EXPLORATORY LOW-SPEED WIND-TUNNEL STUDY OF CONCEPTS 

DESIGNED TO IMPROVE AIRCRAFT STABILITY AND COEiTROL 

AT HIGH ANGLES OF ATTACK 

BY 

David Edgar Hahne 

B.S June 1981, Vi rginia Polytechnic Institute and State University 

A Thesis submitted to the Faculty of 

The Gradueaie School of Engineering and Applied Sciowe of 

The George Washington University 

in partial satisfaction of the requirements for 

the degree of Master of Science 

December 1985 
ABSTRACT 

A wind-tunnel investigation of concepts to improve the high angle-of-attack 

stability and control characteristics of a high performance ai rcrsft has been 

conducted. The effect of vertical tail geometry on stability and the effcctiveness of several conventional and unusual control concepts has been determined. 

found that vertical tail location, cdnt angle and leading edge sweep could 

i nfl uence both 1 ongi tudi nal and 1 ateral -di rectional stabi 1 i ty . 

cepts tested were found to be effective and to provide control into the poststall angle-of-attack region. 

These results were obtained over a large angle-of-attack range. It was 

The control conii 
ACKNOWLEDGEMENTS 

The author wishes to thank the National Aeronautics and Space Administration 

for providing the opportunity to perform this research, prepare this thesis, and 

complete all the other requirements for this degree. The author is grateful to 

Professor &hn P. Campbell who acted as academic and thesis advisor. His 

guidance and encouragement have been instrumental in the successful completion 

of my degree program. 

guidance throughout this test program, Ms Jennifer D. Irby for her speedy 

preparation of this document and Mr. Stanley H. Husch for his exceptional work 

in preparing the figures contained herein. Special thanks also go to Mr. Mark 

A. Croom and Mr. Luat T. Nguyen for their helpful insights in ana’lyzing and 

i nterpreti ng the resul ts presented in thi s thesi s. 

I would also like to thank Mr. Joseph L. Johnson for his 

iii 
TABLE OF CONTENTS 

Page 

TITLE.................................... .......................... i 

ABSTRACT......... .................................................. ii 

aCKNOWLEDGEMENTS.. ................................................. iii 

TABLE OF CONTENTS....................... ........................... iv 

LIST OF TABLES.... ................................................. v 

LIST OF FIGURES...............,...................... ............... vi 

LIST OF SYMSOLS.................................................... viii 

Chapter 

I INTRODUCTION...... .......................................... 1 

I1 MODEL AND 4PPARATUS.............. ........................... 3 

I11 WINO TUNNEL TESTS............. .............................. 4 

IV ANALYSIS TECHNIQUES ......................................... 5 

V RESULTS AND DISCUSSION........ .............................. Y 

VI SUMMARY OF RESULTS........................... ............... 20 

REFERENCES ......................................................... 22 

TABLES..... ........................................................ 23 

FIGURES.................................... ........................ 25 
LIST OF TABLES 

Page 

1. Mass and geometric characteristics of airplane ................ . . 23 

V 
LIST OF FIGURES 

Page 

1 . 

2 . 

3 . 

4 . 

5 . 

6 . 

7 . 

8 . 

9 . 

10 . 

11 . 

12 . 

13 . 

14 . 

15 I 

16 . 

17 . 

18 . 

19 . 

20 . 

Effect of improved control power on maneuverability ................ 25 

Effect of improved control power on configuration optimization ..... 25 

Sketch of basic model configuration ................................. 26 

FOreDduy geometry ................................................... 27 

Vertical tail geometry .............................................. 28 

tongi tudinal and 1 &era1 control surfaces ........................... 29 

Horizontal tail and canard positions ................................ 30 

Photograph of model in 12-Foot Low-Speed Tunnel ..................... 31 

Definition of body axes system ...................................... 32 

Effezt of vertical tail geometry on longitudinal characteristics .... 33 

Effect of .d wtical tal 1 geometry on 1 stera1 -di rectional 

characeekistics ..................................................... 31 

Effect of vertical t?il sweep or! longitudinal characteristics ....... 35 

Effect of vertical tail sweep on laterdl-directional characteristics 36 

Effect of vertical tail spanwise location on longitudinal 

characteristics ..................................................... 37 

Effect of vertical tail spanwise location on lateral-directional 

characteristics ..................................................... 38 

Effect of vertical tail cant angle on longitudinal characteristics . 39 

Effect of vertical tail cant angle on lateral-directional 

characeeristics ..................................................... 40 

Effect of leading edge flaps on longitudinal characteristics ........ 41 

32 

a . Longitudinal characteristics 

b . Lateral-di rectional characteristics 

Effect of scheduled leading edge flaps .............................. 

Trailing edge extension flap effectiveness .......................... $3 

a . Slotted flap 

b . p1 ain flap 

vi 
21 . Effectiveness of ailerons for pitch control ......................... 44 

22 . Horizontal tail effectiveness ....................................... 45 

23 . Effect of canard fairing on longitudinal characteristics ............ 46 

24 . Effect of canard fairing on lateral-directional characteristics ..... 47 

25 . Effect of canard deflection on lateral-directional characteristics .. 49 

26 . Canard effectiveness ................................................ 49 

27 . Centerline vertical tail effectiveness .............................. 50 

28 . Effect of rudder hingeli ne sweep on rudder effectiveness . 6r = .30° . . 51 

29 . Tip rudder effectiveneses ........................................... 52 

30 . Effect of aileron geometry on roll control . 63 = -2o0............... 53 

31 . Effectiveness of wing tip deflection for roll control ............... 54 

32 . Departure resistance of test configuration .......................... 55 

33 . Effect of ARI gain on control respcnse .............................. 56 

34 . Maneuverabil i ty comparison to ccirrei>t ai rcraft ...................... 57 

a . Pitch rate acceleratio2 

b . Yaw rate acceleration 

c . Roll rate acceleration 

35 . Effect of inertia coupling on maximum roll rate ..................... 58 

vi i 
LIST OF SYMBOLS 

m 

S 

s 

V 

Y 

2y/b 

a 

. aspect ratio 

wing span, ft 

lift coefficient, Lift/I;S 

rolling-moment coefficient, Rolling moment/qSb 

pi tching-moment coefficient, Pitching moment/qSc 

yawing-moment coefficient, Yawing mornent/qSb 

side- force coefficient, Side force/qS 

mean aerodynamic chord, ft 

mass moments of inertia about xb, Yb, zb body axes, sIugs-ft2 

mach number 

ai rcraf t mass, slugs 

angular veloctiy about the Xg, Yb, and z;, body axes, deglsec 

roll rate about the velocity vector, &g/sec 

angular acceleratioq about the Xb, 

freestream dynamic pressure, 1 bs/ft2 

wing area, ft2 

Laplace transfo m variable 

Y,,, and Zb body axes, deg/sec;! 

linear velocities along the xb, Yb, zb body axes, ft/sec 

freestream velocity, ft/sec 

body axes 

span ordinate, ft 

nondirnensiortal semi span lwation 

angle cf attack, deg 

angle of sideslip, deg 

viii 
Q 

P 

incremental roll ing lnornent coefficient 

incremental yawing momcnt coefficient 

Incremental side force coefficient 

aileron deflection, deg 

centerline vertical tail deflection, deg 

canard def 1 ect j on, deg 

trailing edg? extension flap deflection, deg 

leading edge flap deflection, deg 

horizontal tail deflection, deg 

rudder deflection, deg 

aft swept tail rudder deflection, deg 

forward swept ta5l rudder deflection, deg 

tip rudder deflection, deg 

wing tip deflect'on, deg 

root of cha-acterfstic equation 

afrcraft roll attitude, deg 

a5r density, slugs/ft3 

Stability derivatives 

2v 2V 

2v 2v 

ix 
Subscripts 

IS inboard 

M6 mid board 

06 outboard 

X 
CHAPTER 1 

I NTROBUCTION 

The emphasis an sustained nonafterburning supersonic cruise makes aircraft 

with high-f ineness-ratio fuselages, \OH-aspect-ratio, highly-swept wings and 

highly-integrated control surfaces attractive because of low cruise drag 

(Reference 1). These sane features, however, make the aircraft more susceptible 

to the development and sudden breakdovrn of strong vortical flowfields at angles 

of attack and speeds encountered during maneuvering. 

above 15" angle of attack where the vortical flowfields can produce r~onlinearitier and regions of unstable behavior in the longitudinal, lateral ana directional aerodynamics. In addition strong spanwise flowfields arise which reduce 

This is especially true 

the effectiveness of conventional control devices. Such aerodynamic characteristics complicate control system design especially for unstable aircraft 

where powerful control surfaces are the key to a sirccessfcrl active flight 

control system. 

As illustratzd conceptually in figure 1, maneuverability aecreases with 

increasing angle of attack such that very little capability is available at and 

above the stall. 

tibility to loss of control and spins. 

On some configurations, limiters are imposed to avoid suscep 

Providing controls that maintain a high 

level of effectiveness will allow future high performance aircraft to exploit a 

much larger angle-of-attack envelope, including brief excursions into post stall 

flight conditions. A major potential payoff is configuration optimization 

flexjbility which results frov the fact that the hish levels of conzrol effectiveness combined wi th appropriate control 1 dws and fly-by-w i re technology WS 11 

allow increased reliance on st&ility and control augmentation. This will allow 

the designer greater freedom in optimizing the configuration to meet other 

requirements such as minimum drag at cruise conditions (figure 2). Typically, 

I 
designs that have origi naily been optfmlzed for supersonic cruise. performance 

have had to be modified to obtain acceptable low-speed characteristics at tfie 

expense of crui se performance (2eference 2) . Advanced contro? s technology can 

minimize this required compromise in future high performance alrcreft. 

As part of a NASA Langley Research Center Flight Dynamics Branch and McDonne1 1 Aircraft Company cooperative research program, an inveitigation of advanced 

control concepts for high performance aircraft applications has been pertorm2d. 

The primary objectives were: 

maintain desired level of effectiveness throughout the maneuvering angle-ofattack range, including stall and post-stall conditions; and (2) to stiidy the 

impact that these advances will have on aircrdft maneuverability and the assc- 

(11 to develop aerodynamic control devices that 

ci ated fl i ght dynamics. The generic configuration of thi s study has emerged 

front a nUmb%!i- of studies aimed at developing design yuidelincs far providing 

efficient supersonic cruise, transonjc rcancvzjering capability cowparable to 

current high performance aircraft a~d gmd iw-speedhigh angle-ot-attack stabi1 ity and control (Reference 3). This thesis presents the results of the latest 

phase of tnis group of studies to develop and evaluate conventional an4 unusual 

control concepts for advanced aircraft. This work. includes an eva;uatian of the 

effects of tail geometry on stability as well as an evalrratior: of control effectiveness. Compari sons of manebgerabi 1 i ty Hi th a hi ghly-maneuverabl e ai rcraf t 

are also presented. 

The test program was conducted in the Langley Research Center's IP-Foot 

Low-Speed Wind Tunnet using a strut supporreci scale mode? of tie generic CORfiguration. Data viers taken cwr a large angle-of-attack range at several 

si des1 5 p angi es . 

-2- 
CHAPTER 2 

MODEL AND APPARATUS 

The basic model and its support system are shown in Figure 3. Details of 

the model forebody, vertical tails, control surfaces, and horizontal tails and 

canards are shown in figures 4, 5, 6, and 7, respectively. 

65" swept leading edge with an aspect ratio of 1.95. 

canards all had thin flat-plate cross-sections (for simplicity) with beveled 

leading and trailing edgzs. The wing leading edge was divided into three flap 

segments and the trailing edge had two different flap/aileron configurations 

along with flaps on the inboard trailing edge extension. 

The arro:: wing had a 

The wings, tails and 1 

It was also possible 

to deflect the wing tips as ailerons. 

and could be replaced by horizontal tcsils. 

to use canards. 

The trailing edge extension was removable 

The fuselage could also be adapted 

The rnadel could be configured with three di Fferent vertlca'l 

tail arrangements: 

edge extensions) and a single centerline tail. The vertical tails also had 

several fore and aft positions as well. as the abi!ity to cant either inboard or 

outboard. The fuselage, with flow through ducts, was strut mounted to minimize 

unfaverable effects on vortex bursting as described in Reference 4. 

forward and aft svrept Win tails (mounted on the traiiing 

F3rce and moment data were obtained using a six component strain-gage 

balance mwnted inside the model fuselage. The voltage outputs of the 

strain-gage balance were converted to digital signals by a MEFF 620 

amp1 i fier/mul tiplexor. The NEFF 620 providea the required signal processing for 

use Mith the HP-98458 microco.+pttter tcsed fer data reductiors and storage. The 

data acquisition system r,amp!<d each balance output 100 times over 10 seconds 

dqd averaged the data. Non-diinensional coefficients were then calculated and 

stored on magnetic discs. 

-3- 
CHAPTER 3 

WIND TUNNEL TESTS 

Static 

Low Wed W 

in figure 8 

force tests were conducted in the Langley Research Center's 124oot 

nd Tunnel. X photograph of the model in the test section is presented 

The tests were conducted at a freestream dynamic pressure of 4 psf 

(M = . 05) which corresponds to a Reynolds number of 0.62 x lo6 based on wing 

mean aerodynamic chord. 

0" to GO" at angles of sideslip of 0" and 25". 

wall effects were made to the data. 

both angle of attack and angle of sideslip. 

Data were obtained over an angle of attack range from 

No corrections for base drag or 

F?ow angularity corrections were made for 

In the first part of the test program, studies were made of the effect of 

spanwise location for the aft wept tails and tail cant (from 30" inboard co 30" 

outboard) for the forward and aft swept win vertical ta-ils. Once an optimm 

cant angle was obtained, rudder effectiveness was determined, An all moving 

centerline tail was also investigated. Various pitch control devices were examined includins horizontal tails, canards, trailing edge extension flaps and 

ailerons deflected symmetrically. 

unswept hinge1 ine ailerons and constant chord ailerons. 

were also investigated for roll control. 

Two types of ailerons were investigated: 

Deflected wing tips 

All data were initially obtaired about the body axes (figure 9). All 

lateral-directional data are presented about the body axes, but all the longitudinal data are presented about the stability axes. 

-4- 
CHAPTER 4 

AIJALY SI S TECHNIQUES 

A calculated paramter which is useful far the analysis tlf high angle-ofattack lateral-directional stability is C"g,dyn (Reference 5) , defined as 

Negative values of this parameter indicate a susceptibility to a divergence 

(nose slice). This term is derived from one of the coefficients of the lateral 

open-loop characteristic equation in Hhich A, 6, C, 0 and E are constants. 

Ax4 + Bx3 + CA' + Dx f E = 0 (1) 

This eqlratian will have unstable roats if any of che coefficients or Routh' s 

discriminant (BDC - AD2 - B2E3 becon:es negative. 

coefficient becomes negative a directional divergence usually occurs. 

Neglecting the products of inertia the C-coefficient is given by: 

In particular, if the Cwhere 

a nd 

-5- 
Neglecting products of derivatives (which in general are <<1) 

Y = 42 LK~C,~ cos a- K~c,~ sin a 1 (3) 

* *.. i5y removing the common terms and dividing equation 3 cy .x, A'.? 

obtained Normally, this equation is reduced further by assuming cos a = 1, but 

for this study the cosine term was retained for exactness because of the large 

angles o attack involved. The work in Reference 6 h as shown that these assump 

tions have very little effect on the results. 

Another type of divergence results when lateral control response to pilot 

inputs is reversed. In this case, the lateral control divergence parameter 

(LCDP) is used. 

transfer function h eference 7 1 

This parameter is derived from an approximation of the +/&a 

where f(s) is the approximated lateral open-loop characteristic equation. 

For thii derivation only the steady state terms (f(o)) are considered. 

For normal roll response, $/&a < 0. 

neglected rind the transfer function becmes 

Since Cysa is generafly <<1 i t can be 

Multiplying byl.!?)2f (0) (a p,sf:ive value) eqtiation 5 reduces Lo 

'C 

-6- 
For almost a1 1 configurations the control derivative Ceda is negative, therefore 

dividinq by it dl1 reverse the inequality and yield the final result 

where positive values indicate normal response to roll control and negative 

values indicate a reversed response. If an aileron-rudder interconnect (ARI) 

system is used then equation 6 becomes 

where K is the ARI gain. 

The equations used for calculating 6, +, and i, come from the rncmment 

equations of the eqtiirtions of motion. By neglecting a?’ of the rate terms and 

assuming tl;at the moments are generated entirely by colitrsf deflection, the 

following relationships can be obtained: 

Because of inertia coupling the maximum roll rate (about the stability axis) 

a1 1 cwabl e is detemii ned by the icrna~nt of avai 1 ab1 e pi tcii contrcl . 

about the velocity vectGr, cspeci6lly at high angles-of-attack, a typically 

fuselage-heavy fighter aircraft will experience a nosed p pitching moment 

When roll i ry 

-7- 
(Xeference 81. 

assumption is made that the only source of pitching moment is from control 

deflection, the steady state pitching moment equation becomes 

For an aircraft symmetrical about the X-2 plane, and if again the 

(I~ - I,) pr +-q s C kae 5@ = o (11) 

S ubs t i tuti ng 

i, = p COS a S 

and 

r = p, sin a 

(12) 

(13) 

Using nose-down control deflections it is possible to determine the maximum stability axis roll rate allowable without incurring an uncontroll ible pitch p. 

This relationship is given by 

-8- 
CHAPTER 5 

RESULTS AND DISCUSSION 

Static Stability and Control 

Vertical Tail Geometry.- Figure 10 shows the effect of the centerline and 

twin tails on longitudinal characteristics. The centerline tail has no effect on 

lift or pitching moment but the twin tails reduce CL~~~ and also i~crease the 

magnitude of the pitching imnent;. 

(Reference i) indicated thdt this decrease in lift and the related increase in 

Earlier work with a similar configuration 

pitching moment were duc to early vortex bursting caused by the twin tails. 

A comparison of the effect of the centerline tail and the twin tails on 

lateral-directional characteristics is made in figure 11. The centerline tail 

maintains djre(:tiGnal statjjljty IC, ) un:il n = 25" and contjnoxs to provide 

stabilizing increment until 3 = 30'. With r'urtl-ier increases in angle of attack 

the cerrterlinc tail is destanifiring. Past u = 53" tile r;B"iabilizing effects of 

the forebody (Reference 1) dominate. The twin tails have a stabilizing effect 

throughout the angle-of-attack range tested. 

tional stability, however, between a = 15" and 2?O. Adverse sidewash from the 

a 

Th!re is a sharp loss of direcvring leading edge vortex zystem is the most probable cause of tnis reduced stability region. 

Dihedral effect (-CL ) is also affected by the tail geometry. At low 

9 

angles of attack both tail configurations produced an increase in effective 

dihedral. 

of the adverse side force on the tail which is well above tne aircr3f.C center of 

gravity. k'hen this destabilizing rolling mcment is added to the ?o,.r stability 

Between a = 30' and 50" the centerline tail is destabilizing because 

caused by asymmetric vortex bursting on the dng a region of instability results 

-9- 
from a = 35" to 40". The twin vertical tails force the wing vortices to burst 

symmetrically, which greatly increases the effective dihedral betwezn a = 25" 

and 40". 

A comparison of the longitudinal chardcteristics for the aft and forward 

swept twin tails is made in figure 12. The forward swept tails had a slight 

effect on lift between a = 16" and 19" causing a more abrupt break in the 

pitching moment curve. The longitudinal position of the forward swpt tail had 

very little effect on either lift or pitching moment. 

Figure 13 shows the effect of forwawl and aft swept vertical tails dn 

lateral-directional characteristics. The forward swept tails show improvement 

in directional stability over the aft swept tails in the region from a = 15" to 

20". Above cs = ZOO, however, the aft swept tails provide much more directional 

stability. Nith the forward swept tails in the aft position (thir; positiotl 

gilies the forward %ept tails approximately the same tail vrjltirne 65 the aft 

sviept tails) directional stability is increased to match the aft wept tails st 

the lower angles of attack. 

quate levels of stability with minimal differences between them except between u 

= 10" and 20" wkere the aft swept tails exhibits less effective dihedral. 

Laterally, all three configurations maintain adeThe effect of spartwise location of the twin vertica; tails on longitudinal 

and lateral-di rectional characteristics Mas a1 SO investigated and the results 

are shown in figures 14 and 15, respectively. 

edge extecsion was reinoved and the aft swept tails used. 

no change except for a sliynt increase in maxirntlm lift for the outbotird tail 

position. 

inboard position (a = 12'). 

effect on dihedral. ?here Was an improvement, however, in directional stability 

For both positions the trailing 

The lift curve showed 

This position elso czused the pitcb-up t3 occur 2' earlier thm the 

SpanPtise location of the vertical tails had no 

- 10 - 
by moving the tails outboard. Directional stability is maintained with the outboard tails through the region from CL = 15" to 25" where the inboard tails exhibit a reduction in stability. 

Of all the vertical tail parameters varied, tail cant angle had the rnost 

impact on both the longitudinal and lateral-directional characteristics. The 

longitudinal characteristics are presented in figure 16. Cant angle has little 

effect on lift with the exception of the 30" outboard cant. This configuration 

exhibited an increase in maximum lift. This can be attributed to the additional 

lifting surface created with surh an extreme outboard cant. Vertical tail cant 

angle also had a noticeable effect on the pitching moment. For all cases the 

initial pitchup occurred at about a = 8". 

cants affected pitching moaent more than either 15" cant. The 30" outboard cant 

increased the severity of the pitch-^! p' and caused a stab1 c pi iclri ng moment SI ape 

Both the inboard and outboard 30" 

past: CI = 15". The 30" inboard cant, hawever, dr?cr.eazed the pitch p but delayed 

a stable pitching moment slope until a I- 20" as did the other con:'igiirittiC:ns. 

In Reference 1, which showed this same trend, it was found that the cant angle, 

not the spanwise location of the vertical tail tip, affected the pitching moment 

characteristics. 

understand the fluid mechanics involved. 

Extensive flow visualization studies will t ? needed to Fully 

Figure 17 shows the effect of vertical tail cant angle on the lateraldirectional stabi 7 i ty of this configuration. ael ow 14" angle of attack, cant 

angle has little effect on directional stability. Past a = 14", canting the 

tails outboard, much like moving the vertical tails spanwise more outboard, 

increased rli recticinal stabil i ty. Effective dihedral was a']. SQ intre2sed, siioS+'ir;g 

the saw trends as shom by the tests of Reference 1. Cantirg the tails inboard 

had the opposite effect. That is both lateral and directJondl stability were 

- 11 - 
decreased. 

not surprising that inboard canting of the tails places the vertical tails in an 

unfavorable f1o.d field. 

Since in sideslip the windward wing vortex moves more, inboard it is 

Leading Edge Flaps.- The deflection of a three-segmented leading edge flap 

was investigated primarily to reduce the pi tchu p tendencies. 

the pitching moment data for the various flap settings tested. 

figurations, the static margin increased at low angles of attack, and the pitchup was delayed and the severity of the pitch p reduced. The data indicate that 

the inboard-most section of flap contributes most to this improvement in 

pitching moment characteristics apparently by reducing the strength of the wing 

vortex. 

flaps (that is, using a different flap deflection at each angle of attack) it is 

Figure la shows 

For all conThe datd of figure 19 indicate that by scheduling the leading edge 

possi bl e to reduce the severity of the pi tch-u p signi f icsrrtl y w5 tkaut sacriFicf ng directional stabil ity. 

dul i ng f7 dp deflection would reduce the a:ii9uni crf nose-d&~ crJntrol needed between 

Decreasing the' wverity of the pi tch4.i p by schea = 8" to 15" and allow t,he clean configuration to be flown in the angle-ofattack range used for supersonic cruis? flight (a : - 6'). 

Langi tudinal Control .- Fcx methods for obta,nincj adequate pitch control 

were tested: 

flaps, horizontal tails, and canards. 

sented in figures 20 through 26. 

flaps on the trailing edge extensions, use of the ailerons as 

The data OR these configurations are preThe trailing edge extensiovl could utilize either 

flap, both capable of deflecting ttp and down. Figure 

power available for the slottec flap &rile Figure 20 

a plain fiap or a slotted 

23 (a) shows the control 

b) snows the same Gala forthe p1aiR flap. 

rnomcint, a?lowing the Configuration to be triinrned to ,tbout 40' angle of attack. 

The two concepts provide the Sdme amount of nose-up pitching 

- 12 - 
The plain flap, however, provides slightly mwe nose-down nownt than the 

slotted flap at the lower angles of attack and maintains this advantage until 

'LMAX 

Itre use of ailerons defl cted symmetrically was investigated as J primary 

source of pitch control and a so as an augmentation of the trailing edge extension 

flaps. 

ha1 f the pitching moment that the plain trai 1 ing edge extension flap produced. 

This coupled with rapid loss in effectiveness past a = 15" makes the ailerons 

unacceptable for primary pitch control. However, in conjunction with the flaps, 

The data of figure 21 show that the ailerons produce a little more than 

the ailerotis increase the available pitch power in the lower angle-of- attack 

range where additional pitch control Has needed. The ailerons also extended 

the trim angle of attack to 42", approximately 7' beyond the stall. 

By removing the trailing edge exteiisions, it is possible to add harizontal 

tails to the confjgaration. A clear benefit uf the horizontral tail configuration 's that it reduces the severity of the pitchup around o = 15" 

(figure 22 . The available control power using the horizontal tail is camparable to that of the flapaileron combination and the maximum trim angle of 

attack is ncreGsed to 44". Unfortunately, on this configuration the use of 

horizonta'l tails requires either fuselage mounted twin vertical tails or a 

single centerline vertical tail, both of which have poor lateral-directional 

stability characteristics (Reference I). 

The use of canards required the addition of a fairing along the upper surface of the inlet area. The effects of this Fairing on the longitudinal and 

latcral-directional characteristic; Here examinerf before the canards-*ere adderr, 

The longitudinal data (figure 23) show a slight Increase in lift around CL~~~, 

but a much greater change in pitching moment. The addition of the fairing 

- 13 - 
decreases longitudinal stability below a = 12", delays the stable break in the 

pitching moment curve to a higher angle of attack (ct = 38") and increases the 

nose-u p pitching moment considerably. 

making adequate nose-down control a problem and high angle-of-attack stability 

All of these results are undesirable, 

poor. 

24). Directional stability is dnchanged and effective dihedral is actually 

The fairing had far less effect on lateral-directional stabilfty (figure 

increased in some regions. 

Figure 25 5 hows that when the canards arc deflected to -40' to produce a 

nose-down moment, Cn, becomes negative at a lo@er angle of attack than for a 0' 

deflection. While there is some effect on Ceg, the changes are no! significant. 

The data of figure 26 indicate that even rJiih a deflection of -40" the canard is 

unable to produce a nose-down moment past a = 12'. 

Directional control - .- Adequate rudder power, especially at high angles of 

attack, is a'lways a coficern with hic,h performance aircraft, 

directional contrca? were evalua't?d as part of this investigation: an clll moving 

Four concepts for 

centerline tail, a conv~ntional rudder, a rudder with a forward swept hinge 

line, dnd tip rudders. 

cepts to be very effective at high angles of attack. 

The work of Reference 9 has shown these last two conThe all moving centerline tail, as was expected, was extremely powerful at 

low angles of attack (figure 27) with more than adequate yaw control from 13" 

of deflection. More yadng moment can be generated with a 30" deflection, but 

the effectiveness drops off sharply past a = 25' as the tail becomes shielded by 

the wing and fuselage. For angles of attack above 30" the larger deflection 

provides only as much yawing rnoincnt as the 10' deflectfen. All rudder Dower 

with either deflection is lost by 3 = 45". 

A comparison of the conventional rudder on the aft swept tail and the 

forward swept redder on the forward swept tail is presented in figure 28. The 

- 14 * 
forward position of forrJard swept tail (FST) coincided with the position of the 

aft swept tail (AST) while the rearward position of the FST was designed to have 

approximately the same tail volume as the AST. 

of rudder power between the RST and the FST is a result of the larger rudder 

area of the FST. The conventional rudder starts to lose effectiveness at a = 

15" because the flow becomes increasingly parallel to the rudder hingeiine until 

by a = 40' all rudder power is lost. The forward swept rudders, however, are 

designed to keep the flow more perpendicular to the rudder hingeline at high 

angles of attack and are therefore able to maintain adequate dirzctional control 

well past 

The difference in the magnitude 

Tip rudders (figure 51 were also studied for obtaining high angle-of-attack 

directional control. As angle of attack increase; the hingeline of the tip rudders becomes more ycrpendicular to the frcestrearn. As ai result of this, the tip 

rudders start to becone effective at % 25" (figure 291 as the rrrdder ofi the 

AST starts tu lose its effectiveness. 

control until the stall where, like the FST rudders, effectiveness is reduced. 

With a larger rudder on the AST, it should be possible to achieve good levels of 

yaw control by using the AST rudders and the tip rudders in conjunction with 

each other. 

The tips rgddcrs nieintairi good yaw 

Lateral Control .- Two basic concepts were evaluated for roll control : 

ailerons and deflectable wing tips. 

dith unswept hingellnes were tested and the results are presented in figure 30. 

Both constant chord ailerons and ailerons 

There is very little difference between the tlrc concepts. 

effectiveness past 3 = 15" because of spanijise flaw on ',ne wiiig, and by a = 30" 

neither are very effective. ':p until this point, however, the constant chord 

ailerons supply slightly more rolling moment. 

acth start losing 

* 15 - 
In order to take advantage of the spanwise flow on the wing, deflectable 

wing tips with a hingeline that is about 45" to the wing chord were tested. Selow 

a= 20" these deflectable wing tips provided slightly le5s rolling moment than the 

constant chord ailerons (figure 31). After a small drop in effectivenesc just 

past a = ZOO, however, the deflectable wing tips maintain an almost constaflt anount 

of roll control to an angle of attack of GO" while the ailerons continue to lose 

effectiveness. By combining these two concepts ft is possible to provide sufficient rolling moment throughout the angle-of-attack range. 

Estimated F1 ight Dynamics 

The last three sections of this chapter deal with the flight dynamics and 

maneuverability predicted using the static data. The configuration used to obtafn 

these rewltr was the forward swept tails using buth aiierom and trailing edge 

extensiori flaps for pitch control and both ailerons and deflectable w'ing tips 

for roll control (table 1). Data 3n the current aircraft used For cofnptrison of 

rnaneuverabil i ty were obtained from Reference 10. 

Departure Resistance.- The data of tigure 32 show that this configuration 

should be very resistant to yaw departur - (nose slice). Positive valuer cf 

cn fhdyn 

cn 5, dy n 

are maintained throughout the angle-of-attack range tested. 

becomes more positive with increasing angle of attack except between 

1i1 general, 

= 16" and 24" and in the pos,t-stall region (a > 35'1. Even though Cn5 is neWtive or very small for this configuration between a = 20" and 46O, the stable 

values Of Ct3 from a = 0" to GO" shculd be cnotigh to preveot a n05e slice frm 

occurring. 

- 16 - 
Control Reversal .- When adverse yawing moment due to aileron deflection 

becomes too high, too much sideslip is generated and control response to a 

pilot roll input will be reversed. This control reversal is predicted when LCDP 

(as defined in Chapter 3) becomes negatide. FGr. ailerons only, this configuration will experience a reversed response to roll inputs past a = 20". In 

order to prevent 'chis, it is necessary to use rudder deflection to offset the 

adverse yaw dlre to aileron deflection. 

pilot or by using an aileron-rudder interconnect (ARI) system. The ARI system 

This can be accomplished manually by the 

provides a prJportiona1 amount of yaw control for a given roll inplrt in order to 

ensure proper control response. With the proper ARI gain (K = 0.351, correct 

response to pilot roll commands can be maintained until a = 45" (figure 33). 

Maneuverabil i 9. The maneuverability of an aircraft is generally measured 

by how fast it can turn about its three axes. This requires a complete data 

base, such as that used in sirnulation str;dies, and knokledgi? of restrictions 

that could limft the maximum allowable rates and accelerations. A less exact 

method is to use the equations presented in Chapter 4 to calculate the maximum 

rate accelerations that can be commanded using a constant velocity and altitude. 

This gives an indication of how fast an aircraft can initiate a turn. Figdre 34 

presents the rate accelerations that can be commanded for a ctirrent highly 

maneuverable aircraft (Aircraft A) and the configuration of this study (Aircraft 

B) . 

The pitch rate acceleration data of figure 34(a) show that belovr the stall 

(both aircraft stall at a = 35") Aircraft A can initiate 2 pftch attitude change 

faster than Aircraft B. Aircraft B, hcwever, is capable of higher pitch rate 

acceleration beyond the stall. Also, Aircraft B shows little change in pitch 

rate acceleration with changing angle of attack. This will provide the pilot 

P 

- 17 - 
with virtually the same longitudinal respsnse to his pitch control inputs 

throughout this angle-of- attack range. 

The results of the yaw rate acceleration calculations are presented in 

figure 34(b). The data indicate that Aircraft A is capable of generating higher 

yaH rate accelerations than Aircraft 5 thrcughout the angle-of-attack range 

tested. A consequence of lower attainable yaw rate acceleration for Aircraft 8 

will be lower effectiveness of a stability augmentation system &AS) in the yaw 

axis. Exactly how important this is depends upon the basic airframe static and 

dynamic stability characteristics. 

Figure 34(c) shows the roll rate acceleration capabilities of the two 

aircraft. Aircraft A is capable of initiating a roll rate quicker than Aircraft 

B below the stnll whereas the opposite is true above the stail. 

the region approaching the stall and above for aircraft vlith highly swept wings 

to exhibit plring rock. tendencies. 

It is comrmn in 

The higher 1-37] rats acceleration capabi’i ity 

of Aircraft 5 in this angle-of-attack range will allow more effective use of 

artificial stabilization to prevent this ding rock. 

High performance aircraft, vthich are typically fuselage heavy, tend to 

pitch p when rolling about the aircraft velocity vector (stability axes) 

because of inertia coupling. This places a limitation on the maximum stability 

axes roll rate that can be maintained. This maximum roll rate, as shown in 

equation 15, is related to the amount of nosedown longitudinal control power 

avail able to overcome the pi tchu p caused by inertia coup1 ing. 

description of itiertia coupling can be found in Reference 8. 

the results of this evaluation for the test configuration and current aircraft. 

aelow the stall, Aircraft A is limiteO less by inertia coupling than Aircraft 3, 

allowing Aircraft 4 to make quicker rolls about the velocity vector. 

A detailed 

Figure 35 shows 

In the 

- 18 - 
post-stall region, however, Aircraft B is able to sustain a higher roll rate 

that is almost constant throughout the remainder of the angle-of- attack range. 

- 19 - 
CHAPTER 6 

SUMMARY OF RESULTS 

The results of this investigation to evaluate conventional and unusual 

control concepts and to study the effect of tail geometry on stability and 

coritrol may be summarized as follows: 

1. A twin vertical tail configuration offered the best overall stability 

characteristics based on static data. 

2. It was found that spanwise location and cant angle of the vertical 

tails could be used to tailor longitudinal and lateral-directional stability. 

3. It was possible to use scheduling of leading edge flaps to rcdtrce the 

severity of the pi tchu p without sacri ficing supersonic cruise performance. 

4. A combination of trailing edge extension flaps and symmetric ailero:i 

deflection provided good pitch corttrol past the stall angle of aLmk. 

5. Rudders on forward swept tzils provided the siniplest method of 

obtaining high angle-of-attack yaw control. 

could be possible to achieve the same amount of yaw control with an aft swept 

tail rudder plus a tip rudder. 

Results also indicate that it 

6. Deflectable iring tips, which taie advantage of wing spanwise flow, were 

found to provide good rn!l control when used in conjunction with ailerons. 

7. Static stability levels indicated that the forward swept tail configuration shoul d be highly departure resi stant. 

8. It was shown that proper aileron-rudder interconnect gains cauld postpone roll control reversal fi*$m 51 = 20" t:, = 45" and preLent: a control induced 

departure. 

- 20 - 
10. Based on attainable rate accelerations and taking into account limitations resulting from inertia coup1 iq, the rnaneuverabil i ty of the forward swept 

tail configuration was close to that of a highly maneuverable aircraft below the 

stall and exceeded that of the current aircraft in the post-stall region. 

- 21 - 
REFERENCES 

1. Klein, J. R.; Walck, K. J.; and Hahne, 0. E.: "Airframe Component Effects 

on the Aerodynamic Stability and Control Characteristics of a 

Supersonic Cruise Fighter Aircraft at High Angles of Attack". AIAA84-2110, August 1984. s 

2. 

3. 

4. 

5. 

6. 

7. 

8. 

9. 

Grafton, Sue B.; 3 nd Nguyen, Luat T.: "Wind Tunnel Free-flight Investigation of a Model of a Cranked-Arrow-Uing Fighter Configuration". NASA TP-2.110, March 1985. 

Uood, R. M.; Miller, 0. S.; Hakne, D. E.; Eleidling, L.; and Klein, J.: 

"Status Review of a Supersonically Biased Fighter Wing-Design Study". AIAA-83-1857, July 1983. 

cbhnson, J. L.; Grafton, S. B.; and Yip, L. P.: "Exploratory Investigation 

of the Efiects of VO rtex Bursting on High- Angle-of- Attack lateralairectional Stability Characteristics of Highly-Swept Win5s". AIAA83-0.163, March 1980. 

Moul, Martin T.; and Paulson, 33 hr; W.: 'Dynamic Lateral Behavior of High 

Greer, H. Douglas: "Summary of Directional Divergence Characteristics of 

Performance Aircraft". 

Several High Performance Aircraft Configurat<ons". NASA TN D-6993, 

November 1972. 

N4CA RM-L58€16, August 1958. 

Etkin, Bernard: ''Dynamics of Atmospheric Flight"; b hn Wiley 8~ Sons, Inc.; 

Nguyen, Luat T.; Gilbert, blilliarn P.; 3 nd Ogburn, Marilyn E.: 

New York, Mew York, 1972. 

System Techniques for Improved Departure/Spi n Resistance for Fighter Ai rcraft" . NASA TP-1689, August 1380. 

Past %all Maneuverability", Journal of Aircraft, vbl. ZG, No. 7, 

July 1983. 

"ControlRansom, Stephen: "Configuration Development of a Research 4i rcraft Hi th 

, 

10. Nguyen, L. T.; Qgburn, M. E.; Gilbert, W. P.; Kibler, K. 5.; &om, P. W.; 

3 nd Deal, P. L.: "Simulator Study of Stalf/Post-Stall Characteristics 

of a Fighter Airplane With Relaxed Longitudinal Static Stability", 

NASA TP-1538, December 1979. 

- 22 - 
Table I Mass and geometric characteristics of the airplane 

Weight(Estimated): 

TOGW. 1 bs ...................................................... 44. 000 

Moments of Itxrtia (Estimated): 

‘x. Slugs . tt ................................................. 19312 

Iy. sluys . ft2 ................................................ 163899 

Iz. slugs . ft ................................................ 176171 

Overall Fuselage Length. ft ........................................ 66.67 

Wing: 

Span. ft ......................................................... 36.7 

Area. ft ........................................................ 691.9 

Mean aerodynamic chord. ft ....................................... 23.1 

Aspect ratio ..................................................... 1.95 

Leading edge sweep. deg ............................................ 65 

Swept Ai 1 eron: 

Area (one side). ft2 .......................................... 15.3 

Unswept A! 1 eron: 

Area (one side). ft2 .......................................... 10.1 

Deflectable Wingtip: 

Area (one side). ft2,., ........................................ 7.6 

Aft Swept Vertical Tail: 

Area (each). ft2 ................................................. 79.6 

Span. ft .......................................................... 9.2 

Root chord. ft ................................................... 14.0 

Tip chord. ft ..................................................... 3.2 

Aspect ratio ..................................................... 1.06 

Leadjng edge sweep. deg .......................................... 62.8 

Conventional ?udder area. ft2 ..................................... 7.6 

Tip rudder area. ft2.+.......................................... 11.7 

Forward Swept Vertical Tai 1: 

Area (each). ft2 ................................................. 83.7 

Span. ft ........................................................ 10.4 

Roo i cho-d. ft .................................................. 11.2 

Tip chord. ft ..................................................... 4.8 

Aspect ratjo ..................................................... 1.29 

Leading edge sweep. dey ........................................... -30 

Rudde” areas ft2 ................................................. 18.4 

. 23 . 
Table 1 Concluded 

Horizontal Tail fall moving): 

Area (each). ft2. ............................................... 47.67 

Span. ft ......................................................... 8.67 

Leading edge sweep. deg .......................................... 42.5 

Aspect ratio ..................................................... 1.58 

Trailing Edge Extension: 

Length. ft ....................................................... 15.7 

Width. ft ......................................................... 4.7 

Flap area (each). ft2 ............................................ 23.0 

Canard: 

Area. ft ......................................................... 28.8 

Span. ft .......................................................... 5.98 

Root chorJ. ft .................................................... 7.63 

Tip chord. ft ..................................................... 2.08 

Aspect ratio ...................................................... 1.24 

Leading edge weep. deg ............................................. 50 

. 24 . 
Maneuverability 

Current 

Future -..-..e-- 

Angle of attack 

Figure 1. Effect of improved control power on caneuverabi 1 i ty. 

WITH INPROVED 

/- CONTROL POiJER 

Figure 2. Effect of improved control power on confisuration 

o j t i i,i i z a t i o n . 
cu 

I 

I 

n c 

N 

0 

n 

I 

F 

i 

I 

I 
M 

h 

e' 

u\* I . 

h 

L 
Centerline 

All dimensions in inches 

Figure 5. Vertical tail geometry. 
wing tip 

Figure 6. Longitudinal and lateral control surfaces. 
v) 

c 

0 

4.J 

.C- 

.C 

VI 

0 

a. 

v 
. c 

Y 

c 

c 

3 

c 

c-) 

0 

0 

LL 

N 

4 

c 

.c 

c 

a 

Q 

0 

E 

O 

Y- 
W 
m C 

No tail 

Celiterline tail 

Twin aft swept tail 
C 

C 

"P 

0 

0 No tail 

0 Centerline tail 

0 Twin aft swept tail 

-10 0 10 20 3Q 4Q 50 60 

sf, deg 

Figure 11. Effect of vertical tail geometry on ldterdl-directiondl 

characteristics. 
0 

a C 

c 

Tail sweep Position 

Aft 

Forward Forward 

Forward Aft 

- 

Figure 12. Effect of vertical tail sweep on longitudinal characteristics. 
Tail weep Position 

0 Aft 

0 Forward Fowa rd 

0 Forward Aft 

- 

0 

B 

0 

-10 0 10 20 30 40 50 . 60 

Figure 13. Effect of vertical tail sweep on lateral-directional characleristics. 
0 

E 

CJ 

n -= VI \o 

N ** 

00 

+-- 

I_ 

c 

rg 

c 

U 

ZJ 

tr 

3l 

E 

0 

.C 

.C 

c 
C 

C 

"6 

-10 0 ' 10 20 30 40 50 60 

Figure 15. Effect of vertical tail spanwise location I;n ldtecal-diceczional 

characteristics. 
0 O0 

0 15O IS 

b 30" iB 

-10 0 10 20 30 40 50 60 

a, del 

Figure 16. Effect of vertical tail cant angle on longitudjnal characteristics, 
Cant angle 

0 30° OB 

0 15O OB 

0 O0 

A Eo I5 

b 3oo IF3 

Q C 

yB 

C 

"P 

P 

-10 0 10 20 30 40 50 60 

Figure 17. Effect of vertical tail cant angle on lateral-d<rectionat 

characteristics. 
m c 

Q 

bf B OB 

0 oo oo oo 

D 30" Qo Oo 

0 300 30" o* 

n 30° 3oo 30° 

Figure 18. Effect of leading edge f?dpS on longitudinal characteristics. 
0 

F 
C 

C 

% 

0 

0 

-18 0 10 20 30 40 50 60 

a, d@g 

b. Lateral-dirt?ctional characteristics 

Figure 19, - Concl uded 
rn C 

0 ~ 

0 

-10 

Figure 20, Tra 

LJ-J I I I 

10 20 30 40 50 60 

6, deg 

a. Slotted flap 

1 ing edge extension f ap effectiveness. 
m c 

3 

-10 

6F 

0 -30' 

a oo 

0 3oo 

/ 

b. Plain flap 

Figure 20.- Concluded 
cL 

0 

0 -20" -30' 

0 oo GO 

0 30° 0" 

Q0 30° 

b 30° 30° 

_i I i I i J 

-10 0 10 20 30 40 50 60 

Q, deg 

Fiyure 21. Effectiveness of ai 1 erons for pi tch cont-01 . 
0 

-10 0 10 20 33 40 50 60 

Figure 22. t!oriion:d! tail eftectiveness. 
m c 0 F 

Ct 

Figure 23, 

-10 

Effect 0 

Canard fairing 

0 Cff 

0 On 

y--Y--- 

0 14) 20 30 40 $13 ti0 

a, dq 

f canard fdi-iny on longitudinal chardcte-ist ics. 
C 

yB 

0 

C 

"B 

Canard fairing 

0 off 

0 On 

.e- I 

Figure 24. Effect of canard fairing on lateral-directional characteristics. 
, 

ax 

c 

V 

. 

M 

V 

a 

M 

L 

w 

+J 

U 

m 

L 

m 

r 

U 

.c 

.C 

c 

rg 

c 

0 

c, 

V 

0 

L 

73 

I 

*P 

c 

aJ 

+J 

a 

C 

0 

C 

0 

c, 

V 

aJ 

ual 

U 

3 

L 

a 

c 

a 

U 

rc. 

0 

+J 

V 

aJ 

uQw 

.C 

.r 

c 

c 

.C 

c 

rc) 

hl 

CU 

L 

J 

3 .? 

LL 
C 6 

0 oo 

0 -4Q0 

Figure 26, Canard effectiveness. 
0 

0 -loo 

0 

LI 

ri 

Figure 27. Center1 irle vertical tai 1 effectiveness. 

LI 

ri 0 ij = -30° 

‘f 

0 

c: del 

Figure 23. Tip rudder effectiveneses. 
Q 

C 

u 

c 
0 

LI 

4c" 0 

Figure 31. Effectiveness of wing tip deflection for roll control. 
. 

c 

C 

44 

'FI 

c 

3 

7, 

C 

0 

V 

44 

w 

Y 

..J 

v- 

.C 

.C 

YD 

QJ 

U 

C 

m 

d.4 

m 

m 

QI 

c 

aA 

L 

3 

44 

c 

'4 

.C 

i? 

. (v 

m 

Y 

L 

3 

39 

LL 

.C 
3 

$1 

Q 

e 

n 

0, 

.?- 

.la. 

n 

3 

<=\A- \ 

a, dq 

Figure 35. Effect of inertid coupling on maximum -011 rate.