Wind-tunnel investigation of an stol aircraft ...€¦ · TECH LIBRARY KAFB, NM 0132298 WIND-TUNNEL...

WIND-TUNNEL INVESTIGATION OF A N STOL AIRCRAFT CONFIGURATION EQUIPPED WITH A N EXTERNAL-FLOW JET FLAP

by Lysle P. Purlett and Jumes P. Shivers

Langley Research Center Langley Station, Humpton, Va.

N A T I O N A L AERONAUTICS A N D SPACE 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 9

https://ntrs.nasa.gov/search.jsp?R=19690023897 2020-07-10T00:47:50+00:00Z

TECH LIBRARY KAFB, NM

.--. 0132298

WIND-TUNNEL INVESTIGATION OF AN STOL AIRCRAFT CONFIGURATION

EQUIPPED WITH AN EXTERNAL-FLOW J E T FLAP

By Lysle P. Parlett and James P. Shivers

Langley Research Center Langley Station, Hampton, Va.

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - CFSTI price $3.00

WIND-TUNNEL INVESTIGATION O F AN STOL AIRCRAFT CONFIGURATION

EQTJIPPEDWITHANEXTERNAL-FLOWJET FLAP

By Lysle P. Parlett and James P. Shivers Langley Research Center

SUMMARY

The present investigation was performed to provide information on the static longitudinal and lateral characteristics of a proposed short take-off and landing (STOL) t rans- port configuration utilizing the jet-flap principle. Longitudinal tests were conducted at engine gross-thrust coefficients of from 0 to 3.4 through a range of angle of attack which included the stall; and lateral tes ts were made, both power-off and power-on, through a sideslip range of +30° at angles of attack of 0' and loo.

Untrimmed lift coefficients up to 7.8 were attained at a gross-thrust coefficient of 2.83 in the tail-off condition. With the tail on, nearly all high-lift conditions were charac- terized by a marked longitudinal instability (or pitch-up tendency) which began at an angle of attack of 7'. The instability was apparently caused by the tip vortices which, under the influence of the highly loaded center section of the wing, were drawn into the region of the horizontal tail. The tail-on configuration was directionally stable and had positive dihedral effect at all flap and power settings tested; and in the take-off and landing conditions increasing power increased directional stability and decreased dihedral effect. With one outboard engine not operating, the model could be trimmed laterally and directionally up to lift coefficients of 4.2 in the take-off condition and 5.7 in the landing condition. Above these lift coefficients the model could not be tr immed in roll, but t r im in yaw could still be attained.

INTRODUCTION

The external-flow jet-flap principle is incorporated in a recently proposed design for a medium-size four-engine jet transport intended to have short take-off and landing (STOL) capabilities. Previous investigations (refs. 1, 2, and 3) have demonstrated that an external-flow jet flap can produce the high lift coefficients required for short-field operation, but that the high lift coefficients may be accompanied by serious t r im and stability problems. and unsymmetrical span loading of powered lift which would vary with configuration. order to broaden the knowledge in the jet-flap field by testing a configuration significantly

These problems are attributed primarily to downwash characteristics In

I

different from those of past investigations, the NASA undertook to test a model of the proposed STOL transport. The wing of this configuration is more highly tapered and the engines a r e located relatively closer to the fuselage than in the previous investigations. The tes ts provided general aerodynamic data for the take-off, cruise, and landing conditions, with emphasis on t r im and stability studies in the high-power, high-lift conditions. Longitudinal and lateral forces and moments were measured at angles of attack up to 28O, at sideslip angles up to 30°, and at engine gross-thrust coefficients up to 3.4. In t e rms of t r im flight conditions fo r the proposed full-scale aircraft represented by the model, a gross-thrust coefficient of 3.4 would result in a thrust-weight ratio of approximately 0.6.

SYMBOLS

A sketch of the axis system used in the investigation is presented in figure 1. Lon- gitudinal forces and moments a r e referred to the wind-axis system; lateral and directional forces and moments a r e referred to the body-axis system.

b wing span, f t (m)

CD drag coefficient, D/qS

CL lift coefficient, L/qS

CL,O lift coefficient, power off

CL, r jet -induced circulation lift coefficient

cz rolling-moment coefficient, Mx/qSb

Cm pitching - moment coefficient , My/qSE

Cn yawing-moment coefficient, MZ/qSb

CY side-force coefficient, Fy/qS

aCY CYp = ap

C

2

engine gross-thrust coefficient, mV,/qS

local wing chord, ft (m)

- C

D

FA

FN

FY

it

L

MX

MY

MZ

m

S

T

VE

(Y

P

6aux

6f

length of mean aerodynamic chord, ft (m)

drag, 1b (N)

axial force, lb (N)

normal force, lb (N)

side force, lb (N)

incidence of horizontal tail, deg

lift, lb (N)

rolling moment, ft-lb (N-m)

pitching moment (referred to 0.25E), ft-lb (N-m)

yawing moment, ft-lb (N-m)

engine mass flow rate, slugs/sec (kg/sec)

free-stream dynamic pressure, lb/ft2 (N/m2)

wing area, f t2 (m2)

thrust, lb (N)

engine exit velocity, ft/sec (m/sec)

model body axes

angle of attack, deg

angle of sideslip, deg

deflection of auxiliary flap, deg

flap deflection, deg

jet deflection, deg

3

rudder deflection, deg

spoiler deflection, deg

vane deflection, deg

downwash angle, deg

JQTg$ T flap turning efficiency,

Designations for flap settings a r e given in figure 2(b).

MODEL AND APPARATUS

The investigation was conducted on the four-engine high-wing jet transport model illustrated by the three-view drawing of figure 2(a). A typical section through the flap system and the relationship of the flaps to the engines a re shown in figure 2(b). The leading-edge slat shown in figure 2(b) was extended for all test conditions. The flap com- binations a r e defined in the table of figure 2(b) and a plan view of the wing semispan is presented in figure 2(c). Photographs of the model a r e presented in figure 3, and dimen- sional characteristics a r e listed in table I.

The engines were of the ejector type (in which thrust resulting from gas flow through primary nozzles is augmented by a secondary flow of ambient air induced by the primary flow) and had the same external geometry as a current turbofan engine. Flow of the primary gas (compressed nitrogen) to the section of the engine simulating the turbine was controlled independently of primary flow to the fan simulator so that the desired thrust was obtained at the desired bypass ratio (8 to 1). For some of the tests, thrust deflector plates were installed on the outboard fan simulators as shown in figure 2(b).

It may be noted that the use of these ejector engines did not allow inlet and exit mass flow rates to be simulated correctly at the same time, but for the present tes ts the exit mass flow was considered to be the more important of the two.

The model was mounted on a six-component strain-gage balance and was strut- supported in the test section of the Langley full-scale tunnel. throat test section of 30 by 60 feet (9.14 by 18.29 meters), which allows models of the present s ize (8-foot (2.5-m) span) to be tested at high lift coefficients without introducing significant tunnel wall effects.

This tunnel has an open-

4

TESTS AND PROCEDURES

In preparation for the present tests, single-engine calibrations were made to deter- mine net thrust and mass flow ra tes as functions of nitrogen drive pressure in the static condition and at the test free-s t ream airspeed at zero angle of attack. The tests were then run by setting the drive pressures for the fan and turbine simulators, respectively, to the desired values and holding these pressures constant through the ranges of angle of attack or sideslip.

J e t deflection angles and flap turning efficiencies were determined from measure- ments of the normal and axial forces made in the static thrust condition with flaps deflected. The static thrust used in computing turning efficiency w a s taken directly from the single-engine calibrations at the appropriate drive pressures.

During the wind-on tes t s various changes were made to the flap geometry or to control-surface deflections; each condition was usually tested at values of 3.4 through a range of angle of attack of -4O to 28'. All tail-off tes ts were made with both the horizontal and vertical tails removed. Sideslip runs were made over a range of angles of sideslip from -30° to 30'. dynamic pressure of 11 lb/ft2 (527 N/m2), which corresponds to a velocity of 97 ft/sec (29.6 m/sec). aerodynamic chord of the wing.

Cp of 0 to

Nearly all wind-on tests were made at a free-stream

The Reynolds number was approximately 0.8 X lo6 based on the mean

No wind-tunnel jet boundary corrections were applied to the data because such corrections were computed for a somewhat larger high-lift model during a previous investigation (ref. 3) and were found to be negligible.

PRESENTATION OF DATA

The test data a r e presented in the following figures. The four main headings corre-

Figure

Longitudinal characteristics, tail off . . . . . . . . . . . . . . . . . . . . . . . . 4-7

Longitudinal basic data, tail off, auxiliary flaps deflected 8

Summary of auxiliary-flap performance 9 Analysis of jet-flap effectiveness . . . . . . . . . . . . . . . . . . . . . . . . . . 10-12

spond to those in the Discussion section.

Lift Characteristics

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Longitudinal Stability and Trim, Symmetric Thrust Longitudinal characteristics, tail on, basic configurations . . . . . . . . . . . . 13 - 16 Photographs of smoke flow showing vortex system 17 . . . . . . . . . . . . . . . .

5

Figure Longitudinal characteristics, tail on, configurations intended

18-23 Horizontal -tail effectiveness . . . . . . . . . . . . . . . . . . . . . . . . . . . 24

to remedy longitudinal instability . . . . . . . . . . . . . . . . . . . . . . . .

Lateral Characteristics, Symmetric Thrust Lateral characteristics, tail off . . . . . . . . . . . . . . . . . . . . . . . . . . 25-27 Lateral characteristics, tail on, basic configurations . . . . . . . . . . . . . . 28-31 Lateral characteristics, tail on, controls deflected . . . . . . . . . . . . . . . 32-35

Lateral and Longitudinal Characteristics, Asymmetric Thrust Lateral characteristics, tail on, asymmetric thrust and control . . . . . . . . . 36-38 Lateral t r im capability, one engine out 39 Longitudinal characteristics, asymmetric thrust and control . . . . . . . . . . 40-45

. . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . Effect of thrust distribution on lift 46

DISCUSSION

Lift Characteristics

Basic longitudinal data for the model in the tail-off condition at flap deflections representing the cruise, take-off (two deflections), and landing configurations a r e presented in figures 4 to 7. The data show that the stall angle and maximum lift coefficient increased with increasing thrust coefficient, and that as flap deflection increased, the effects of thrust on the lift characteristics became more pronounced. lift coefficients up to 7.8 (untrimmed) at a gross-thrust coefficient.of 2.83. As would be expected, high lift coefficients a r e accompanied by large nose-down moments because of the rearward location of the flap loads.

The leading-edge slat was extended for all test conditions.

The landing flap deflection (fig. 7) produced

With the basic landing flap setting LDG, which produced the highest lift coefficients, auxiliary flaps were investigated as a means of providing glidepath control during a landing approach. Data which show the longitudinal characteristics with various auxiliary- flap deflections in the approach condition a r e presented in figure 8 and a re summarized in figure 9. thrust can produce increases in drag. These drag increases reflect the large induced drag which accompanies the induced l i f t at high flap settings in a high-lift system, and suggest that the auxiliary flap might be an impractical device for glidepath control, at least, in the usual sense, with large main-flap deflections. If the flap deflection fo r landing were lower, i t is possible that the auxiliary flap would appear in a more favorable light as a glidepath control system.

Figure 9 shows that with the basic landing flap setting LDG, increases in

The effectiveness of a jet-flap system is usually analyzed in t e rms of C L , ~ , the jet-induced circulation lift coefficient. The quantity CL, r is significant because it

6

represents a lift component not solely attributable either to the upward component of the deflected engine thrust or to the power-off lift of the wing, and is therefore an indication of the ability of the integrated engine-wing-flap system to utilize engine power to produce additional increments of lift coefficient. A typical resolution of total lift coefficient into its three components is shown for a 60° flap setting in figure 10. represents the circulation lift normally developed by the wing and flap system in a moving airs t ream in the power-off condition. In the powered condition, the engine slipstream impinges on the flap system and is thereby deflected downward through the angle the te rm C p sin 6. + ar] represents the lift contribution due to this redirection of engine gross thrust. The flow of the engine slipstream through the flap system and downward from the trailing edge as a jet sheet not only produces the force represented by C p sin 6. + a! 17, but also induces a flow which augments the circulation over the wing. This increased circulation gives r i s e to the third l i f t component, the jet-induced added circulation lift CL, r.

C L , ~ as the basis for comparison, the effectiveness of the engine-wing-flap system of the present model is compared to that of the model of reference 2 in figure 11. The comparison is not exact because the data for reference 2 a re for a jet deflection angle of 60°, whereas the most nearly comparable jet deflection in the present investigation was 65'. comparison. The C L , r values produced by the model of reference 2 do not neces- sar i ly represent the ideal, but they have been considered generally representative of those to be expected from an efficient external-flow jet-flap system. jet-induced circulation l i f t for the present model compares unfavorably with that of the model of reference 2 throughout the range of for which both models were tested. Analysis of the probable effects of geometric differences between the models seems to indicate that it is important to have the engine efflux flattened and spread more widely across the span than is the case for the present model.

The CL at C b = 0

6j; and

( J )

( 1 )

With

It is believed, however, that this slight discrepancy does not materially affect the

Figure 11 shows that the

C p

Because the jet-induced lift is highly dependent on the direction and velocity of the engine slipstream as it leaves the flap system, it appears that for best jet-flap performance the flap system should be capable of turning the slipstream efficiently through large angles. The slipstream angle Sj and the static turning efficiency r] for the present model a r e shown in figure 12, which is a plot of the ratio of normal force to thrust FN/T against the ratio of net axial force to thrust FA/T. angles near 600 the turning efficiency was approximately 0.7, which is low enough to be at least partially responsible for the relatively low values of CLJ.

Figure 12 shows that at turning

Longitudinal Stability and Trim, Symmetric Thrust

The longitudinal stability and t r im characteristics of the model with the tail on a r e plotted against angle of attack for various thrust levels and flap settings in figures 13

7

to 16. Some of the tests on which this group of figures is based were performed with the landing-gear pod on, and some with the pod off; figure 13 is presented f i rs t to establish that the pod had virtually no effect on longitudinal characteristics? and that data for other configuration variations may therefore be compared regardless of the presence or absence of the pod.

Longitudinal characteristics for the cruise condition (although with leading-edge slat extended) are presented in figure 14. The most noteworthy characteristic shown by figure 14 is the longitudinal instability which develops under power-on conditions at an angle of attack of approximately 7O. landing flap settings? respectively, also show instability developing at an angle of attack of 7O, and this instability becomes progressively worse with increasing angle of attack until near the stall the model was more unstable than it was with the tail off. This destabilizing effect of the tail indicates that the tail was operating in a downwash field which increased in intensity at a rate greater than that at which the angle of attack increased? with the result that the downwash factor 1 - - was negative. A brief smoke-flow study (sample

photographs a r e presented in fig. 17) showed that the tail was indeed immersed in a downwash field that for the present model may have been particularly powerful because of the high concentration of lift on the inboard sections of the wing, which caused a large and powerful tip vortex to be located far inboard in the region of the horizontal tail. The model of reference 2 did not have such a pronounced instability as the present model; therefore? attempts were made to achieve a more nearly uniform spanwise lift distribution by means of thrust deflection and flap changes. tests made with the thrust deflectors installed on the two outboard engines. The data of figure 18 show that the deflectors produce some slight increment in l if t but that they do not improve the stability. Apparently the outboard engines a r e located so far inboard that increasing the spread of their efflux somewhat does not alter the spanwise lift distribution (or the downwash) sufficiently to relieve the vortex in the region of the horizontal tail.

Figures 15 and 16, presenting data for the take-off and

de da!

Figure 18 presents the results of

The results of an additional modification? that of drooping the ailerons 40' to increase the lift on the outer part of the wing, a r e presented in figure 19. These data show that the combination of aileron droop and thrust deflection produce a noticeable? although insufficient, contribution toward stability.

Although the use of spoiler deflection is not normally associated with a take-off configuration, the effects of symmetrically deflecting the inboard section of the spoiler were investigated in the present case as a possible means of improving longitudinal stability by reducing lift on the inboard part of the wing, thereby making the lift distribution more uniform. had negligible effects on lift and stability.

Figure 20 shows, however, that inboard-spoiler deflections of loo and 60'

8

Early in the present test program the flap deflection AT02 (see fig. 2(b)) was investigated as an alternate take-off flap setting. arrangement was discarded in favor of the take-off (TO) flap arrangement which has been the subject of the foregoing take-off discussion; figure 21 presents the AT02 data, however, primarily as further evidence that inboard concentration of lift is the source of the present longitudinal .stability problem. Figure 21(b) shows that the AT02 flap deflection has a slightly higher maximum lift, but, at high lift, slightly more instability than the TO flap setting.

On the basis of tes t results, the AT02

In the landing flap configuration, as in the take-off configuration, symmetric con- t rol deflection proved ineffective in relieving the longitudinal instability. 23 show that in the landing configuration the longitudinal stability characteristics were virtually unchanged either by 40° of aileron droop or by 60' of symmetric spoiler deflection.

Figures 22 and

If it is assumed that the downwash caused by inboard concentration of lift is a major factor in the longitudinal instability, speculation may then be made on various means by which a more favorable downwash distribution might be obtained. One possible means would be to move the engines outboard. The engines a r e presently located fairly close inboard to minimize lateral out-of-trim moments in the engine-out condition; it may be, however, that in their present location their adverse effect on longitudinal stability would outweigh the lateral considerations. Another possible means of improving lift distribution might be to reduce the taper ratio. The present wing is highly tapered, with all of the power applied to inboard, long-chord areas; therefore, reducing the length of inboard chords while lengthening the outboard chords would probably spread lift outboard somewhat. Another possibility for making the spanwise lift distribution more uniform is the use of wing sweep, since sweep has the effect of inducing outward spanwise flow. If the engine slipstream is thereby induced to spread outboard, it is possible that the jet-flap effect would be extended to outboard a reas which a r e not now developing high lift. In configurations such as the present one in which the engines a r e located fairly close inboard, sweeping the wing also has the advantage of causing a reduction in t r im requirements by locating the flap load closer to the aerodynamic center. longitudinal stability might be to relocate the horizontal tail of the configuration. Smoke flow studies (fig. 17) showed that after leaving the wing tip, the tip vortices move toward the airplane center line as they move rearward, passing over the outboard a reas of the horizontal tail. If the tail were moved forward, these vortices might pass far enough outboard of the horizontal tail to avoid the present downwash effect. Such a result is suggested by the fact that the model of reference 2, which had a shorter tail length and lower tail height, had much better longitudinal stability than the present model.

Another possibility for improving

9

Figure 24 presents the effect of varying the incidence of the horizontal tail, and shows that the tail, which has an area of 0.37 of the wing area, is capable of trimming the large nose-down moments produced by the wing and flap at high-lift conditions.

Lateral Characteristics, Symmetric Thrust

Tail off.- The tail-off static lateral and directional characterist ics of the model a r e presented in figures 25 to 27 for the cruise, take-off, and landing configurations at three thrust levels and at angles of attack of 0' and 10'. As might be expected, the model with the tail off is directionally unstable in all flight conditions, and it is to be noted that the instability increases with increasing thrust. The data of figures 26 and 27 also show that in the take-off and landing configurations the dihedral effect -Ci negative as thrust is applied.

goes from positive to ( d Tail on.- Figure 28 presents lateral and directional data for the tail-on configuration

in the cruise condition. In this condition the model is laterally and directionally stable, and the stability characteristics a r e not noticeably affected by changes in thrust.

The lateral and directional characteristics of the configuration with flaps deflected to take-off and landing settings a r e presented in figures 29 to 31. The data show that the model is laterally and directionally stable in the power-off condition, but that in some cases power effects a r e pronounced. The application of power to the basic take-off configuration (fig. 29) produces a marked increase in directional stability, which contrasts with the destabilizing effect it produced in the tail-off condition. In the basic take-off and in the landing configurations, increasing power caused reduction in the dihedral effect. Theory and experience would lead to the expectation that increased. power (with consequently increased lift) would produce increased dihedral effect; the reduction in the present case may have been due to asymmetric exposure of the horizontal tail to the wing tip vortices in sideslip conditions. With the landing flap setting LDG, but with inboard spoilers deflected 60° (fig. 32), the model was also laterally and directionally stable.

The control moments produced by asymmetric deflection of various control surfaces a r e presented in figures 33 to 35. Deflection of the inboard spoiler I O o (fig. 33(a), flap setting TO) produces virtually no moments; deflection of the outboard spoiler 30' (fig. 34(b), flap setting LDG) produces large rolling moments accompanied by small favorable yawing moments. Rudder effectiveness (fig. 35, flap setting TO) is, as would be expected, unaffected by engine thrust.

Lateral and Longitudinal Characteristics, Asymmetric Thrust

Under conditions of asymmetric thrust, the lateral characteristics of a configuration are usually the primary concern and will consequently be discussed in this section prior to the longitudinal characteristics.

10

Lateral and directional characterist ics of the basic take-off configuration for the condition of left outboard engine not operating are plotted against angle of attack and against angle of sideslip in figures 36(a) and 36(b), respectively. As would be expected of a powered-lift configuration, out-of-trim yawing and rolling moments a r e large and increase with increasing thrust or angle of attack. show that the configuration is laterally and directionally stable with one outboard engine out; the change in slope of the Cz curve at p = 150 may be further evidence that the wing tip vortex acting on the horizontal tail has a noticeable influence on dihedral effect. Figure 37 presents lateral and directional characteristics for the basic take-off configuration with left outboard engine not operating and with the ailerons, spoiler, and rudder deflected in the direction to res tore tr im. Figure 37(a) shows that t r im in yaw was produced throughout the angle-of-attack range, whereas t r im in roll was not achieved at angles of attack greater than 13' (CL = 4.2). It should be noted that these data are for a spoiler deflection of 30°; better roll t r im capability could reasonably be expected if full spoiler deflection (60O) had been utilized.

The slopes of the curves in figure 36(b)

The preceding discussion of the ability of the control surfaces to res tore lateral t r im after loss of thrust from one outboard engine was for the take-off configuration; for the landing configuration, an analysis of roll t r im capability with an engine out is presented in figures 38 and 39. The engine-out curve of figure 39 (landing configuration, flap setting LDG) is plotted from the basic data of figure 38 and represents the rolling moments and lift coefficients which would exist if, after loss of all thrust from the left outboard engine (assuming all four engines were initially operating at a Cp of 0.50 per engine), the Cp of the remaining left-hand engine were increased to 0.71. The spoiler-deflected curve (plotted with sign reversed for comparison) is based on rolling-moment data presented in figure 34(b). These control power data were obtained under conditions of symmetric thrust, but they a re considered to be applicable to the engine-out condition because in engine-out operation the spoiler would be deflected on that wing on which two engines were still operating. Lift coefficients for the spoiler-deflected curve a r e those of the engine-out curve decreased at each angle of attack by the amount resulting from spoiler deflection. Figure 39 shows that, under this engine-out condition, t r im in roll could be maintained at lift coefficients up to 5.7. Capability for t r im and maneuver at somewhat higher lift coefficients would be expected i f more than 30° spoiler deflection were used and if ailerons were also employed.

The longitudinal characterist ics of the configuration under the conditions of lateral asymmetry which have been discussed a r e presented in figures 40 to 45. Loss of one outboard engine in either the take-off or landing configurations (figs. 40 and 41, respectively) results in markedly worse longitudinal stability characteristics in addition to the expected loss in maximum lift. Spoiler deflection (figs. 42, 43, and 45) produces a slight loss of lift and, in cases where lift is spoiled over outboard areas (figs. 44 and 45), increases

11

I I I I I I I I I 1,1111, 1.11.11.1 I. .

longitudinal instability. Deflection of all lateral control surfaces (fig. 44) produces no particular longitudinal effects other than those which the preceding discussion has attributed to the spoiler. dent of the thrust distribution. out condition as compared with a symmetric thrust condition, provided the same total is maintained by increasing the thrust of the remaining operable engines. This characterist ic has been observed in connection with other jet-flap configurations (ref. 3) and would certainly be 'expected of the present configuration, in which the outboard engines are very close to the inboard engines.

Figure 46 shows that fo r a given total Cp, the lift is indepen- This fact means that there is no loss in lift for an engine-

C p

SUMMARY OF RESULTS

A wind-tunnel investigation of the aerodynamic characterist ics of a transport model with a tapered wing equipped with an external-flow jet flap has yielded the following results :

1. In the power-on condition the configuration had a marked longitudinal instability, This instability became more severe or pitch-up tendency, at angles of attack above 7'.

as angle of attack o r thrust was increased.

2. The instability was apparently caused by the wing tip vortices being drawn into the region of the horizontal tail by the high concentration of lift over the inboard a reas of the wing. Configuration or thrust changes which increased inboard lift concentration (outboard-spoiler deflection, outboard engine inoperative) caused increased instability. Changes intended to produce more nearly uniform spanwise lift distribution (drooped ailerons, thrust deflectors on outboard engines, and inboard-spoiler deflection) produced only negligible improvements in stability.

3. In the powered-lift conditions, loss of thrust of one outboard engine produced large rolling and yawing moments. Deflection of lateral controls could maintain t r im at lift coefficients up to 4.2 in the take-off configuration and 5.7 in the landing configuration.

4. The model with tail on was laterally and directionally stable under all power-on test conditions. Increases in thrust produced decreases in dihedral effect and increases in directional stability.

5. The static turning efficiency and the jet-induced circulation lift which it produced were low compared with those of previous investigations.

Langley Research Center, National Aeronautics and Space Administration,

Langley Station, Hampton, Va., April 30, 1969, 721-01-00-31-23.

12

REFERENCES

1. Lowry, John G.; Riebe, John M.; and Campbell, John P.: The Jet-Augmented Flap. Preprint No. 715, S.M.F. Fund Paper, Inst. Aeronaut. Sci., Jan. 1957.

2. Campbell, John P.; and Johnson, Joseph L., Jr.: Wind-Tunnel Investigation of an External- Flow Jet- Augmented Slotted Flap Suitable for Application to Airplanes With Pod-Mounted Engines. NACA TN 3898, 1956.

3. Pariett, Lysle P.; Fink, Marvin P.; and Freeman, Delma C., Jr. (With appendix B by Marion 0. McKinney and Joseph L. Johnson, Jr.): Wind-Tunnel Investigation of a Large J e t Transport Model Equipped With an External-Flow Jet Flap. NASA TN D-4928, 1968.

13

TABLE 1.- DIMENSIONS OF MODEL

Fuselage: Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.156 f t ( 2.791 m) Maximum height . . . . . . . . . . . . . . . . . . . . . . . . . . 1.060 f t ( 0.323 m) Maximu.m width (excluding landing-gear pods) . . . . . . . . . . 1.00 f t ( 0.305 m) Maximum cross-sectional area (including landing-gear pods) . . 1.588 ft2 (0.148 m2)

Wing: . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Area 9.80 ft2 (0.910 m2)

S p a n . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.28ft ( 2.53m) Root chord. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.91 f t ( 0.582 m) Tip chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.46 f t ( 0.140 m) Mean aerodynamic chord length . . . . . . . . . . . . . . . . . 1.34 ft ( 0.408 m) Spanwise station of mean aerodynamic chord . . . . . . . . . . 1.65 f t ( 0.503 m)

Sweep of quarter-chord line . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8'

Incidence at root . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -0.9' D i h e d r a l . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2'50'

Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.0

Taper ratio 0.242

3'

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

W i n g t w i s t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Trailing-edge flaps:

f 1 Span . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.11 to 0.43 wing semispan Chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.26 local wing chord

Span . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.43 to 0.75 wing semispan Chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.26 local wing chord



f2

Auxiliary flap 1

Auxiliary flap 2

Aileron: Span . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.75 to 1.00 wing semispan Chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.40 local wing chord

Leading-edge slat: Span . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.11 to 1.00 wing semispan Chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.13 local wing chord

14

TABLE 1.- DIMENSIONS O F MODEL - Concluded

Vane: Span . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.11 to 0.75 wing semispan Chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.16 local wing chord

Span . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.18 to 0.43 wing semispan Hinge line . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.69 local wing chord Trailing edge . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.79 local wing chord

Span . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.43 to 0.75 wing semispan Hinge line location . . . . . . . . . . . . . . . . . . . . . . . . 0.69 local wing chord Trailing edge . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.79 local wing chord

Area. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.65 f t2 ( 0.339 m2) Span . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.27 f t ( 1.30 m)

Tip chord (theoretical) . . . . . . . . . . . . . . . . . . . . . . . 0.44 f t ( 0.134 m) Tail length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.65 f t ( 1.72 m) Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.0 Sweep of quarter-chord line . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16.1'

Area (aft of hinge line) . . . . . . . . . . . . . . . . . . . . . . . 0.86 f t2 (0.0798 m2)

Inboard spoiler:

Outboard spoiler:

Horizontal tail:

Root chord (theoretical at fuselage centerline) . . . . . . . . . . 1.27 f t ( 0.387 m)

Elevator:

Hinge line location . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.75 local chord

Area . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.24 ft2 ( 0.208 m2) Span . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.31ft ( 0.399m) Root chord (top of fuselage) . . . . . . . . . . . . . . . . . . . . 2.04 f t ( 0.622 m) Tip chord (at horizontal tail) . . . . . . . . . . . . . . . . . . . . 1.37 f t ( 0.418 m) Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.77 Sweep of quarter-chord line . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45.5O Tail length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.53 f t ( 1.38 m)

Area (aft of hinge line) . . . . . . . . . . . . . . . . . . . . . . . 0.62 ft2 (0.0576 m2) Span . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . l . l O f t ( 0.335m) Hinge line location . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.68 local chord

Vertical tail:

Rudder:

15

? I X - .,-

a

Wind direction -q

I z i

Wind direction .-

Figure 1.- Axis system used in presentation of data.

16

\ , , i cI4 rwt horizontal tai l

I -

n

E. 15

'"I"'

I

4.095 (IO. 4011 Outboard 4.629 (11.7581 Inboard

. . -- I I

2" 501

(a) Three-view drawing showing principal dimensions i n inches (centimeters).

Figure 2.- Drawings of model.

0.020c

Fan primary nozzles

T

rente plane - __

c- -- .

\ Turbine primary nozzles c --_ -

m h

d - - z u(

- h r u s t deflector plate, approximately 5 in. (12. 7 cml wide

I

Flap configuration Inboar~flaps' f

Symbol Nominal 4 ]Over lap Gap

TO Take-off 3 0 / 11.5 0 9.6 , O

AT02 ' Alternate r. .i. 9 . 6 ' 0 0 1.4 0 1.4

LDG Landing LA10 Landing LA20 Landing 0 1.4 LA30 ' Landing 60 1 0 1.4

CR Cru ise ' 0 - -

outboa: f flaps' [ Aux i l iary flap 1 and 2 Vane -

4 Overlap Gap j baux Overlap Gap 6" Overlap I Gap

60 0 1.4 0 0 0 42.5) 0 2

40 9.6 0 0 0 0 42.5 0 60 0 1.4 0 0 0 42.5 0 ! 60 0 1 1.4 10 2 1 42.5 I 0 ; j 60 , 0 I 1.4 , 30 t 1 42.5 1 0 2 1

0 . . 0 - - - 1 - -

M) o i 1 . 4 o o o 42.51 o 2 I

60 ~ 0 ' 1.4 , 20 1 1 42.5 0

(b) Wing section and flap configuration schedule. Overlaps and gaps i n percent local wing chord. Deflections i n degrees,

Figure 2.- Continued.

b 0.11 2

L ,/' z

\

Aux. 1 b 0.752 7 b 0.4311 2 F u se I age

(c) Wing planform showing f lap segments.

Figure 2.- Concluded.

~------. - -~~

Figure 3.- Photographs of model installed in tunnel. L-69-1395

20

I

15

10

5

0

1

C m o

-1

4

2

CL

0

CD

-2

-4 -10 0 10

a, deg 20 30

Figure 4.- Tail-off longitudinal data for cruise flap setting CR.

-1

21

, , ..-.-.-- ..... . l11111111...11 I Ill 1.11 I

0

1

Cm o

-1

8

6

4

CL

2

0

CD

-2

1 i i

0 10 20 30 1

Figure 5.- Tail-off longitudinal data for take-off flap setting TO.

0 Cm

-1

22

10

- L 5 D

0

0

Cm -1

-2

8

6

CL

4

2

cD 0

-2 -10 0 10 20 30

0 1.99 2.83 3 .41

0 -1

Cm

-2

Figure 6.- Tail-off longitudinal data for an alternate take-off flap setting ATOI.

23

0

1

0

Cm -1

-2

-3

6

CL

4

2

CD

0

-2 -10 0 10 20 30 1

3 0 cl 1.95 A 2.83

Figure 7.- Tail-off longitudinal data for landing flap setting LDG.

0 -1 Cm

-2 -3

lo

0

1

0

Crn -1

-2

-3

10

8

CL

6

4

2

CD

0 -10 0 10 20 30 1

(a) Auxi l iary flaps deflected 100; LA10.

0 0 A

CCI

0 1.99 2.83

0 -1 -2 -3

Figure 8.- Longitudinal data for auxi l iary flaps as glidepath control. Tail off.

25

10

0

t i 1

0 17 A h

0 1.99 2.83 3.41

0

t i ! !

Cm -1

-2

i

-3 I I 1 I

I 1 t

t

i f

i

I

I

! !

10 t

8

CL

6

4

2

CD

0 -10 0 10 20

a, deg 30 1 0 -1

Cm -2 -3

(b) Auxiliary flaps deflected ZOO; LAB.


26

L D -

10

5

0

1

0

Cm -1

-2

-3

10

8

6

CL

4

2

CD

0 -10

r 0 20 30 1

(c) Auxiliary flaps deflected 30°; LA30.


0

A

0 1.99 2.83 1

0 -1 Cm

-2 -3

27

4'

cL

c D

8

6

4

2

0

2

0

7

f

1 I

I I 1

I 1

1

t .t 1

t' 1

1 1 1 1 t i. i t L 1

t t

I

1

P C

6 a;x 10 1 30 *O I

2 3

Figure 9.- Summary of auxil iary-flap performance. Landing flap settings; a = 7.50.

28

7

6

5

4

3

2

1

0 1 2

4

3 CP

Figure 10.- Resolution of jet-flap lift in to components. Tail off; bf = 60°; a = loo; LDG.

29

%, r

Figure 11.- Comparison of jet c i rculat ion lift produced by model of present investigation with that of reference 2. Landing flap setting LDG.

30

.8

.6

FN T

. 4

.2

6f = 60°

.2

Line of constant 77 - - - Line of constant 6

j

/6j = 600

] q = .7

.4

.8 ~

q = 1.01

.6 1.0

Figure 12.- Summary of turning efficiency and turning angle.

31

5 L D -

0

2

1

Cm

0

-1

8

6

CL

4

2

c D

0

-2 -10

3

T I i 1

Y

I

I

P-

I R i

t I

l.L 'i I I

Figure 13.-

10

I 1 I :

20 30

Effect of landing-gear pod on longitudinal characteristics.

2

Gear pod

Off On Off On

1 0 Crn

Landing flap setting LDG; ta i l on.

I I

I

I i

i I

I r I

-1

32

r - ---

lo

CL

CD

0

-5

1

c m 0

-1

4

2

0

-2

-4 -10 0

I 1

i ' I

i -1 I

10 20 30

cP

0 0 0 1.99 A 2.83

Figure 14.- Tail-on longitudinal data for cruise flap setting CR.

1 0 -1

33

I 111 I 11111 I I 11111111 II 11.1 I. I

L D -

c m

CL

CD

34

10

5

0

1

0

-1

8

6

4

2

0

-2

I I ,I ..., ---, , . .. . -

-10 10

Figure 15.-

20 30

Tail-on longitudinal data for take-off flap setting TO.

1 0

Cm

-1

0

1

0

-1

-2

6

C L

4

2

CD

0

-2 -10 0 10

T

20

0 A n

‘ I

1

Figure 16.- Tail-on longitudinal data for landing flap setting LDG.

ccI I 2.83 On 2.83 Off On t 0

0 -1

Cm

35

r--

I

36

--

'" 0-

~ , $ , -'

;i Q)

":;; c. 0 >-

~

"~.--

? Q)

.><

'" >-

a; "0 0 E

'0 >. "0 ::::J

-::;;

~ 7 Q)

.>< 0 E VI

l

(b) Three-quarter rear view. L-69-1397


37

----------------- -.

0

1

'm o

-1

8

6

C L

4

2

0

CD

Figure 18.-

I 0 10

a, deg

I

0 0 A D

d " I

20 30

Effect of t h rus t deflectors on longitudinal characteristics.

0 1.99 2.83 2.83

1

On On On Off

t

I

0 Cm

Take-off flap setting TO; ta i l on.

-1

38

L D -

10

5

0

1

Cm o

CL

CD

-1

8

6

4

2

0

-2 0 10 20

k, 1’1

30

0 1.99 2.83 2.83

1 1

1 Cp Deflectors A i l e rons

O n Deflected O n Deflected On Deflected Off Undeflected

0 Cm

Figure 19.- Effect of t h r u s t deflectors and drooped ailerons on longi tudinal characteristics. Take-off flap setting TO; aileron deflection, 400; ta i l on.

10

0

1

-1

8

6

CL

4

2

0

CD

-2 10

C,, Spoiler

0 Deflected 1.99 Deflected 2.83 Deflected 2.83 Undeflectec

20 30 1 0 Cm

-1

(a) Spoiler deflection, loo.

Figure 20.- Effect of symmetric inboard spoiler deflection o n longitudinal characteristics. Take-off flap setting TO; ta i l on.

40

L D -

10

5

0

1

C m o

-1

8

6

CL

4

2

0

CD

-2 -10 10

4 deg 20 30

(b) Spoiler deflection, 60'.


C,, Spoiler

0 Deflected 1.99 Deflected 2.83 Deflected 2.83 Undeflected

1 0 Cm

-1

41

10

0

1

c m 0

-1

6

CL

4

2

0

CD

-2 -10 0 10 20

Figure 21.-

30

(a) Alternate take-off configuration AT02.

1. A 2.83 t~ 3.41

Longitudinal characteristics of an alternate take-off flap setting.

1 0

Tail on.

-1

42

1

c m 0

-1

8

6

CL

4

2

0

CD

-2

0 n

-10

c

20 30

(b) Comparison with basic take-off flap deflection at Cp = 2.83.


tt Flap configuration 1

AT02 TO

0 C m

43

i - I

I

I ro

0

2

1

0

-1

8

0

A n

0 Deflected 1.99 Deflected 2.83 Deflected 2.83 Undeflected

C m

2

CD

0

I ! 1 , -2 10 20 30 0 -1

Figure 22. Effect of ai leron droop on longitudinal characteristics. Landing flap setting LDG. Ai leron deflection, 400 (each); ta i l on.

44

L D -

Cm

CL

CD

5

0

1

0

-1

8

6

4

2

0

-2 -10 30

clJ Spoile

0 Deflected 1.99 Deflected 2.83 Deflected 2.83 Undeflecte

1

Figure 23.- Effect of symmetric inboard-spoiler deflection on longitudinal characteristics. Landing flap setting LDG: spoiler deflection, 600: ta i l on.

45

I111111111III II.III. 11111111 I I 1 I 1 111111,1111 I I I

0

0

cnl -1

-2

8

6

CL

4

2

CD

0

-10 0 -2

Figure 24.-

20 30 0

(a) Horizontal-tail incidence, IOo.

Effect of varying ta i l incidence on longitudinal characteristics.

t.

LcI

0 1. 99 2.83

-1 c m

Landing flap setting LDG.

I I

I

-2

46

5

L D -

0

1

cnl 0

-1

8

6

CL

4

2

CD

0 -10 0 20 30

(b) Horizontal-tail incidence, -5'. -100.


0 0 0 A

10 - 5 -10 - 5

0 Cm

-1

47

I

5

L D -

0

2

1

Cm

0

-1

8

6

CL

4

2

CD

0

20 30 2

ccl

0 0 17 1.99 A 2.83

1 Cm

0 -1

(c) Horizontal-tail incidence, -15O.


48

CY

Cn

1.0

0

-1.0

.1

0

-. 1

.1

0

1

(a) Basic data.

i i r 0 0 0

ztn

0 0 1. 99 1.99

10

i

i 1

'1 I 1

I

p 1 Q

0

Figure 25.- Tail-off lateral characteristics for cruise flap setting CR.

49

cyp -.01

-. 02

0

c”p -.002

-. 004

0

-. 002 0

I[ 0 I7

1

- T

i

2

I

3

(b) Static stability derivatives.


50

CY

Cn

CZ

1.0

0

-1.0

.1

0

-. 1

.1

0

-. 1

ciJ

0 1.9 2.8

-30 -20 - 10 0 10

(a) Basic data.

Figure 26.- Tail-off lateral characteristics for take-off flap setting TO. a = loo.

51

0

cyp -.01

-. 02

0

cnP

-. 002

.002

0

-. 002 0 1 2



52

CY

Cn

CZ

1.0

0

-1.0

.1

0

-. 1

.1

0

-. 1 -30 -20 - 10

(a) Basic data.

0 0 A

0 1.99 2.83

Figure 27.- Tail-off lateral characteristics for landing flap setting LOG. u = loo.

53

0

CYp -.01

-. 02

0

cnp -.002

-. 004

.002

0

-. 002 0 1



3

54

1.0

CY

-1.0

.2

.1

Cn

0

CZ

-. 1

.1

0

-. 1 -30 -20 - 10 0

P, deg

(a) Basic data.

0 0

10 30

Figure 28.- Tail-on lateral characteristics for cruise flap setting CR. a = Oo.

55

0

-. 02

-. 03

-. 04

.006

.004

0

0

-. 002

-. 004

-. 006

L

0



2

56

2.0

1.0

CY 0

-1.0

-2.0

. 3

.2

.1

Cn 0

-. 1

-. 2

-. 3

. 1

cz 0

-. 1 -30

1

I

1

1

I : -20 -10 0

10 1 20 30

P, deg

(a) Basic data.

Figure 29.- Tail-on lateral characterist ics for take-off flap setting TO. a = 100.

57

0

-. 02

-. 03

cyP

-. M

-. 05

-. 06

. 010

.008

.006

cnP

. OM

.002

0

0

CZB -.002

0 -. 004

2 3



58

2.0

1.0

C Y 0

-1.0

-2.0

. 3

.2

. 1

C n 0

-. 1

-. 2

-. 3

. I

cz 0

-. 1 -20 -30 -10 0

P, deg

(a) Basic data.

10 20 30

Figure 30.- Tail-on lateral characteristics for a n alternate take-off flap setting AT02. a = 100.

59

0

-. 01

-. 02

cyP

-. 03

-. 04

-. 05

.008

.006

cnP .004

.002

0

0

QP -. 002

0 -. 004 1 2 3



60

CY

Cn

C l

I

Figure 31.-

2.0

1.0

0

-1.0

-2.0

. 3

. 2

.1

0

-. 1

-. 2

-. 3

.1

0

-. 1 -30 -20 - 10 0

P, deg

(a) Basic data.

10

Tail-on lateral characteristics for landing flap setting LDG.

20

a = 100:

30

it = -100.

61

I

-. 02

-. 03

-. 04 cyP

-. 05

-. 06

.012

. 010

008

0

.002

0

0

-. 002 cz B

004

0 1



62

1.0

0

CY

-1.0

-2.0

.2

.1

Cn

0

-. 1

.1

Figure 32.-

0

-. 1 ,

-30 -20 -10

Lateral characterist ics with inboard spoilers deflected.

0

0 I7

cP

0 1.99

10

Landing flap setting LDG.

20

t

30

Spoiler deflection, a@; a = IOo; it = @; ta i l on.

63

1.0

CY 0

-1.0

. 1

Cn 0

-. 1

. 1

C l 0

-. 1 10 0

t 0 0 0 1.99 1 n 2.83

c

I 2c

(a) Inboard-spoiler deflection, 100.

Figure 33.- Lateral characteristics with right spoilers deflected. Take-off flap setting TO.

64

CY

Cn

CZ

1.0

0

-1.0

.1

0

-. 1

.1

0

-. 1

1 b

1 1 1. P 1

I

I

a P 0

1 1 T 0 A

10

(b) Outboard-&oiler deflection, 100.


clJ

0 2.8

20

65

1.0

CY

-1.0

. 1

Cn 0

-. 1

.1

CZ 0

-. 1 - 10

0 A

10

0 2.83

i L

20

t

30

(c) Outboard-spoiler deflection, 300.


66

CY

Cn

CZ

1.0

0

-1.0

.1

0

-. 1

. 1

0

-. 1

I I I I I I I I I I I I I 1-1 1 I I I I I I I I w I I I I I I I I -

0

-10 0 10 a, deg

(a) Outboard-spoiler deflection, loo.

20 30

Figure 34.- Lateral characteristics wi th r i gh t spoiler deflected. Landing flap setting LDG; ta i l on.

67

1.0 a71:

0

-1.0

.1

0

-. 1

0

- I I

I I / I

11 R ( I

0

I I I i F

r

L

I I

I I

r i

-

-

10

(b) Outboard-spoiler deflection, 300.


0

20 30

68

CY

1.0

0

-1.0

.1

Cn

0

0

CZ

- 1 * o

0 A

m fl 1 I

F I

10

cP

0 2.83

20 30

Figure 35.- Rudder effectiveness, take-off flap setting TO; a = 100.

69

1.0

-1.0

.1

3 Left Right Right

inboard inboard outboard

0.50 0.50 0.50

0

C n

-. 1

-. 2

0

-. 1

-10 -. 2

Left outboard

0 0

0 10 20 a, deg

(a) Effect of varying angle of attack.

30

Figure 36.- Lateral characteristics w i th left outboard engine not operating. Take-off flap setting TO; ta i l on.

70

-20 -10 0 10 20

(b) Effect of varying angle of sideslip; a = 100.


LLLLE

30

71

CY

C, per engine

Left Right inboard inboard

Right outboard 1 [ outboard

- 1 0. 71 0. 71

1 1 10

1 4 I t I

1 I

1

t 1 i 20

CZ II I -. 1

-. 2- 30

(a) Variation wi th angle of attack.

Figure 37.- Lateral characteristics wi th left outboard engine not operating. Take-off flap setting TO; aileron deflection, ZOO (each); spoiler deflection ( r ight outboard), 300; rudder deflection, -300.

72

CY

1.0

0

-1.0

-2.0

.2

.1

Cn 0

-. 1

-. 2

.1

C l 0

-. 1 -20 -10

I i

1 1

1 1 1

1

1 1

0 I 3 9 deg

10

( b ) Variation wi th angle of sideslip; a = 100.


30

73

1.0

CY O

-1.0

.1

C n 0

-. 1

.1

0

cz -.l

-. 2

-. 3 -10

0 0

.71

CP per engine

Left Left Right Right outboard inboard inboard outboarc

0.50 .71 . 7 1

0

0. 71 0.50 .71 . 7 1 . 7 1 .71

, / I , . . , I : : ! :

20 30

(a) Variation wi th angle of attack.

Figure 38.- Lateral characteristics wi th lef l outboard engine not operating. Landing flap setting LDG.

74

2.0

1.0

CY 0

-1.0

-2.0

. 3

. 2

. I

C n 0

-. 1

-. 2

-. 3

.1

0

C l

-. 1

-. 2 -30

C,, per engine Left outboard

7 0 4 0 3 .71

Left inboard

0.71 .71 .71

Right inboard

0.50 .71 .71

-20 - 10

Right outboarc

0.50 .71 .71

il I,

0 P, deg

10

(b) Variat ion wi th angle of sideslip, a = IOo.


20

75

CZ

-. 10

-. 08

-. 06

-. 04

m

Engine ou t

Spoiler deflected

Left outboard

0

.50

def I ected

I l i ~ ~ i i i ~i I ! ! ! I ! ! I ! ! K

C,, per engine

inboard inboard outboard Right 1 0. 71

.50

0.50

.58 Oo50 .50 1 . . . . , . . . , . , . . , . . 1 . . ,-... . , . . .&-.-, _..._, , . . . I_L2 ~ . . . .. . , 1 . . I ,m

0 1.0 2.0 3.0 4, 0 5.0 6.0

C 1.

Figure 39.- Roll ing moment t r i m capability of outboard spoiler, one outboard engine not operating. Landing flap setting LDG.

76

1

' 1

-1

i

6

l i i . I

4

I

cL

2

0

CD

-2 -10

Figure 40.-

0 10 a, deg

20 30

0 A n

l l l l l l l cp Pe

Left Left Jutboard inboard inboard

0.50 0 0.50 0 .71 .71 .71

Longitudinal characteristics with lef l outboard engine not operating.

i I

engine

Right

.71

.71

1 iliitf Right

Iutboa rd

0.50 .71 .71

1 0 Cm

Take-off flap setting TO.

-1

IS

77

1

cnl 0

! ! / ! I . .

i I S 1 i I I

50 71 71

0.50 1 -1

8

6

4

2

0

I .71 .71

CL I

CD

1 -10 0 10 20 30 0 c m

-1 4 deg

Figure 41.- Longitudinal characteristics with left outboard engine not operating. Landing flap setting LDG.

78

10

0

1

cm 0

CL

CD

-1

8

6

4

2

0

-2 - 10 0

/ I

I i n i

10 4 deg

20 30 1

Ccc Spoiler

0 Deflected 1.99 Deflected 2.83 Deflected 2.83 Undeflectec

Cm 0 -1

Figure 42.- Longitudinal characterist ics with r ight inboard spoiler deflected IOo. Take-off flap setting TO.

79

10

0

1

c m 0

-1

6

4

CL

2

CD

0

-2 -10 20 30

(a) Spoiler deflection, IO0.

' Spoiler ,IJ D ef I ected

83 Deflected 83 Undeflected

1 0 C m

-1

Figure 43.- Longitudinal characterist ics wi th r ight outboard spoiler deflected. Take-off flap setting TO.

80

L D -

c L

CD

10

5

0

1

CIn 0

-1

6

4

2

0

-2 0 1.0 20

1

i I

~ I

30

Spoil€

0 0 Deflected A 2.83 Deflected D 2.83 Undeflected

1 0 C m

( b ) Spoiler deflection, 30°.


81

5 L D -

0

1

Crn o

-1

6

4

CL

2

CD

0

-2 -10 10 20 30

0 A 0

0 Deflected : 2.83 Deflected 1 2.83 Undeflected

1

Spoik

0 C rn

-1

(c) Spoiler deflection, 600.


a2

1

C m o

-1

6

4

CL

2

0

CD

-2

0 n

Controls

Def I ected U ndeflectec

I

- 10 0 10 20 a, deg

30 1 0

C m

ill -1

Figure 44- Longitudinal characteristics wi th left outboard engine not operating. Take-off flap setting, TO; aileron deflection, 20° (each); spoiler deflection ( r i gh t outboard), 30°; rudder deflection, -30°; Cu = 2.83.

83

L D -

CL

CD

5

0

2

Cm 1

0

8

6

4

2

0 -10 10 20 30

0 0 30 0 10

0 1.99 30 A 1.99 10 n 1.99 0

2 1

Cm

Figure 45.- Effect of r i gh t outboard spoiler deflection on longitudinal characteristics. Landing flap setting LDG.

0

84

cL

6

5

4

3

2 ,

1

0

Left fl outboard

0 0.71 0 , . 50 D O 0 0

CP per engine

Left Right nboard inboard

Right ou tboa rd

0. 71 0. 7 1 .50 .50 . 7 1 . 7 1 . 7 1 .50

0. 71 .50 . 7 1 .50

0

1 2 3 4 5

Figure 46.- Effect of t h r u s t distribution on l i f t . Landing flap setting LDG; a = Oo.

I

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