading-edge sweep is shown as figure 3. This model is essentially the s a m e model as that investigated and reported in reference 1 except that the airfoil sections are different. The airfoil sections of the present wing are presented t o scale in figure 4.
Other dimensions of the basic model are given in table I.
I II I I Ill 111ll111l l I l I l l l 1 1 1 l 1 1 1 1 1 1 1 1 1 1 1 1 l l l l I I l I I I I
During the investigation the basic fuselage nose cross-sectional shape B1 w a s altered to a circular section B2 as shown in figure 5. The cross-sectional-area distri- bution was identical for the B1 and B2 nose sections. The B2 nose blended into the main body at fuselage station 29.59 in. (0.7516 m).
Several modifications were made to the forewing area between the fuselage and mov- able wing and also between fuselage stations 29.50 in. (0.7493 m) and 54.13 in. (1.3749 m).
The modifications r e f e r r e d to as forewing flaps are portions of the forewing capable of being deflected as shown in figure 6. Forewing flaps F1, Fa, and F3 had 76O leading- edge sweep and hinge lines perpendicular to the plane of symmetry as shown in figure 6(a).
Forewing flaps F4, F5, and F6 had leading-edge sweeps of 76O, 73O, and 720, respec- tively, as shown in figure 6(b). Forewing flap F4 had a hinge-line sweep of 70° whereas forewing flaps F5 and F6 had hinge-line sweeps of 66O. Forewing flap F7 had a leading-edge sweep of 70° and was not deflected.
Three horizontal-tail configurations were tested on the model. Tail H1 corre- sponds to the basic horizontal tail tested in reference 1 which had 200 of anhedral as shown in figure ?'(a). Configuration Ha w a s the same surface as H1 but with the dihedral angle removed. A l a r g e r horizontal tail H3 made of 0.25 in. (0.0064 m) aluminum plate and having no dihedral was utilized for some tests. Other details concerning the horizontal tails are given in figure 7(a) and tables 1 1 and III.
Canards C1, C2, and C3 used in the investigation a r e illustrated in figure 7(b).
They were made of 0.125-in.-thick b r a s s with rounded leading edges and blunt trailing edges. They had an aspect ratio of 3.0 and were mounted with 50 percent of their root chord at fuselage station 10.40 in. (0.2642 m). Other details of the canards a r e given in figure 7(b) and table IV.
TEST AND CONDITIONS 7- by 10-foot tunnel. The The investigation was made in the Langley high-speed model was sting mounted, as shown in figure 3. Forces and moments were measured by an internally mounted six-component strain-gage balance. To insure a turbulent boundary layer, transition s t r i p s of No. 100 carborundum grit approximately 1/8 inch wide were affixed to the model at 7 percent of the length of the body and at 7 percent of the chord of the wing and all appendages.
The tests were made at a dynamic p r e s s u r e of 56.9 lb/ft2 (2723 N/m2), a Mach num-
ber of 0.2, and an average Reynolds number of 10.7 X lo6 based on the fuselage length. In
addition to the normal angle-of-attack runs at Oo sideslip, some selected configurations also were tested at rt50 sideslip.
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ~ -- ~ ~~ _._..
The drag data were corrected t o correspond t o a pressure at the base of the engine nacelles and the base of the fuselage equal to free-stream static pressure. The nacelles had constant-diameter internal passages, for which an axial-force coefficient of 0.00254 f o r all three nacelles was calculated as being the internal skin friction due t o flow through the nacelles. This value was subtracted from the drag coefficient obtained experimentally.
The jet-boundary corrections calculated f o r the drag and angle of attack by the method of reference 3 for the reference sweep condition are as follows: The jet-boundary corrections t o the pitching-moment data were found t o be negligible.
The data were also corrected f o r wind-tunnel blockage by the method presented in refer- ence 4. The angles of attack and sideslip were corrected for deflection of the sting- support system under load.
PRESENTATION OF DATA The results of the investigation are presented in the following figures: Figure Longitudinal stability character istics :
Effect of movable-wing sweep and incidence . . . . . . . . . . . . . . . . . . 8 t o 11
12 to 13 Effect of forewing flap sweep. . . . . . . . . . . . . . . . . . . . . . . . . . .
14 to 16 Effect of forewing flap deflection. . . . . . . . . . . . . . . . . . . . . . . . .
Effect of canard size and deflection . . . . . . . . . . . . . . . . . . . . . . . 17 t o 19
Effect of fuselage-nose c r o s s section . . . . . . . . . . . . . . . . . . . . . . 20
Horizontal-tail effects study . . . . . . . . . . . . . . . . . . . . . . . . . . . 21 t o 22 Longitudinal control characteristics: 23 t o 25 Effect of horizontal tail and forewing flap . . . . . . . . . . . . . . . . . . . .
Lateral stability characteristics:
Effect of fuselage-nose c r o s s section . . . . . . . . . . . . . . . . . . . . . . 26
IIIIIIIII l l l l l
DISCUSSION Longitudinal Stability
Movable-wing sweep. - The movable-wing-sweep effects on the aerodynamic charac-
BIWHIVNNF1, sf = 00, and iw = Oo, are shown in
teristics of the basic configuration figure 8. It will be noted that pitch-up tendencies exist in the data for all sweep angles and become more pronounced as the wing-sweep angle is increased from Oo t o 25O. For the wing-sweep angle of 160, for which most of the tests were made, the pitch-up started as low as an angle of attack of 80. The trend of Cm with a ! o r CL is similar to that experienced with the model of reference 1 which had the same fuselage and tail configura- tion but, as previously mentioned, different wing airfoil sections. In the earlier investi- gation as in the present one, the pitch-up was attributed t o airflow separation over the wing combined with increasing lift on the forewing. Wings experienced early separation as indicated by tuft studies beginning at 5 O in the case of reference 1 and at 8O in the present study.
A plot of the low-lift aerodynamic parameters against sweep angle (fig. 9) shows that the aerodynamic center varies linearly with sweep over the range of sweep angles tested. The overall variation amounts to about 48 percent of the reference chord for movable-wing-sweep angles from 00 t o 25O.
Movable-wing incidence.- The effects of positive changes in movable-wing incidence are shown in figure 10. The data at high angles of attack show the pitch-up tendency was essentially unchanged for all incidence angles tested. At low angles of attack, the increased incidence gave negative t r i m changes of sizable magnitude and a small forward shift in the aerodynamic center as indicated in figure 11. The change in t r i m is equivalent t o about 4O of up elevator (fig. 25). The most noteworthy effect of incidence change, how- ever, was the increase in (L/D),, which amounted t o about 2.3 units. The increase in (LID),, probably results from a more favorable matching of the loading on the outboard panel with that on the highly swept forewing.
Forewing sweep.- One of the modifications tested which improved the low-speed high-angle-of -attack characteristics but which would compromise the high Mach number cruise configuration is lower sweep on the forewing, as indicated in figure 12. Reduction of the sweep angle from 760 with F1 t o 70° with F 7 reduced the pitch-up tendency con- siderably; a sizable reduction was also indicated for the configuration F5 with 730 sweep. The improved characteristics probably result from a more linear lift variation with angle of attack over the forewing, decrease of forewing area, smaller induced effects on the adjoining movable wing, and improved flow conditions at the horizontal tail. The improved flow conditions at the tail are in evidence in figure 13 where the horizontal-tail contribution to stability AC, is shown to be greater at an angle of attack above 12O for the forewings having 70° and 730 of sweep than for the forewing having 760 sweep.
In addition t o the improvements in pitch characteristics, there was also an increase in ( L / D ) m z . Reduction of the forewing sweep from 76O to 70° resulted in an increase of 1.8 in ( L / D ) m a as shown in figure 12.
Forewing flap -_ deflection.- The deflection of the forewing flaps with 76O sweep and . . . .
axes perpendicular to the plane of symmetry F1, F2, and F3 gave considerable reduc- tions in pitch-up tendencies. The reductions in pitch-up which increased as the a r e a of the forewing flap increased appeared to result mainly from losses in lift on the fixed apex as indicated in figure 14. Some of the reduction in pitch-up may also be due to improved flow conditions for the outboard movable panel as well as the horizontal tail. Evidence of the improved flow conditions at the tail is shown in figure 15 in the plot of AC, against a. ACm at high angles of attack is greater with the The tail contribution t o stability forewing flap deflected loo than with the flap undeflected.
The pitching-moment improvements shown by the deflected flaps were not without compromise to the other characteristics. There were losses in (L/D),= f o r the large deflections of the flaps which were required for material improvements in the high-angle- of -attack stability. Additional losses in (L/D),, would occur in trimming out the neg- ative pitching moment resulting from flap deflection.
Figure 16 shows the effect of deflecting forewing flaps F4 and F6 about hinge lines which have sweep angles of 70° and 660, respectively. These flap configurations as in the case of F1 and F2 gave considerable reduction in the pitch-up tendency but unlike F1 and Fa gave only minor changes in the variation of C, or CL with a!
in the low-lift range.
Neither flap eliminated the pitch-up, but with 300 deflection the variation of C, with CL o r a! w a s closer to being linear with F6 than with Fq (which had the l a r g e r leading-edge sweep). Both flaps show increases in (L/D),, over the basic configura- tion (F4 undeflected) for a wide angle-of -attack range. This characteristic indicates the possibility of improving the low-speed lift-drag ratios of airplanes designed princi- pally from supersonic considerations.
Canards.- - A comparison of the aerodynamic characteristics of the basic model with those of the canard (undeflected) configuration is given in figure 17. These data indicate a destabilizing action on the model at high angles of attack with the horizontal tail off, but with the horizontal tail on, the pitch characteristics a r e similar to those of the basic model.
The stabilizing effect of the canards on the horizontal-tail characteristics is shown in fig- u r e 18 and is similar to that achieved with deflected forewing flaps or reduced sweep on the forewing. When the canards a r e deflected positively (fig. 19), a stabilizing tendency is
I l l 1 1 l 1 1 1 l l l I I l l I
noted as the result of the canard stall. For example, a loo deflection of the largest
canard C3 resulted in a loss of lift and an increase in CL for pitch-up from 0.9 t o 1.3.
(See fig. 19(c).) A corresponding increase of a ! also was noted. The stalling of the canards, however, produces a slight loss of lift and an increase of drag that materially reduce the L/D of the configuration.
In contrast to the losses of lift from the deflected canards, the presence of the unde- flected canards materially increased the L/D of the basic configuration. (See fig. 17.)
In the case of These C2, the untrimmed increment in (L/D),, was about 3.0.
L/D along with those obtained from other modifications indicate that increases in methods are available for increasing the low-speed L/D of configurations designed principally for supersonic cruise.
Fuselage.- Because of the reduced area in the pitch plane, it would be expected that the fuselage with the circular sections would give l e s s positive pitching moments B2 than the basic fuselage B1. This configuration, which was tested with forewing flap F6 deflected 22.5O, is shown in figure 20 and it will be noted that the circular fuselage pro- vided a slight reduction in pitch-up.
Horizontal tail.- A comparison of the aerodynamic characteristics of three horizontal-tail configurations is given in figure 21 and of the three configurations the basic tail with -20' dihedral HI is the most effective in reducing pitch-up. The effectiveness of H2 is reduced at high angles of attack. Although H3 has a greater degree of initial stability, it tends to diminish in effectiveness as the angle of attack increases above 8O.
(See fig. 22.) The -2OO dihedral of HI probably places it in a more favorable field of downwash than H2 or H3.
Longitudinal Control The longitudinal control parameter 8cm/a6h determined from the data of fig- u r e s 23 and 24 for 50 deflection of the horizontal tail is presented in figure 25. The horizontal-tail control is positive everywhere negative values of 8Cm/ 86 within the
( h)
angle-of -attack range investigated. The values of 8Cm /ash became more negative (this condition is indicative of an increase of control) at an angle of attack above 3 O because of the emergence of this horizontal control from the wing wake. The increased effectiveness of H3 over H1 is approximately proportional t o the increase in horizontal-tail area.
Lateral Directional Stability The data of figure 26 show the effect on the lateral directional stability parameters of changing from the basic nose B1 t o the circular nose B2. This change in cross- sectional shape resulted in Cn decreasing with a ! > loo. The unusual P
and cya
variations of C and C y that is, the positive increase of the parameters in the upper "P P' B1. It w a s suggested that a force angle-of-attack range, were noted in reference 1 for opposed to the normal cross-flow components was being developed on the nose of the fuse- lage as a result of the particular cross-sectional shape. Previous experience with two- dimensional noncircular c r o s s sections (ref. 5) had shown that cross-wind forces opposite t o those generally expected can develop on cylinders at certain Reynolds numbers and section orientation. It w a s presumed that this property was responsible for the increase in C and C y exhibited in figure 26 for configurations with vertical tail off and on.
"P P This supposition was confirmed when it was shown (fig. 26) that the values of and Cn P were considerably less positive at high angles of attack for the circular nose model c y P than for the basic nose model. In fact, the circular nose model with vertical tail became directionally unstable above Q! = 210.
Although these data suggest the possibility of improving the directional stability a long nose by proper selection of characteristics of an airplane configuration employing a cross-sectional shape, care must be used in applying these data because of Reynolds number effects. Very large differences exist in the cross-section Reynolds number range represented by these tests and those that would exist on a full-scale airplane in the landing o r take-off attitude. In addition to Reynolds number and cross-sectional forebody shape nose pointedness appears t o be an important factor in the development of favorable side forces, as indicated in reference 6. These data, which are for some isolated bodies having varying cross-sectional ellipticity and which may be employed as fuselage forebodies, show that the side force can be dependent on forebody fineness ratio, SUMMARY OF RESULTS An investigation has been made in the Langley high-speed 7- by 10-foot tunnel to determine methods of reducing the pitch-up tendency of a variable-sweep wing configura- tion employing a highly swept forewing and an outboard pivot. The results of the low- speed investigation are summarized as follows: 1. The basic model exhibited an increasing pitch-up tendency as the movable-wing leading-edge sweep angle was increased from 0' t o 25O.
For the wing position of 16O, for which most of the investigation was made, the pitch-up tendency started as low as an angle of attack of 8 O .
2. The pitch-up tendency of the model could be reduced materially by u s e of deflected forewing flaps, by substitution of a lower sweep angle of the leading edge of the forewing or addition of deflected canards.
3. Improvements in the maximum lift-drag ratio over the basic model were obtained with several deflected forewing flaps, reduced sweep angle of the forewing, addition of .. ,_.. . . ,, ._ . .. .. .
canards with little or no deflection, or positive change in incidence of the movable wing.
4. The modifications which improved the pitch-up characteristics and lift-drag ratios generally resulted in improved horizontal-tail effectiveness.
fuselage nose from the basic shape to 5. Changing the cross-sectional shape of the circular reduced the directional stability at angles of attack above 100 by reducing the favorable side-force variation with angle of attack.
Langley Research Center, National Aeronautics and Space Administration, Langley Station, Hampton, Va., March 28, 1966.
REFERENCES 1. Lockwood, Vernard E.; McKinney, Linwood W.; and Lamar, John E.: Low-Speed Aero- dynamic Characteristics of a Supersonic Transport Model With a High-Aspect-Ratio Variable-Sweep Warped Wing. NASA TM X-979, 1964.
2. Mechtly, E. A.: The International System of Units - Physical Constants and Conversion
Factors. NASA SP-7012, 1964.
3. Gillis, Clarence L.; Polhamus, Edward C.; and Gray, Joseph L., Jr.: Charts for Deter- mining Jet-Boundary Corrections for Complete Models in 7- by 10-Foot Closed Rectangular Wind Tunnels. NACA WR L-123, 1945. (Formerly NACA ARR L5G31.)
Blockage Corrections for Three-Dimensional-Flow Closed-Throat 4 , Herriot, John G.: Wind Tunnels, With Consideration of the Effect of Compressibility. NACA Rept. 995, 1950. (Supersedes NACA RM A7B28.)
5. Polhamus, Edward C.; Geller, Edward W.; and Grunwald, Kalman J.: P r e s s u r e and as Affected by Reynolds Number Force Characteristics of Noncircular Cylinders With a Method Included for Determining the Potential Flow About Arbitrary Shapes.
NASA TR R-46, 1959.
6. Spencer, Bernard, Jr.; and Phillips, W. Pelham: Transonic Aerodynamic Character- istics Of a Series of Bodies Having Variations in Fineness Ratio and Cross-Sectional Ellipticity. NASA T N D-2622, 1965.
TABLE 1.- BASIC MODEL DIMENSIONS Fuselage:
Length. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 91.00 in. (2.3114 m)
Base area . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.080 sq ft (0.0074 sq m)
Chamber area. . . . . . . . . . . . . . . . . . . . . . . . . 0.034 sq ft (0.0032 sq m)
Wing (reference):
Leading-edge sweep . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 6 . 0 deg
Root chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.93 in. (0.2522 m)
Tip chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.29 in. (0.0836 m)
Span. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 85.84 in. (2.1803 m)
Area. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 . 9 4 sq f t (0.3660 sq m)
Aspect ratio. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.99
Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.33
Wing (actual):
Leading-edge sweep . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16.0 deg
Root chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10.03 in. (0.2548 m)
Tip chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.26 in. (0.0828 m)
Span. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 85.40 in. (2.1692 m)
A r e a . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.94 sq f t (0.3660 sq m) Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.85
Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.325
Horizontal tail HI:
Leading-edge sweep angle . . . . . . . . . . . . . . . . . . . . . . . . . . . 65.4 deg
Root chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.40 in. (0.3658 m)
Tip chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.86 in. (0.0980 m)
Span, total. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18.96 in. (0.4816 m)
Exposed area . . . . . . . . . . . . . . . . . . . . . . . . . 0.889 sq ft (0.0826 sq m)
Dihedral angle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -20.0 deg
Vertical tail:
Leading-edge sweep angle . . . . . . . . . . . . . . . . . . . . . . . . . . . 37.3 deg
Root chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.00 in. (0.3556 m)
Tip chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.80 in. (0.0711 m)
Span, exposed . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10.38 in. (0.2637 m)
Area, exposed. . . . . . . . . . . . . . . . . . . . . . . . . 0.606 sq f t (0.0563 sq m)
Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 . 2 4
I I llIlIIlllIll I I I I I I1 I I
TABLE I I . - HORIZONTAL-TAIL PROJECTED DIMENSIONS Horizontal tail HI:
Root chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.40 in. (0.3658 m)
Tip chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.86 in. (0.0980 m)
Semispan, exposed . . . . . . . . . . . . . . . . . . . . . . . . . 7.26 in. (0.1844 m)
Area, exposed. . . . . . . . . . . . . . . . . . . . . . . . . 0.889 sq f t (0.0826 sq m)
Dihedral angle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -20.0 deg
Horizontal tail H2:
Root chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.40 in. (0.3658 m)
Tip chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.86 in. (0.0980 m)
Semispan, exposed . . . . . . . . . . . . . . . . . . . . . . . . . 7.72 in. (0.1961 m)
Area, exposed. . . . . . . . . . . . . . . . . . . . . . . . . 0.889 sq f t (0.0826 sq m)
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.0 deg
Dihedral angle Horizontal tail H3:
Root chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17.34 in. (0.4404 m)
Tip chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.86 in. (0.0980 m)
Semispan, exposed . . . . . . . . . . . . . . . . . . . . . . . . . 9.86 in. (0.2504 m)
Area, exposed. . . . . . . . . . . . . . . . . . . . . . . . . 1.343 sq f t (0.1248 sq m)
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.0 deg
Dihedral angle TABLEIT.1.- H1 AND H2 AIRFOILORDINATES . .- Distance from Surface ordinate, leading edge, percent chord per cent chord __ 0.0
0 .o
10 .o .542
20 .o .959
30 . O 1.265 40 .O 1.445 50 .O 1.500 60 .O 1.445 70 .O 1.265 80 . O .959
90 .o .542
100.0 .o
*
Distance frorr :eference chord.
TABLE 1V.- CANARD DIMENSIONS Canard C1:
Root chord, cr . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.62 in. (0.1173 m)
Tip chord, c t . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.71 in. (0.0180 m)
Span, b , . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.96 in. (0.2022 m)
Area, Sc . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.147 sq ft (0.0137 sq m)
Canard C2:
Root chord, cr . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.08 in. (0.1798 m)
Tip chord, c t . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.08 in. (0.0274 m)
Span, bc . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.24 in. (0.3109 m)
Area, Sc . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.346 sq f t (0.0322 sq m)
Canard C3:
Root chord, cr . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.88 in. (0.2255 m)
Tip chord, ct. . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.36 in. (0.0345 m)
Span, bc . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15.36 in. (0.3901 m)
Area, Sc . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.546 sq f t (0.0507 sq m)
I
2 Side force
Relative
--
__f
D rag
_--
\ \ \ \ ‘L-J
D rag
wind
Figure 1.- System of axes used in presentation of data.
Reference dimensions 42.70 fl.0846) Chord 9.93 in. (0.2522 m l Span 8 5 8 4 in. (.?.I803 in) 3 . 9 4 s9 f t (0.3660 sqm/ A r e a Sweep f6.0" t *2& (0.0828) (23114) -.
Figure 2.- Three-view drawing of model. A l l linear dimensions are in inches (meters).
Figure 2.- Concluded.
L-62-8932 Figure 3.- Photograph of model i n Langley high-speed 7- by 10-foot t u n n e l .
Semispan station ____ Reference l i n e Reference l i n e - ~- Reference ‘line Reference l i n e Reference l i n e ~ - Reference l i n e - ..- .
--- ~.
17.168 (0.4861) . * . ‘ 1
Reference l i n e G I - . -----------__ . .. .
--\_ 13.642 (0.3465) Reference l i n e - -~ 1 _.
10.881 (0.2764) Reference l i n e - - -- - - - - 6.925 (0.1759)
-
Reference l i n e .
- -_ - Figure 4.- A i r f o i l sections of wing. Dimensions are given in inches and parenthetically in meters.
1 ' 1
I I I 1 2.96 592 11.83 14.79 1275 1825 2071 2203 2.367 258026.63 29.59 888 (0.3005) (0.3757) /0.4509)10.4636) (0.5260) (0.55961 (0.6012) 10.6553)@.6Z4) (0.7516) (0.0752) (0.1504) (0.2256) Fuse luge sf af ions I I - _ _ - - B a s i c f u s e l age,B, I I I I 1 : C i r c u l o r f u s e I age, B2 I
I
I Figure 5.- Comparison of fuselage-nose shapes B1 and B2. Dimensions of stations are given in inches and parenthetically in meters.
IIIII Ill 1 1 1 1 1 1 1 l 1 l 1 I l I l l 1 11l11lllIl I 1 I I 1 I
I F3 ---- - - (a) F1, Fp. and F3.
Figure 6.- Geometry of forewing flaps. A l l linear dimensions a r e given in inches a n d parenthetically in meters.
I Fusefoge sfofions 54f3 (137491 5675 (1.44141 I fO.Ol69J .66~ . F&eloge reference fine I I ' I Deffecfions 8 f measured i n o pfone perpendicufar fo hinge line Secfion A - A , f y p i c o f --
I--'
(b) Fq. F5, Fg' a n d F7.
Figure 6.- Concluded.
c 2 / 5 4 (0.5471) (0.096GI
c ~-
u l6.84 (0.4271) - . .
, -~
t
9.86 ' (0.25001 __ ~ I .- ._.__ 8.90 rai , I Z72 (0.19611
I
.25 inch p l a t e w i t h rounded leading edge and blunt trailing Fuselage l i n e 2 1 4 2 + . - .
i i1
(0.0loh , . l4.40 10.36581
-- 1234 (0.4404) (a) Horizontal tail, Figure 7.- Geometry of horizontal tails. A l l linear dimensions are given in inches and parenthetically in meters.
. -
-
- f0.40
I Bfunt t r a i l i n g edge
$
Rounded leading edge
I I
(b) Canards.
Figure 7.- Concluded.
.8 .5 4 c , .3 / - v -.6 -4 -2 0 .2 4 .6 B L O L 2 L 4 L 6 L 8 CL Figure 8 . - Effect o f wing-sweep angle o n aerodynamic characteristics. BlWHiVNNF1. A = 16O; bh = Oo; x/c = 5.715; bf = Oo; iw = Oo.
Figure 8 . - Concluded.
I
0 5 /O 20 25 30
Figure 9.- Variation of aerodynamic parameters C bCm/dC,, a n d (L/Dlmax w i t h wing-sweep angle. B I W H I V N N F 1 ; 6 h = 0 ' ; 6f = Oo; iw = Oo.
h' iw , d e !
0 0 0 1 . 9 0 9.2 5 . 4 i
-4 1
I I1 - 86 .2 . u 11 Figure 10.- Effect of movable-wing incidence o n aerodynamic characteristics. BiWHiVNNFI; A = 16O; 6h = 00; x/c = 5.715; 4 = 00.
w Figure 10.- Concluded.
/ 2 3 4 5 Figure 11.- Variation of t h e aerodynamic parameter C d C , / d C , , a n d ( L / D + , , a x w i t h wing-incidence angle.
b' BIWHIVNNF1; A = 16'; 6 h = 00; 4 = W.
CD I r S wee,c -2 0 .2 . 4 .6 L O 6 2 L 4 L 6 6 8 20 22 CL (a) H I .
Figure 12.- Effect of forewing-flap sweep angle on the aerodynamic characteristics. BIWHVNN; I\ = 16'; dh = 0 ' ; 15, = Oo; iw = Do.
0 .2 .4 . 6 . 8 10 12 /.4 16 - 4 O 4 8 I2 1 6 20 24 (a) Concluded.
Figure 12.- Continued.
W W 1.11 I I I I. I, I I I , . I. I . . . . . -..-.--.- ----. .. ~ . I I I I 1111.1.111111 I . . I I 111 I II I I 111 II I I I .6 .5 .4 CD / veep ang de9 O F ; 76 7 70 . 4 .6 .8 L O L 2 L4 L 6 /.8 -2 0 CL (b) Horizontal t a i l off.
Figure 12.- Continued.
I - . 6
-. 8
4 . 0 -4 0 4 8 I2 I6 20 24 -.2 0 .2 4 .6 .8 10 l2 L 4 1.6 (b) Concluded.
Figure 12.- Concluded.
..2
./
-./
./
- ./
-2
-4
- 4 0 4 8 /2 /6 20 24
Figure 13.- Effect of forewing-flap sweep on the increment i n pitching moment and l i f t coefficients due to the addition of the horizontal tail.
BIWHIVNN; A = 1 6 ' ; 6h = 0 ' ; 4 = 0 ' ; iw = 0 ' .
.8 .7 .6 .5 .3 .2 . / 4 .6 .8 LO M L 6 818 CL (a1 F1.
F i g u r e 14.- Effect of forewing-flap deflection o n t h e aerodynamic characteristics. B I W H I V N N ; A = 16O; 6 h = Oo; x/c = 5.715; iw = 00.
w Q, (a) Concluded.
Figure 14.- Continued.
I I I I
I I
! !
I ' I
I
I I I
I
ii
I
I
I
I i
I
I
I
f
l
a
I
i
i
I
b - 1 I
I
I
I I I I
I I
I
I
- ?1 W b 6 LZ M .8 LO . 4 . 6 0 .2 CL (b) F2 Figure 14.- Continued.
A (b) Concluded.
Figure 14.- Continued.
I I I
I
I
I , I I I I I
I i
I I . 7
I
I I I .6 I
i
I
.5 I I
I
I
.4 CD
t
.3 I .2 . / &
' f J d e g
I I
IO
f I
' I
I !
i6 I2 a,deg -4 L 6 /.8 20 22 .8 6 0 L 2 64 . 6 CL Figure 1 4 . - Continued.
. 2 Cm Q 0 __ -4
0 lo -
0 20 __ -.6 (c) Concluded.
Figure 14.- Concluded.
- . 7
- 4 0 4 8 1 2 16 20 24
a,@7
Figure 15.- Effect of forewing-flap deflection o n t h e increment in pitching-moment and lift coefficients due to the addition of the horizontal tail.
BIWHVNNF1: A = 16'; i , = Oo.
I II I I I II I I I 1 1 1 1 1 1 1 1 1 1 1 . 1 1 1 I l l 11111111111.11111111111111111 111 I I I I I I I I I I I I . 1 1 1 1 1 1 1 . 1 1 n 1 1 1 9 . 1 1 , , , , , 1 . 1 1 1 . 1 9 - 1m.11. . , . 7 CD .3 .2 ./ 0 0 3 z5 3 1 5 . 0 A 22.5 30.0 b 45.0
l l l l l l
.8 10 1.2 14 -.2 0 .2 4 .6 16 L 8 CL (a) F4; x/c = 5.715.
Figure 16.- Effect of forewing-flap deflection o n t h e aerodynamic characteristics. BIWHIVNN; A = 16O; 6 h = 0 ' ; iw = 00.
(a) Concluded.
Figure 16.- Continued.
0 0 030.0 -.2 0 .2 1 . 6 .8 L 6 1.8 20 22 ( b ) Fg; X / C = 5.735.
Figure 16.- Continued.
--022.5 -- -030.0 E I Z -?2 0 .2 4 .6 .8 1 0 12 14 / 1 6 -4 0 4 8 I2 I6 20 24 (b) Concluded.
Figure 16.- Concluded.
I 1 1 1 1 1 1 1 1 1 1 1 II
I
I I
.8 I
i
.7
j
.6 I .5 b P .4 c ,
r I
! I
.2
b 1
.I -7 L 2 L 4 L 6 1.8 20 22 -2 0 .2 . 4 .6 .8 LO CL (a) H1.
Figure 17.- Effect of canard size o n the aerodynamic characteristics. B I W H I V N N F 1 ; A = 16O; 4 , = Oo; 6 , = 00; 4 = 00; i , = W.
4a
I
em . 2 -4 - . 6 -. 8 - 10 -4 0 4 8 I2 I6 20 24 --2 0 .2 4 .6 .8 LO L 2 14 1 6 (a) Concluded.
Figure 17.- Continued.
. 7 .5 C D .3 ./ C 0 O f f 0 CI 0 c 2 A c 3 1 6 L 8 20 22 4 . 6 -.2 0 .2 (b) Horizontal tail off.
Figure 17.- Continued.
(b) Concluded.
Figure 17.- Concluded.
Ill I
- 4 0 4 8 /2 /6 20 24
0, deg
Figure 18.- Effect of canard Size on the increment in pitching moment and lift coefficients due to t h e addition of t h e horizontal tail.
B ~ W H ~ V N N F ~ ; A = 16O; 6 , = 00; 6 , = 00; 6f = Oo; iw = @.
^^ /6 /2 =, deq -4 -2 0 .2 4 .6 .8 M / 1 6 18 2.0 2.2 (a) C1; x/c = 5.615.
Figure 19.- Effect of canard deflection on the aerodynamic characteristics. B1W2H1VNNF1: A = 16O; 6 h = 00; 6f = Oo; iw = @.
(a) Concluded.
Figure 19.- Continued.
zc I
I
i I
I
E
I
t
4 f
-4 u .2 4 . 6 .8 L O /.2 L 4 L 6 /.8 20 2.2 CL (b) Cp; X / C = 5.515.
Figure 19.- Continued.
(b) Concluded.
Figure 19.- Continued.
i
4i
OI I *'4 1 . 6 .8 LO LZ L 4 L 6 L 8 20 22 CL (c) C3; X/C = 5.401.
Figure 19.- Continued.
(c) Concluded.
Figure 19.- Concluded.
zn / / .6 .5 .4 .3 .2 ./ c -4 .z I . 6 . 8 L O L Z /14 L 6 1.8 20 22 CL Figure 20.- Effect of fuselage cross section on the aerodynamic characteristics. BWH1VNNF6; A = 16'; &h = Oo; 4 = 22.5O; x/c = 5.735; i , = Oo.
I Q) Figure 20.- Concluded.
. 8 . 7 . 6 .5 . 4 c, .3 .2 i ./ L 0 -2 - 4 2s 2L a,deq 8 -4 -00 -.6 - 4 -. 2 0 .2 4 . ( L O L 2 1 1 4 1 1 6 1 8 L Figure 21.- Effect of horizontal-tail size o n t h e aerodynamic characteristics. BIWHVNNF1; A = 16O; 9 = IOo; 6 h = Oo; x/c = 5.715; iw = 0 ' .
I.
. .
H Off HI H2 H3 -- -.6 -4 e.2 0 .2 4 .6 .8 l.0 12 14 l.6 7 -8 -4 0 4 8 I 2 I6 20 24 CL Figure 21.- Concluded.
Figure 22.- I n c r e m e n t in pitching-moment and l i f t coefficients due to t h e addition of t h e horizontal tail. B ~ W H V N N F ~ : A = 16O; df = loo; iw = 0 0 .
.8 I6 L
i"
0 O f f 0 0 0 -5 0 .2 .8 6 0 6 2 64 LB Figure 23.- Effect of horizontal t a i l o n t h e aerodynamic characteristics. BlWHVNNF1; A = 16'; x / c = 5.715; iw = 0 ' .
-.6 -4 m . 2 0 .2 4 . 6 .8 l.0 L 2 14 1.6 -% -4 , 0 4 8 I2 16 20 24
CL Q'e
(a) Concluded.
Figure 23.- Continued.
CD .3 .2 ./ P v .+ .v .v L O 1 4 / 1 6 L 8 - . 6 -4 72 O .2 (b) H I ; 6f = 10'.
Figure 23.- Continued.
(b) Concluded.
Figure 23.- Continued.
.8 . 7 . 6 .5 CO .3 .2 - " : 6 -4 2 0 2 4 6 .8 b 0 L 2 (cl H3; 6f = 10'.
Figure 23.- Continued.
(c) Concluded.
Figure 23.- Concluded.
3 C D I ?
Of a -.2 0 .2 4 .6 . 8 L O L 2 L 4 1 . 6 L 8 20 22 CL Figure 24.- Effect of horizontal t a i l on t h e aerodynamic characteristics. B ~ W H ~ V N N F ~ ; A = 16'; af = 450; x/c = 5.715; i , = 0 ' .
.2 -4 -.6
- .8
- ' < 2 0 .2 4 .6 .8 LO 12 14 16 -4 0 4 8 I2 I6 20 24 Figure 24.- Concluded.
- * 01
-.02
-. 04
-. 05
-. 06
-4 0 4 8 /2 16 20 24
a, dep
Figure 25.- Effect of horizontal-tail size and forewing-flap deflection o n the horizontal-control parameter dC,,,b6h. BIWHVNNF. A = 16O; iw = 0'.
I
-8 -4 0 4 8 /2 /6 20 24
4, deg Figure 26.- Effect of fuselage nose cross-sectional shape and vertical t a i l on t h e lateral directional stability characteristics in pitch.
h = 16'; oh = 0 ' ; 6 f = 0 ' ; i , = 0 ' .
NASA-Langley, 1966 L-4933 b .
“ T h e aeronautical and space activities of t h e United States shall be
conducted so as t o contribute . . . t o the expansion of humdn knowl-
edge of phenomena i n the atmosphere and space. T h e Administration shall provide f o r the widest practicable avd approprjdte dissemination of information concerning its actisities and the results thereof .” -NATIONAL AERONAUTICS AND SPACE ACT OF 1958
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