Document
NASA TECHNICAL NOTE
NASA TN D-6573
!
Z Z
FULL-SCALE WIND-TUNNEL TESTS
OF A SMALL UNPOWERED JET
AIRCRAFT WITH A T-TAIL
by Paul T. Soderman and Thomas N. Aiken
Ames Research Center
and
U.S. Army Air Mobility R&D Laboratory
Moffett Field, Calif. 94035
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION • WASHINGTON, D. C. • NOVEMBER 1971 .L i I. Report No. _ 2. Government Accession No. 3. Recipient's Catalog No.
NASA TN D-6573
I
4. Title and Subtitle 5. Report Date FULL-SCALE WIND-TUNNEL TFSTS OF A SMALL UNPOWERED November 1971 JET AIRCRAFT WITH A T-TAIL 6. Performing Organization Code 7. Author(sl Performing Organization Report No.
A-3135 Paul T. Soderman and Thomas N. Aiken 10. Work Unit No.
9. Performing Organization Name and Address 126-13-01-43-00-21 Ames Research Center, NASA 11. Contract or Grant No.
Moffett Field, Calif., 94035 13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address Technical Note National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington, D. C., 20546 15, Supptementary Notes 16. Abstract The aerodynamic characteristics of a full-scale execfitive type jet transport aircraft with a T-tail were investigated in the Ames 40- by 80-Foot (12.2- by 24.4-meter) Wind Tunnel (subsonic). Static longitudinal and lateral stability and control characteristics were determined at angles of attack from -2 ° to +42 ° .
The aircraft wing had 13 ° of sweep and an aspect ratio of 5.02. The aircraft was tested power off with various wing leading- and trailing-edge high-lift devices. The basic conf'tguration was tested with and without such components as engine nacelles, wing tip tanks and empannage_ Hinge-moment data were obtained and downwash angles in the horizontal-tail plane location were calculated. The data were obtained at Reynolds numbers of4.1×106 and 8.7X106 based on wing mean aerodynamic chord.
The model had static longitudinal stability through initial stall. Severe tail buffet occurred near the angle of attack for maximum lift. Above initial stall the aircraft had pronounced pitch-up, characteristic ofT-tail configurations. A stable trim point was possible at angles of attack between 30 ° and 40 ° (depending on e.g. location and flap setting).
Hinge-moment data showed no regions that would result in adverse effects on stick force. Comparisons of wind-tunnel data and flight-test data are presented.
! 17. Key Words (Suggested by Author(s)) 18, Distribution Statement Deep stall Unclassified Unlimited T-tail aircraft General aviation 22. Price" 19. Security Classif. (of this report) / 20. Security Classif. (of this page) 21. No. of Pages $3.00 Unclassified l Unclassified 99 ii For sale by the National Technical Information Service, Springfield, Virginia 22151 NOTATION b wing span, 10.40 in (34.1 ft) c wing chord measured parallel to the plane of symmetry, m (ft) _ /" b/2 c2 dy, 2.14 m (7.04 ft) mean aerodynamic chord, S ,,o drag coefficient (wind axes), drag CD qooS rolling moment CI rolling-moment coefficient about stability axis, OCl qooSb
%
_"7" lateral stability parameter, per deg ac__2/ aileron effectiveness parameter, per deg CI6 a aCSa lift CL lift coefficient (wind axes), qooS OCL CLfi _8"-_ flap effectiveness parameter, per deg d pitching moment Cm pitching-moment coefficient about _ (stability axes), qooSd • ,CJ_n, longitudinal stability parameter, per dog Cmo_ _)o_ Cn yawing-moment coefficient about moment center shown in figure 2(a) (stability axes), yawing moment qooSb aCn directional stability parameter, per deg Cnb3 a# OCn rudder effectiveness parameter, per deg Clt81.
OSr side force Cy side-force coefficient (wind axes), q_S it horizontal-tail incidence angle, deg q dynamic pressure, N/m 2 (lb/sq ft) Vood R Reynolds number, P S wing area, 21.50 m 2 (231.77 ft 2) V_ free-stream velocity, m/sec (ft/sec) spanwise distance perpendicular to tile plane of symmetry, m (ft) Y O_ angle of attack of fuselage, deg angle of sideslip, deg; positive - nose to left # _Sa trailing-edge aileron deflection angle, deg; positive - left aileron trailing edge down 8e elevator deflection angle, deg; positive - trailing edge down 8f trailing-edge flap deflection measured from wing chord line, deg 8r rudder deflection angle, deg; positive - trailing edge left 8s spoiler deflection angle, deg average downwash at the tail location with respect to free stream, deg Y r7 wing semispan station, sweep angle of quarter-chord line, 13° At/4 v free-stream kinematic viscosity, na 2/sec (ft 2/sec) leading-edge contours defined on figure 2(d) Subscripts L left max maximum R right t tail u uncorrected A change free stream Hinge Moments of positive Positive hinge moments tend to move the control surface in the direction deflection. The average chord aft of the hinge line was used for the reference length.
iv
Aileron
hinge moment where Sa = 0.544 m z (5.85 ft 2) Cha = qSada da =0.38 m (1.24 ft) Rudder hinge moment where Sr = 0.609 m 2 (6.56 ft 2) Chr = qSrdr dr = 0.46 m (1.51 ft) Elevator hinge moment where Se =0.635 m 2 (6.83 ft 2) Che = qSede de =0.29 m (0.96 ft) Horizontal stabilizer hinge moment where S h = 5.02 m 2 (54.0 ft 2) Chh- qShdh eh =l.17m(3.83ft) Note: Se is the area of the right or left elevator; S h is the total area of the horizontal stabilizer.
FULL-SCALE WIND-TUNNEL TESTS OF A SMALL
UNPOWERED JET AIRCRAFT WITH A T-TAlL Paul T. Soderman and Thomas N. Aiken Ames Research Center and U.S. Army Air Mobility R&D Laboratory Moffett Field, California 94035 SUMMARY The aerodynamic characteristics of a full-scale executive type jet transport aircraft with a T-tail were investigated in the Ames 40- by 80-Foot (12.2- by 24.4-m) Wind Tunnel (subsonic).
Static longitudinal and lateral stability and control characteristics were determined at angles of attack from -2 ° to +42 ° .
The aircraft wing had 13 ° of sweep and an aspect ratio of 5.02. The aircraft was tested power off with various wing leading- and trailing-edge high-lift devices. The basic configuration was tested with and without such components as engine nacelles, wing tip tanks, and empennage.
Hinge-moment data were obtained and downwash angles in the horizontal-tail plane location were calculated. The data were obtained at Reynolds numbers of 4.1XI06 and 8.7X106 based on wing mean aerodynamic chord.
The model had static longitudinal stability through initial stall. Severe tail buffet occurred near the angle of attack for maximum lift. Above initial stall the aircraft had pronounced pitch-up, characteristics of T-tail configurations. A stable trim point was possible at angles of attack between 30 ° and 40 ° (depending on c.g. location and flap setting).
Hinge-moment data showed no regions that would result in adverse effects on stick force.
Comparisons of wind-tunnel data and flight-test data are presented.
INTRODUCTION Most small aircraft, including executive jet transports, are designed with a minimum of wind-tunnel data. Furthermore, flight tests are likely to be qualitative rather than quantitative. As a result, the designer has little opportunity to verify his design predictions.
Therefore, to aid designers the present investigation was conducted to determine the static longitudinal and lateral stability and control characteristics through deep stall of a full-scale executive jet aircraft. The deep stall testing was conducted to see if the aircraft exhibited unfavorable characteristics at high angles of attack because of its T-tail. Some of these problems and relatedresearch can be found in references 1 through4. Unfortunately,it cannotbe determined
from wind-tunneltestsof unpowered aircraft whetherthe poweredaircraft canbecome lockedin
deepstall.
AIRCRAFT AND APPARATUS
In figuresl(a) and (b) the model is shownmountedin the Ames40- by 80-Foot(12.2-by
24.4-m) Wind Tunnel. Pertinent dimensionsof the basic model configurationsare given in
figures2(a)and(b).
Wing
Thewinghada quarterchordsweep of 13 °, an aspect ratio of 5.02, a taper ratio of 0.507, and a dihedral angle of 2.5 °. The airfoil section was an NACA 64A 109 modified by increased camber and chord at the leading edge (fig. 2(d)) which was minimum at the root and maximum at the wing-tip tank junction.
High Lift Devices Flap details- The basic wing had a single slotted, extendable (Fowler) flap (fig. 2(c)) located from the edge of the fuselage at 7.1 percent to 61.2 percent r/. Maximum flap angle was 40 ° at the lower Reynolds number and 38 ° at the higher Reynolds number because of air load effects. A center section flap that extended under the fuselage was tested (fig. l(b)). There were no gaps between the sides of the center section flap and the main flaps.
Leading-edge contours- The drooped leading edges of the basic wing were removed, part way through the test, and replaced by various leading-edge contours (fig. 2(d)). The dimensions of the leading edges varied linearly from root to tip.
Wing plan form modification- In an attempt to delay the stalling of the wing tip region, fence, were placed first on the tops and then on the sides of the tip tanks (fig. 2(e)).
Lateral Controls Ailerons- The ailerons (fig. 2(b)) had relatively blunt leading edges and balance tabs to decrease stick force. As the ailerons were moved, the balance tabs moved in the opposite direction such that e_tab = -(5/6)6a where C_ta b is the tab angle relative to the aileron chord and 6a is the aileron angle relative to the wing chord.
Spoilers- The chords of the spoilers were 10.25 percent of the wing chord at midspoiler and were located from 22.2 percent to 49.4 percent semispan (see fig. 2(b)). Spoiler angles ranged from 0 ° to 42 °. In addition to the basic wing spoilers, dummy spoilers were tested outboard (fig. 2(f)).
Tail The geometry of the horizontal and vertical tails is described in figure 2(g). Pitch control was provided by an all-movable tail that had an available deflection range of 0.4 ° to -7.0 ° and by a 32 percent chord elevator with balance horn. The elevator angle was variable from 15 ° to -15 °. The rudder (25 percent chord) had a deflection range of 30 ° to-30 ° and had a trim tab that was locked at 0 °. The horizontal stabilizer was used for aircraft trim.
Nacelles Engine nacelle detail and location are shown in figure 2(h). A constant-area circular duct was installed in each nacelle to allow mass flow conditions of 4.81 kg/sec (10.6 lb/sec) of air at standard conditions, similar to that of the jet engines for idle airflow at a Mach number of 0.2. Static and total pressures were measured with rakes at the aft ends of the ducts to determine the actual dynamic pressure of the nacelle flow and the internal naceIle drag (which was removed from the data). The nacelles were removed from the pylons during a part of the test.
Tip Tanks Wing tip tank detail and location are shown in figure 2(b). All data are presented with the tip tanks on unless stated otherwise.
TESTING AND PROCEDURE Forces and moments were measured for the model through an angle-of-attack range from -2 ° to 42 °. Pitching-moment data were computed about a moment center location at 25 percent & The center-of-gravity range for this aircraft is 16 percent C to 31.5 percent _, Tests were conducted at Reynolds numbers of 4.1×106 and 8.7X10 6 based on a mean aerodynamic chord of 2.14 m (7.04 ft) and speeds of 27.8 m/see (54.2 knots) and 59.0 m/sec (115.0 knots), respectively. These speeds correspond to dynamic pressures of 478.8 N/m s (q = 10 psf) and 2156 N/m 2 (q = 45 psf).
Tests were conducted with the basic configuration _ at several tail incidences with variable elevator, rudder, aileron, spoiler, and flap settings. Data were also obtained with landing gear down.
The maximum angle of attack at R=8.7×106 was 16 ° (tail on) because of tail buffet load limitations. Most data, tail on, at angles of attack higher than 16 ° were taken at a Reynolds number of4.1×106 .
_Basic configuration refers to the airplane as shown in figure l(a) with engine nacelles, tip tanks, and empennage on model. Control surfaces were at zero angle unless stated otherwise.
DATA ACQUISITIONAND REDUCTION
Forcesandmomentsweremeasured on thewind-tunnelsix-component balance. Torquetubes
in the elevators andrudderweregaged to providehinge-moment data.
All datawerecorrected for strut tares, nacelle internal flow drag, and wind-tunnel wall effects.
Nacelle internal flow drag was calculated from pressure measurements in the nacelle ducts, and AC D = 0.0005 cos a was subtracted from model drag. Corrections added for wind-tunnel wall effects were Ac_ = 0.506 CLu AC D = 0.0088 CLu 2 ACm = 0.0171 CLu (tail on runs only) ACCURACY OF MEASUREMENT The various quantities measured were accurate within the following limits, which include error limits involved in calibrating, reading, and reducing the data.
_+0.2° Angle of attack +0.5 ° Angle of sideslip Free-stream dynamic pressure -+0.5 percent -+0.5° Control surface settings Coefficients at Force or moment R = 8.7×10 6 Lift +22.4 N (+5 lb) +0.0005 Drag -+13.4 N (+3 lb) _+.0003 Side force _+13.4 N (+3 lb) +.0003 Pitching moment +271 J (+200 ft-lb) -+.0027 Yawing moment -+136 J (_+100 ft-lb) +.0003 Rolling moment --.475 J (+350 ft-lb) _+.0010 RESULTS Table 1 is the index to the figures. The longitudinal data are presented in figures 3 to 18 and the lateral data in figures 19 to 30. Downwash and hinge-moment data are given in figures 31 and 32, respectively.
DISCUSSION
LongitudinalCharacteristics
Flap effectiveness- The longitudinal characteristics of the basic airplane at R = 8.7×106 with three flap settings are shown in figure 3(a). The flap effectiveness parameter, CL6, was 0.015/deg for the 20 ° flap setting and 0.013/deg for the 38 ° flap setting. A theoretical flap effectiveness estimate was made using the simplified lifting-surface theory of reference 5, which gave the value of CL6 as 0.022/deg, almost 60 percent higher than measured. This discrepancy was probably due to a nonoptimum gap setting for the single-slotted, Fowler type flaps. A comparison of small-scale with full-scale wind-tunnel data to be discussed in a later section shows that small-scale flap effectiveness is closer to the theoretical value. This suggests that the flap gap choice was based on small-scale test data and not corrected properly for full-scale Reynolds number effects.
Maximum lift- Figure 3(a) shows the basic stall characteristics of the aircraft at R = 8.7× 106 .
Because of severe buffet on the tail as it penetrated the wing wake, the tail was guy-wired as shown in figure l(a). 2 In addition, some of the data were taken at a reduced Reynolds number of 4.1×106. The tail buffet acted as a strong stall warning. Figure 3(b) shows the longitudinal characteristics at R = 4.1X 106 . Increasing the Reynolds number from 4.1 × 106 to 8.7× 106 caused an increase in maximum lift coefficient of 0.19 (flaps down) and 0.20 (flaps up) as shown in figure 4. The high Reynolds number condition is closer to actual flight conditions. Observation of tufts on the left wing indicated that a region of separated flow developed near the wing leading edge tip tank junction at 8 ° angle of attack (this did not happen with tip tanks off). As angle of attack was increased the region of separated flow spread aft and inboard. Near CLmax the wing root began to stall. Both regions grew with angle of attack until most of the wing stalled and lift dropped.
Static stability- A study of the variation of the stick-fixed pitching-moment coefficient with angle of attack (fig. 5(a)) shows that the airplane was stable through maximum lift (even for aft c.g.
limit of 31.5 percent e). Above maximum lift, the classic deep stall situation occurred that will be discussed later. The data presented for c.g. at 25 percent d give Cma = -0.0186/deg. At stall the aircraft experienced a slight nose down pitching moment. The stick-free static stability characteristics, determined from hinge-moment and pitching-moment curves, are shown in figure 5(b) (data are shown for c.g. at 25 percent d). Freeing the elevator reduced the stability, but the aircraft did not become unstable. For the aft c.g. case (31.5 percent d), 6 f = 0°, o_ = 0 °, Cmo_ was reduced from -0.0185 to -0.005/deg.
Deep stall- As illustrated in figure 6, the airplane was unstable above maximum lift (stick-fixed) with the center of gravity at the quarter chord, and maximum nose-down trim until an angle of attack of 32 ° was reached at which point static stability was again attained. Furthermore, the pitching moments became zero or slightly positive above a = 28 ° flaps down. Thus it may be possible (at Iow Reynolds number) for the airplane to reach a region of positive pitching moment and pitch up to e_= 32 °, a trim condition (power off) if the pilot does not take corrective action.
However, aircraft rolloff may preclude this possibility, as will be discussed in a later section. As shown by the axes superimposed on figure 6(b), at forward c.g. the pitching moments do not become positive, but at the aft c.g. the aircraft would reach the positive pitching-moment region 2The wires had very little effect on the data.
sooner and could pitch up to trim at o_= 39 ° (flaps up or down) while completely stalled.
Figure 6(c) shows that while the effect of sideslip was beneficial, 8 ° of sideslip changed the pitching moment only 0.06 at c_= 32 ° .
With the flaps up, elevator effectiveness was maintained at all angles of attack but pitching-moment increment due to full elevator deflection at angles of attack greater than 24 ° is approximately one fourth that at angles of attack below stall (see fig. 7(a)). Therefore, recovery from deep stall (flaps up, c.g. at 25 percent 6) would be possible using the elevator, but the time it takes to rotate the nose down may be long. With the c.g. in the aft location there is insufficient elevator effectiveness to recover from deep stall. With the flaps full down (fig. 7(b)) there was an almost complete loss of elevator control power above o_ = 24 °. Since the data (flaps up and down) were taken with the horizontal stabilizer leading edge full up, any movement of that control surface would only make the pitching moments more positive.
Figures 8 and 9 illustrate the effects of horizontal stabilizer incidence and removal of the empennage, respectively, on the longitudinal characteristics.
Effect of wing tip tanks, engine nacelles, and landing gear- Figure I0 shows that the wing tip tanks caused an increase in lift coefficient and lift curve slope primarily because of the increased wing area and aspect ratio (reference area was not changed). The drag change was small tip to CLmax. The addition of the tip tanks made the aircraft slightly more stable in pitch.
The engine nacelles caused a decrease in lift, especially with flaps down (fig. 1 1). This decrease was probably due to interference with flow around the wing that redtlced wing lift since the nacelles did not develop negative lift or reduce tail lift. This explanation is substantiated by the increase in nose-down pitching moment with the nacelles on the aircraft. If the nacelles had developed negative lift or if the tail lift had been reduced, the pitching-moment change would have been nose up. The fact that the wing tips were probably not affected by the nacelles accounts for the nose-down pitching-moment change (i.e., the lift loss was inboard).
The landing gear effect on the longitudinal characteristics is small (fig. 12).
High-lift devices- The effects of four wing leading edges are given in figures 13(a) and (b), flaps up and down. For the flap down case the leading edge 14, which had the greatest droop, increased maximum lift beyond the value achieved by l 3 , the basic configuration leading edge.
In an attempt to improve the CLmax of the airplane, fences were placed on the tops and, later, sides of the tip tanks to alleviate flow separation at the junction of the tip tank and wing.
Fences on the sides of the tip tanks caused an increase in lift due to the increased wing area and aspect ratio (fig. 14). In no case was the flow separation alleviated near the tip.
The center body flap (fig. l(b)) caused a very small reduction in lift and drag of the model and a very slight change in pitching moment (fig. 15). The reason for the reduction in lift and drag is unknown.
Drooping the ailerons 13.7 ° increased maximum lift coefficient by 0.1 (fig. 16). Since drooped ailerons reduce roll control, outboard spoilers were tested. These will be discussed in the lateral control section.
EfJ_,ct of spoilers- Runs were made with various right and left spoiler deflections (see figs. 17(a) and (b)). The deflection of one or both spoilers 42 ° caused a nose-down pitching-moment change probably because of an induced increase of tail angle of attack. This supposition checks with figure 17(c) that shows very little change in pitching moment with outboard spoiler deflection. It was expected that the flow field of the tail would not be affected greatly by deflection of the outboard spoilers. The drag was increased 80 percent with full spoiler deflection.
Comparison of wind-tunnel and flight-test data- A comparison of Ames 40- by 80-Foot (12.2- by 24.4-m) Wind Tunnel data, Wichita State University 7-by 10-Foot (2.1-by 3.l-m)Wind Tunnel data and Lear Jet Flight-test data is made in reference 6. Two figures from that paper are presented in this report as figures 18(a) and (b). Results show good agreement between full-scale wind-tunnel and flight-test data. Reynolds number effects account for most of the difference between small-scale and full-scale results.
Lateral and Directional Stability and Control The lateral characteristics of the airplane are shown in figures 19 to 23, and lateral and directional control effectiveness in figures 24 to 29. Stability derivatives Cn/3 and C//3 are plotted versus angle of attack in figure 30. These data show that the airplane had positive effective dihedral (-CI/3) over the normal operating range and was directionally stable statically (positive Cn/3). With the tail removed (fig. 22) the nonzero rolling moment and side force at/3 = 0 ° were probably due to flap misalinement. The flaps had been removed and reinstalled on the model prior to these runs.
The data in figure 23 show that as the model stalled with flaps up, it tried to roll left (left wind down) and with flaps down, it tried to roll right (right wing down). The change in roll direction at stall was probably caused by asymmetric deflection of the flaps. The rolling moment, flaps down, was greater than that produced by full opposite aileron deflection. This severe rolloff in stall would complicate recovery, but it might prevent a deep stall condition.
Control effectiveness- Aileron roll power was fairly constant below stall but decreased rapidly in stall (fig. 24). The airplane had slight favorable yaw due to aileron above 6 ° angle of attack.
Figures 24(b) (d) show the control power due to one aileron. The nonzero side force was probably due to model misalinement in the test section. Figure 25 is a summary plot of C16 a versus angle of attack.
Rudder deflection affected the longitudinal characteristics very little. Figure 26 shows the lateral effects of rudder deflection. The rudder was capable of holding the airplane in sideslip between-15 ° --.</3_< 15 °.
The control power of the basic spoiler is shown in figure 28(a) as plots of Cy, Cn, and C l versus left spoiler angle (right spoiler full down). Figure 28(b) shows the effectiveness of dummy outboard spoilers S_, $2, and $3. These spoilers were more effective than ailerons or inboard spoilers for lateral control. The lateral characteristics of the airplane with the landing gear extended are shown in figure 29. Comparison with the results in figure 19(a) (landing gear retracted) indicates that the landing gear had a small effect on Cy vs./3 but only a slight effect on Cn and C l vs./3.
! \
!
Downwash at the Horizontal Tail An average downwash angle at the horizontal stabilizer was estimated from curves of Cm vs. ot for several tail incidence angles. The intersection of the tail-on curves with the tail-off curve are points where tail lift is zero; and for a symmetrical horizontal stabilizer e=o_+i t Figure 31 shows the results of the above calculation, which were identical for both Reynolds number cases.
Hinge Moments Typical curves of hinge-moment coefficient C h versus angle of attack and C h versus control position are presented in figures 32(a)-(h) for aileron, elevator, rudder, and horizontal stabilizer.
The data were obtained at R = 8.7X 106 to approximate actual flight conditions. These results show no control force reversal for any of the controls within the normal operating range.
CONCLUSIONS A full-scale wind-tunnel investigation was made of a small jet aircraft with a T-tail to determine the longitudinal and lateral stability and control characteristics through deep stall, power off. The following conclusions were drawn from the results of the investigation: I. The airplane had stick-fixed static longitudinal stability at angles of attack up to stall for the full c.g. range. With the stick free, stability was reduced but the aircraft did not become unstable.
2. Before stall the tail experienced severe buffet as it penetrated the wing wake, and, in stall, the airplane tended to roll right wing down or left wing down depending on flap angle. The tail buffet acted as a strong stall warning, that might prevent deep stall entry during actual flight conditions. However, the rolling moment in stall, flaps down, was greater than that produced by full opposite aileron deflection.
3. Above stall, the airplane was unstable in pitch, and the pitching moments could become positive, depending on c.g. A trim condition in deep stall (a = 39 °) with a large reduction in elevator control was possible. With the c.g. in the aft position, elevator control and horizontal stabilizer control were insufficient for recovery from deep stall trim.
4. The airplane was directionally stable, below stall, and had positive effective dihedral.
Ames Research Center National Aeronautics and Space Administration Moffett Field, Calif., 94035, July 6, 1971 REFERENCES 1. Aoyagi, Kiyoshi; and Tolhurst, William H., Jr.: Large-Scale Wind-Tunnel Tests of a Subsonic Transport With Aft Engine Nacelles and High Tail. NASA TN D 3797, 1967.
2. Ray, Edward J.; and Taylor, Robert T.: Effect of Configuration Variables on the Subsonic Longitudinal Stability Characteristics of a High-Tail Transport Configuration. NASA TM X-1165, 1965.
3. Trubshaw, E. B.: Low Speed Handling With Special Reference to the Super Stall. J. Roy. Aeronaut. Soc., vol. 70, no. 667, pp. 695 704, July 1966.
4. Thomas, H. H. B. M.: A Study of the Longitudinal Behavior of an Aircraft Near-Stall and Post-Stall Conditions.
R.A.E. TM Aero 953, 1966.
5. DeYoung, John: Theoretical Symmetric Span Loading Due to Flap Deflection for Wings of Arbitrary Plan Form at Subsonic Speeds. NASA Rep. 1071,1952.
6. Neal, Ronald D.: Correlation of Small-Scale and Full-Scale Wind-Tunnel Data With Flight Test Data on the Lear Jet Model 23. Paper 700237 presented at SAE National Business Aircraft Meeting (Wichita, Kansas), March 1970.
TABLE 1.-INDEX TO FIGURES Figure The model in the wind tunnel .................................... !
Lateral characteristics versus lift coefficient ............................ 2 I E-- _D cq .0 LT_ oo o < o o E i P_ ll (b) Center section flap, nose and tip booms on model.
Figure 1.- Concluded.
3_2 Horz. Vert.
Wing tail tail Aspect ratio 5.02 4.0 -- Taper ratio 0.507 0.469 3.495 21,51 5.02. 34.48 Area, sq m (2],1.77) (54.00) 34.48) Sweep, deg I 3 25 40 (25% Airfoil section 64A-10964A008 64A010 All dimensions in meters (feet) i)5.86 __ 2.47 (8. _o) 4.47 (14.6 7) Moment C/4 Line 2.5 ° Dihedral 2.51 (8.25) 10.40 (54.10) (a) General arrangement of model.
Figure 2.- Geometric details of the model.
4.40 (14.45) (4.57) = I .39 i / 0.431 / (I / Aileron 0.55 l hinge line (I.75) 78% c -_ (4.91) 13° 10.76 5 0 35) 0.24 i (o.
/
/ 73% c / / t Flap leading edge Moment 0.37 center
J25 '
(I.2 I) i (0.93)(I .39) __ Aircraft 2.75 (9.02) (b) Basic wing detail.
Figure 2.- Continued.
3_4 0.268c Flap angle, deg Gap,cm(in) 0 0 2O 1.57 (0.62) 13.97(5.5) 2.46(0.97) 18.55 (7.3) 4O 2.46(0.97) 18.80(7.4) (c) Trailing-edge flap.
Figure 2.- Continued.
Zl 1,z (STD NACA (blunt) _- -'/ 64A-109) / _/-J /6% Chordline Wing and tip tank junction, cm (in) X _f J of STD NACA ×,:,o._, <.,.,> ,,-,.,,,o.,>-/tc--_. "' x, ;
t 4 / --_×_ -
configuration) _ " _ Hinge point of _4 Wing and fuselage junction, cm (in) (STD NACA 64A-109) k drooped 30 ° # Chordline ×1 =8.64(3.4) R l =2.28(0.9) of STD NACA X2=13.71 (5.4) R z = 1.78(0.7) 1,1 linch Scale: I I cm (d) Leading-edge modifications.
Figure 2.- Continued.
I_ 1.61 - (5.28)
---f-
0.48 (I.58) F Area =(5.81 ff 2) 0.18(0.58) / l Side view 1.61 (5.28) 0.48 0.54m z (I.58) Area = 5.81 ft 2 0.1 Top view (e) Fence located on side and top of tip tank.
Figure 2.- Continued.
I
!
A 70 % chord .-7 Symbol "A" dimension,(in) cm Si (2.25) 5.71 $2 (5.50)13.98 S3 (6.75)17.15 (f) Outboard spoiler.
Figure 2.- Continued.
0.45(I / --_ 1.67 0.75 (5.46) (0.83) (4.33) 0.74 (2.42)
/
i.32 1 / P 1.04 1.43 (4.6 9) (3.42) Fuseloge Ventral fin
L 2.80
v I r (9.17) C L Aircraft mom 5_ Hinge , 1.52 pin .JIL(5.OO) o.18 (o.58) - I 0'.52 (I. 70)
I
J
2.24 0.25 (o.75)L -I (7.35) (g) Horizontal and vertical stabilizer.
Figure 2.- Continued. 3-9 (3'61 .I I 4-)
it
0.76 0.63 O.
(2.50)(2.08)i --(1_ line Cross-section ot engine inlet o-- 0.66 (2_.17) 3 0.60__ I 9-( I, 98)-" ]
Lo.3eJ
Cross-section ot wing leading edge-fuselage junction (h) Fuselage-wing cross section at two locations.
Figure 2.- Concluded.
2O I oJ I E o OJUD • m O4 o CO o 6 o _t" II o x r_ o II ej ._ I
_.o
[.l _D (M C_ (I# O0-o II _c_ o X OE]o
/
a
/
---o (M 0 00 _ _- cu 0 co _ _.
_J C)
T
I" E 0 " _5 II _) _8 m o 0 _ o o _) I 0 --001"-- m _ × _-_ !
n _ ODO
P
I e_ ._ !
GO _) _) _1 (_) I" E o _D 0,I 0 CO _- r_ 8O Q.)
_0 o '_ I_ O[3 _.)
°..._ LoO • I "---'-----0 O,J o 8f, deg 0 0 Q 2O 0 38
,i! !
I0
/ /
a,deg
i
0 .4 .2 -.2 -.4 Cm (a) Stick fixed.
Figure 5.-Variation of pitching-moment coefficient with angle of attack; R = 8.7X 10 6 , i t = 0.4 °.
_o
!
od
/
QO -.I
/
_0 II II .,e- CO z o 0o "_ II II
/ °
0 0 0
o o o o
I I I I !
E I* E ¢M E) O0 oJ _ _D _ 0 (M 6ep _e r,O ,--, 0 0 ,_ II
"'---o
x _ H N ,.-,, ._ o
o_
on
N oJ co _D 0 E) (9 o.
.-I I E Oo 00_ _- cO cO 0 ed ,_- _0 00 0 I Sap '_ if)
%
c5 .6 ¢) II 4...a o ,...q 4.-, o X oJ ,..., II ,ID o
,,oo
¢-_ OH o QO o,J o o,J _0 _ c_J 0 cO co _1- cw 0 I _J z.)
I
%
E O_ OD 0 _- _ CO 0 cO • I_0 6J -- I 6ap ' "o 0_1 o aj II 0 fj i X _o _. o II o o II "_ 0 ,_'- O0 GO on<> Od :_'---'o" '---43D- ---..o_ .--..¢_D.- --,t3 o _t cJ 0 0D _D ,_- _ 0 d c) cO i ° _k3 E O0 0o _D CO 0 OJ _" _O O0 0 ¢0 T flap '_ o U ',o -I. o II ;> o _ _ !
o _ ,_ -'T _ O0 "1o tO) 0 b 3O I I E o 0 (xJ _T _ co 0 !
_I- I_ (_I -- 6ap '_ r¢3 o @ ,,
g
o_ × o r,.)
!
II _.0 ______0 II I ,.o _0_ 0q.
¢D eq.
o,J 0 I .d (.2 3] !
i ° "o"_-. o........_ ' - "'u. >--..-.-_ !
E 0_1 . _ _ ..,.,._ I o LD II
d 0
X o I-,i II r._ • I o GO ,¢. _ o.
I I _...I I I o E (kl b o II C_ X o !
II o °,.,.4 |I o "c:l 0 l • !
a (.3 I o,I I E oJ
%
II aOo X o I II _0 II 0o "8" I I a 0 cO _ _.1
3_
I" i° o o
Q
E I#) b [] o I o I X b- © II _= E o o II !
!
CO II I [] o o o.
I _I .r: o o o X _ ,--, _ 0 .N o M o o !
OJ I ---<9 E Oo OJ X OJ "1o II o u" GO "_ -- _0 0 e OJ N O0 q- _ q J i ° -I U I OJ I E o o (M (,D i X _ o I.
tl _o • I .__'- _ I I I m. c3 o ._J
%
E Oo _o oo o o4 ,,.0 cO CO ' o O4 I Sap 'a ro c_ t'_ e_
s
X ° ""f _. o o .{3 _ o o _ e4o II -'T E II o Fg_oo (13 !
e; 0 1:3 1:3 h Q 0 e4 oa o co i • ._1 E O_ I
o
oJ % X ¢) II o _.o J o oq ¢) H o,I L_ II q
5"-
o 1:5 (.P
ii'
/6 I 0,I 0,I oJ 0 I" 4O I I _.0__--- .0 ""0 (NI _0 - ,6 oJ X "_ -- t'_ 0 "U n [i co _ x> Q oo_ 000_ o CO (kl I v.
I I
==9<
_.¢- E _____4D- c_=:=== _0 .(xl X II E cO "o E !
• I .._--¢ *- 0000 i_. onO<] O cj o !
I 1,2 I E O o o c5 I1 ""H 4--1 O,100 Z X od II _D O O_1 ¢J ,_ _0 O 00--o o O O O ¢,) I O _6 i, oq._ --H-O--- GD _D _ O,1 O CO OJ o ._1 O _3 o I o II oJ I- X 0_ E O0 II o oo II c oo o o
go
o "1oo (12 I c c o
o.@
"o oJ "o 0 [] O0 o (1) "1o o I rO "o o i-,, _6 U_ !
_q _O _1" oa 0 O0 _ '_" Oa 0 CO ._J I o,1 I E C) O o o II ,,-q • O,.I _D O X O o,I II -O -- O _D w DO O<] _1 O E II u" # °,,._ (3O o o o _.O o !
rq.
& L_ o. _. '_" _ O J 4P i° i • '--.'"---.-.Z E a.)
OJ c: D _.1 o --_ II !
,.0 1-4 tO 0 "-' tO t_ O4 ...1 _6 o,] I" E o,I X o II U) o. -o _ o -_ o Q) II o o_ m U_'._- u_'_ II OD 0 _sn "t_ ,-1 0 _-- o I !
-o,I t o o,I 0 q i m I J 4T .24 .20 0 Basic model [] Fences on .16 q .12 Cy .O8 .O4 .O4 Cn -.04 .04 / C_ -.0_420 -16 -12 -8 -4 0 4 8 B, deg (b) Lateral characteristics with fences on tops of the tip tanks; a = 0 °.
Figure 14.- Concluded.
|l I" o E II 0<..)
(.,,j X II o o,,,_ I1) r_ o ,_,. ,_" • I o r_ i m I o !
II E I x o II o '7 O0 (v-) c_ c,D II ¢,C) c o c_J -- c2) "10 0 0 "0 E -o GO n7 r_ G) "0 o [] c_ ..0 o /- I c_ c_ i o m.
J _0 (M I" E CD CM (M -o. 000 _1 O9 {D /_ o o o <_<> m ,@ II °_ • _ __ m_ X H II !
• I cD GO {D q oa 0 oa _- I • i • .J (D I
_J _L
--EL_ (xl
i °
J
PJ \ oo_ ffl ©D< _ °_..q R o ff × _k i II II • I CO _D oJ 0 ,q to _. oq. o _.1 (.P .52 !
!
E o _D o El -- 0 "-d o r,.)
!
g
P_ _ M t_ II o o II _c_ E • I -4D I _D O0 o _. _. o m i ° --I 2.8 1.4 r r I I I I Flaps up 0 Flight test 1.2 Ames 40 x 80, R = 8.6 x I06 Wichita St. Univ. 7xlO 1.0 R = 1.4XlO 6 F ff 1.6 .8 i //__Sf: 40 ° C L CL 1.2 .6
/ I I
// 0 Model 23 .4 .8 _' flight test i Ames 40 x 80
/
.4: .2 Wichita St. Univ. -- 7xlO I I I -4 4 8 12 16 2O 0 .04 .08 .12 .16 .20 .24 a , deg Co 1.4 1.4
I I /
I I
Flaps 20 ° Flaps full down ! i 1.2 1.2 / I.O I I
,o
I ,8 ......
.8- /, CL CL .6
--1 I
.6 O_ I .4
o
.4. I 0 Flight test Flight test Ames 40 x 80 Ames 40x 80 Wichita St. Univ.-- ,2 Wichita St. Univ.-- .2 7xlO 7xlO 1 I I i I [ I L .08 .12 .16 .20 .24 .04 .08 .04 .12 .16 .20 .24 0 CD CD (a) Lift and drag characteristics.
Figure 18.- Comparison of wind-tunnel and flight-test data (taken from ref. 6).
_4 0.2 I Flops up Tail - on
I
Cm o _..__ _._ _ -0.2 0.2 Tail-off Cm w Ames 40X80, R = 8.6 X 06 0.2 Wichita St. Univ.
Flops full down Toil - on 7X O, R = 1.4XIO 6 - _ ,lib.
C m -0.2 -0.4 Tail-off -0.2 Cm T - 0.41 , 0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 CL (b) Pitching-moment characteristics.
Figure 18.- Concluded.
.24 .2O [ .16 .12 Cy .O8 .O4 8f, deg 0 0 -.04 .O4 C n -.04 -.08.04 -.0_204 -16 -12 -8 -4 0 4 8 .8, deg (a) or= 0°, R = 8.7X106, flaps 0 ° and 38 ° Figure 19.- Lateral characteristics of basic configuration in sideslip.
.24 .20 a, deg 0 0 D 8.4 O 14.6 .16 Cy \ .12 .08 .04 C n -.04 .08 C_ .04 -20 -16 -12 -8 -4 0 4 8 /3, deg (b) $f= 0 °, R = 8.7X106,or = 0 °, 8.4 ° , and 14.6 ° .
Figure 19.- Continued.
.12 a, deg .08!
.O4 Cy -.04 .04 -.08 a,deg Cn -.04 .04 a,deg 24, :32, I1_ C_ -16 - 2 -8 -4 0 4 _, deg (c)6f=0 °,R=4.lX106,e=0°,12°,16°,24 ° 32 ° .58 Figure 19.- Concluded.
,12
.08
Cy
0 (i)--_--_-_ G.--___. _3
-.04
.O2 -.08
C n -.O2 _C_ ,8,deg 0 0 -.04 .O6 [] -4 0 -8 .O4 .O2 C_
E}-----ID-'-
OI -'020 4 8 12 16 20 24 28 52 :56 e, deg (a) 6f=0 °, R=4.1X106,/3=0 °,-4 ° , _8° Figure 20.- Lateral characteristics of basic configuration versus angle of attack.
.O2
.'3 .:>-- <_ Cy -.02 8f , deg [] 2O -.04 <> 38 .O2 C n -.02 .O4 .O2
i
C_ ,.,-I- -.02 0 z 8 2 16 20 24 28 a, deg (b) R = 8.7× 106, f3= 0 °, 6f= 0 ° , 20 ° , 38 ° 60 Figure 20.- Concluded.
q
d
J
o_
o
f
o" q !
X II o II o 0 0"' I-7 [] o I" GO o.
I I 12) OD <]
P_
o o c,l o._
\
E D o.
i O.
n TO q OI ,_J ,O8 O4 Cy mm O 0 0 0 I 0 I [] 8 -.04 f- 0 0 58 | Z_ 8 58
l
.O4 Crl -.04 O8 .O4 C 1
I
-0420 -I( 0 4 8 -12 -8 -z ,8, deg Figure 22.- Lateral characteristics of the model in sideslip with the empennage removed; R = 8.7X 106 .
.O4
.O2
-.04
.04 -.06 ,O2 C n -.02 .O6 Tail 8f, deg .O4 • off 0 [] on 40 • off 40 .O2 __ Oon O Cl -.02 --'040 4 8 12 16 20 24 28 32 36 4O a, deg Figure 23.- Lateral characteristics versus angle of attack, tail on and off; 8f = 0 ° and 40 °, R = 4.1X 106.
.O4 .15,+15 .O2 Cy .O2 -.02 Cn 0 -.02 ,deg +10,0 +5,0 0,0 C_ -15,+15 0 4 8 12 16 20 24 a, deg (a) 6f = 0°, R = 4.1X106 Figure 24.- Aileron effectiveness.
.O2
_o
i i I 4X Cy -.02 -.04 .02 C n -.02 .O6 .04 8aR, deg +15 _--- +10 .O2 C_
Z
_a
-.02 0 4 8 12 16 2O 24 a, deg (b)Sf=40 °, R=4.|X|06 Figure 24.- Continued.
6_
.O5 s .02 $ g Cy .OJ [ t_ .02 .01 Cn 0 - .01 - .02 .O3 .O2 > .01 C_ a, deg © 0 -.0 [] 8.4
0 14.6
- ,02 -15 -I0 -5 0 5 I0 15 BaR, deg (c) 5f= 0 °, R = 8.7X10 e Figure 24.- Continued.
Cy -.02 .02 r Cn .,.._...._...,.--.--_ ---------K -.02 .O4 a,deg [] 8.7 0 12.9 .O2 0 .5 C?.
-.02 D 5 -I0 -5 0 5 I0 BaR, deg (d) 6f = 38 ° , R = 8.7X106 Figure 24.- Concluded.
.004
I
8f, deg 0 0 [] 38
.002 I
\ \
b
-.002 0 5 I0 15 20 e, deg Figure 25.- C16 a for right aileron versus angle of attack (measured at 6aR = 0°); R = 8.7X 106 .
.O8 .O4 Cy O, deg 0 0
-%
[] 8.4 [ O 16.4 - .08 On 0 - .02 -.04 .O2 C_ - .02 -.04 -3O -20 -I0 0 I0 20 3O St, deg (a) 6f= 0°,# = 0°, R = 4.1X10 6 Figure 26.- Effect of rudder deflection on directional characteristics.
,16 .12 ,,,, - .08 Cy .04 I - .O4< 0 14.6 -.08 .O4 C n - .02 - .04 -.06 _> .04 < -.----41 .02[ <) C_
T
-,02 I0 20 30 -30 -20 -10 _r,deg (b) 8f= 0°,/3 = -4 °, R = 8.7X 10 6 Figure 26.- Continued.
.20 .16 of
t
.12 f f J Cy f .08 .04 a, deg O 0 O, [] 8.4 0 14.6 -.04 .02 C n -.04 -.06 -.08 .06 .04 C_ .02 _ A ' - 50 - 20 - I0 0 I0 20 3( B r, deg (c)6f= 0°,#=-8 °,R = 8.7X106 Figure 26.- Continued.
.02 -. 04 • , -.06 -.08 -.08 _ f f "--" - ,04 f -.02[ -30 -20 -I0 0 I0 2O 30 8r, deg (d) 6f= 0°,/3 =-12 °, R = 8.7XI06 Figure 26.- Continued.
,32 ....X ) ,28 J_ / ,24 tf f_ .20 _.
a,deg 0 0 .08 _.._ _-_- -- [] 8.4 0 14.5 .04 i -.02 - .04 C n - .06 -.08 -.10 .O8 > <_-.
.06 < C_ .04 3-- _ _,__........_..._ ...-----'---" '- .02 -3O -2O -IO 0 I0 2O 3O B r , deg (e) 6f= 0°,/3 = -16 °, R = 8.7X l06 Figure 26.- Continued. 73 .08 _4 .O4 Cy _, deg © .3 [-1 8.7 O {6.6 C n -.02 Q -.04 .O4 .O2 C_.
-.02 -30 -20 -I0 0 I0 20 30 8r, deg (f) _ f = 40 ° ,/3 = 0° , R = 4. l X 106 Figure 26.- Continued.
7/, .12 ._t J .08 :, deg 0 .3 [] 8.7 12.8 -.OE .04 .02 Cn ___. _.
-.02 -.04 -.06 .O6 .04 ........ , _......._>-- - ....
C7. _ ..---() .02 0 _ -:50 -20 -I0 0 I0 20 50 _r, deg (g) 6f = 38°,/3 = -4° , R = 8.7X106 Figure 26.- Continued.
• 16 _I J .12 J J 1f _ a,deg 0 .3
o
[] 8.7
0 t2.8
-.04 C n - .02 -.04 -.06 .O6
.04 _
C_ _ __"- --'- "- .02[-] _)-_-
O3o
- 20 -I0 0 I0 2O 3O Br,deg (h) 6f = 38 °,/3 = -8 °, R = 8.7X 106 Figure 26.- Continued.
.20 .16 _- .._ / _ .12 Cy .08 a, deg 0 .3 .04 I-i 8.7 < 0 12.8 .O2 - .02 Cn -.04 -.06 -.08 -30 -20 -I0 0 I0 20 30 8r,deg (i) 6f= 38 ° ,/3 =-12 ° , R = 8.7X106 Figure 26.- Continued.
.28 i .24 i ...A :] f l / i it I fl ! f .20; f I f 1t f Cy .i6 I t- a,deg 0 0 .08
.o4 >'7
-.02 - .04 C n -.06 -.08 -.10 .08 < .06 CZ .04 I .02( -30 8r,deg (j) 6f = 38 ° ,/7 =-16 ° , R = 8.7XI06 Figure 26.- Concluded.
,_ deg I I Sf, deg _f,deg __ 40 -.0005 Rnx I0 -_'- Rn x 10-6 4.1 4.1 I I -.0010 ,,,O
_:,.-_.._..._ J
-.0015 Cnsr I
I
_f, deg _f, deg 0 __ __ :58 __ m -.0005 Rn x 10 -6 Rn x 10-6 8.7 8.7 -.0010
7-------o--'-o
_.0015'_"" 5 I0 15 20 0 5 I0 15 2O u,deg (:z,deg (a)/3 = 0 ° and -4 ° Figure 27.- Cn6r versus angle of attack for flaps up and down; R = 4.1X l 06 and 8.7X 106 .
/3,deg -8
I I
_f, deg Sf, deg 0 38 - .0005 Rn x I0 -6 Rnx 10 -6 8.7 8.7 • I - .0010 .----.0-- f "--'-0"-
"-o
-.0015 Cn_r /9_ deg -12
" I
I
_f, deg Sf, deg 0 __ 38 -- -.0005 Rn x I0 -6 Rn x 10 .6 8.7 8.7 I -.0010
_.._.._ .--...o-.-..._
:2...__......-.-.o _
-.0015 _ 0 5 I0 15 2O 0 5 I0 15 20 a,deg a,deg (b)/3 = -8 ° and - 12 ° Figure 27.- Concluded.
o0 _D X p.., eJ _ II ;.2¸ "" 0
/ -
I
/
I
0 ..9 (k.l "_ ¢) I I I !
/
i I ('-,1
T -
- g
I .,.._
%
q !
I" c ¢.)
¢.)
J cO o II X _6 lJ o i o6 ¢-q o L_ o o I _f 0 0 0 o.
q o q o I" I" I !
C .16 .12 a, deg [] 2 O8 0 6 ZX 0 Cy 0 -2 .O4 -.04 .O4 C n -.04 .O4 C_ -.04 6 -12 - 8 - 4 0 4 8 2 #, deg Figure 29.- Lateral characteristics with landing gear down; R = 8.7× 106 , 8 f = 38 °, .003 .002 Rn xlO -6 0 8.7 Cn B r-I 4.1 .001 o -.001 -.002 _ c_e - .003 -.004 0 4 8 12 16 a _ deg Figure 30.- Stability derivatives Cnt 3 and C//3 versus angle of attack; 6f = 0 °.
\\ ?-
- \
X N c- X II O0 -o m: m= o l i f_ c 0 Q- L_ Q_ Q.m oJ O O LL LL L_ _f li
_dd
Sap ' • ' alSUO qso_uMoo .t2 8f, deg
/
0 0 .10 [] 2O
/
.O8 Ch o .06 .04 .,,d .02 0 4 8 12 16 a,deo (a) Right aileron hinge-moment coefficient versus angle of attack for three flap angles, 8a R = 0 °.
l , Figure 32.- Hinge-moment coefficients; R = 8.7X ! 06 .
K_ I: !
L
I _= m i = .10
I
a, deg 0 0 8f,deg [] 8.4 O 14.6 Cha .05 [ (,'_ .15 e, deg 0 .4 8f, deg F! 8.7 38 O 2.9 .10 Cho .05 ] - .05 15 10 5 0 -5 -I0 -15 8a,deg (b) Right aileron hinge-moment coefficient versus aileron angle for $f= 0 ° (top) and 6i"= 38 ° (bottom).
Figure 32.- Continued. 87 OJ II !
o_ .e o E E _b o o c_ i_r) l 0,,I o o o.
8..
I I I g .10 .05 _'_ 8f ,deg 0 ,',, d _)g Che , left 0 14.6 \ N l - .05 -.I0 .15 ) 8f,deg Che , right -t0 -5 0 5 I0 15 Be , deg (d) Elevator hinge-moment coefficient versus elevator angle for 6f = 0 ° (top) and 5f = 38 ° (bottom).
Figure 32.- Continued.
.3O ,,, deg 0 0 [] B •20, .I0 L Ch r -:10 -. 20 -.50 -30 -20 -I0 0 I0 20 30 8r, deg (e) Rudder hinge-moment coefficient versus rudder angle at two angles of attack; 6f = 0°.
Figure 32.- Continued.
o O II tO c_ ¢.D "3 .,=-i a) -0 o _.
- I".- O "" I O °_ O<> .d c_ ,===_ el) O o "_ i c,i a0 e,D ,,,=_ [.r., o O ,b
/
o CD I • ,,.-.t ..o o tO o o o OI o O eq .-.; Od I ,.el ¢..
NASA-Langley, 19'71 -- 1 A-3135 95 = l- oJ O o I ("4 O_ o °_._ o & x_ OOOD - OJl_O ..= q..
_o ODO O O N
r o
oJ O N O O o.
I I I r (.)
.2O ,B, deg 8f, deg 0 -8 0 [] -8 58 0 -16 58 A -16 0 .I ]hr -.10 \ \\\ -.20 \ -'3930 -20 -I0 0 I0 20 30 8 r , deg (0 Rudder hinge-moment coettlcient versus rudder angle at two sideslip angles and two flap angles; c_ = 0 °.
Figure 32.- Continued.
NATIONAL AERONAUTICS AND SPACE ADMISTRATION WASHINGTON, D.C. 20546 POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND OFFICIAL BUSINESS SPACE ADMINISTRATION FIRST CLASS MAIL PENALTY FOR PRIVATE USE $300 If Undeliverable ( Section 158 POSTMASTER : Postal Manual) Do Not Return "The aeronautical and space activOies o[ the United States shaft be conducted so as to contribute . . . to the expansion o[ human knowl- edge o[ phenomena in the atmosphere and space. The Administration shall provide/or the widest practicable and appropriate dissemination o[ inJormation concerning its acth,ities and the results thereo[."
--NATIONAL AERONAUTICS AND SPACE ACT OF 1958
NASA SCIENTIFIC AND TECHNICAL PUBLICATIONS
TECHNICAL REPORTS: Scientific and TECHNICAL TRANSLATIONS: Information technical information considered important, published in a foreign language considered complete, and a lasting contribution to existing to merit NASA distribution in English.
knowledge.
SPECIAL PUBLICATIONS: Information TECHNICAL NOTES: Information less broad derived from or of value to NASA activities.
in scope but nevertheless of importance as a Publications include conference proceedings, contribution to existing knowledge.
monographs, data compilations, handbooks, sourcebooks, and special bibliographies.
TECHNICAL MEMORANDUMS: Information receiving limited distribution TECHNOLOGY UTILIZATION because Of preliminary data, security classifica- PUBLICATIONS: Information on technology tion, or other reasons.
used by NASA that may be of particular interest in commercial and other non-aerospace CONTRACTOR REPORTS: Scientific and applications. Publications include Tech Briefs, technical information generated under a NASA Technology Utilization Reports and contract or grant and considered an important Technology Surveys.
contribution to existing knowledge.
Details on the availability of these publications may be obtained from: SCIENTIFIC AND TECHNICAL INFORMATION OFFICE NATIONAL AERONAUTICS AND SPACE ADMINISTRATION Washington, D.C. 20546