~N AILERON DESIGN TO COUNTER

ADVERSE YAW

By

GARY LEE ~~REPS H

Bachelor of Science

Oklahoma State University

Stillwater, Oklahoma

1970

Submitted to the Faculty of the Graduate College of the

Oklahoma State University in partial fullfillment of

the requirements for the Degree of

MASTER OF SCIENCE July, 1985

C 0 P Y R I G H T

by

Gary L. Kreps

May 31~ 1985

1'hes\1 JC1~b ~q ~o.J o o~? • ?-...

AN AILERON DESIGN TO COUNTER --ADVERSE YAW

Thesis ~~PPI'"OV~?d:

··~ ........... -.. --· ..... -· ..... .£..... ... ct:Z. .... -J ..... - ........ -· ·-·----··· ....... -·- ........ -· -· -- .... -..... . a)~ tJ ,

-----------------~·-------------

-.......... -..... -.... /2~ ...... L.2=-~·-····-.. -·-·-- .. ·--·-· Dean of the Graduate College

ii

122~9UB ,

PREFACE

The thesis project was chosen to represent an

original idea to avoid being an assistant on someone"s

research project. This research explores an idea

developed several years ago but with no suitable

opportunity to test. The fact that the idea worked under

test has been very satisfying because the idea can be

justified completely from the standpoint of aviation

safety. An interest in aviation and aeronautical

engineering made the subject of this research an

appropriate choice for a thesis.

The design proposed is intended for light general

aviation aircraft flown by pilots with low experience. The

basic principle of this design is to add parasite drag to

replace lest induced drag on the aircraft wing when

ailerons are displaced. Originally it was predicted that

the design would have to be optimized for a narrow speed

range. The wind tunnel testing showed that the idea works

and can be optimized over a large speed range due to a self

modulating effect of the drag. The lift and drag

characteristics of the wind tunnel model tested showed that

it has promiae as a glide path control device also. The

design has merit over others used to counter adverse yaw

because it is basically a control surface replacement and

iii

could be retrofitted to existing aircraft.

Numerous difficulties were encountered along the way

and the success of the project is due in large part to the

assistance of others. Special credit must be given to my

major adviser, Dr. M. L. Millett, Jr. Dr. Millett is a

full time employee of Boeing Military Airplane Co. of

Wichita, Kansas. He teaches an aircraft design course at

Oklahoma State University in the spring. The time demands

of commuting so far to teach a course indicate tremendous

dedication to education. Agreeing to be a major adviser

for a graduate student placed even greater demands on his

time. Having taught aeronautical engineering at Iowa State

University for twenty five years, in addition to working in

industry, gives Dr. Millett the highest credentials for

serving as a thesis adviser on an aeronautical research

project.

My other committee members, Dr. Robert L. Swaim and

Dr. Flint 0. Thomas provided advice, guidance, and

encouragement.

Recognition must be given to Wichita State University

for making the low speed wind tunnel available at a very

reasonable cost.

Marvin Davidson and Hugh Crane of the Walter H. Beech

Memorial Wind Tunnel gave valuable advice and assistance in

carrying out the experimental phase.

iv

TABLE OF CONTENTS

Chapter- F'iage

I. INTRODUCTION . . . .. . :L

II. DEVELOPMENT OF THEORY :L '? . "-

III. OTHER METHODS OF COUNTERING ADVERSE YAW • • 17

IV. WIND TUNNEL TESTING

v.

VI.

VI I.

RESULTS OF TESTING . • .

DISCUSSION OF RESULTS

CONCLUSIONS AND RECOMMENDATIONS FOR FURTHER STUDY • • •

A SELECTED BIBLIOGRAPHY

APPENDIXES

APPENDIX A - MODEL CONSTRUCTION

21

• 29

• .l~ 1

• • • 47

• 49

. . .. . 51

APPENDIX B - WIND TUNNEL NUMERICAL DATA ... 63

APPENDIX C - WIND TUNNEL COEFFICIENT PLOTS ... 89

v

LIST OF TABLES

Tiabl e Page

I. Wind Tunnel Testing Sequence . . . . . . " . .

vi

LIST OF FIGURES

Fi gur·e

1.

4.

5.

6.

7.

8.

9.

10.

11.

12.

14.

15.

Forces Acting on an Aircraft in Level Flight and in a Turn .••

Lift and Drag Forces on an Aircraft with Ailerons Displaced for a Turn to the Right •.•••

Drag of an Aircraft for Low Subsonic Velocities • • •

Aileron Profiles . . . . . . . . . Drag Cur·ves at -6 Degr·ees Angle of Attack

Drag Curves at -4 Degrees Angle of Attack

Drag Curves at 0 Degrees Angle of Attack

Dr·ag Curves at +4 Degrees Angle of Attack

Drag Curves at +8 Degrees Angle of Attack

Drag Curves at +12 Degrees Angle of Attack

Ikag Curves at ·t-15 Degrees Angle of Attack

Lift/Dr·ag Plots •••

Rib Pr·ofiling Fb:ture • •

Rib Sections

Tunnel Mounting Fixture •

16. Milled Aileron Components ••.

17. Aileron Indexing Brackets . . . . 18. Rigging Structure •••• • •

19. Complete Structure Prior to Covering

20 .. Tail Sting Components ••• . . . . . .

vii

. .

.

. .

. .

. .

•

.

.

.

Page

c· ,.J

7

8

• 15

. . 3:~;

. . ~~4

. . 35

. . 36

. . 37

. . :~;8

. . ~59

• • • 40

. 55

. 55

. . . 57

. 57

. . • ~'59

.. . t39

. 60

• 60

Fiqure P1-aqe

21 ... ·rest Model Installed in Wind Tunnel with Ai 1 erc.1m; in Nt=Jutral Posi ti 1:1r1s . 6'2

2:;:~. Ma:·:imum Split of Ai lercms . . 6'"·' ..:..

:;:~::~. RLtn 1. Cl VS Alpha 89

24. Run 1. Cd vs Alpha . . 90

1""\t:. .. Run 1. Cm vs Alpha 91 ,£,, .. J. . . . . . 26. Run 1. Crm vs Alpha. . . . . 92

';.:!7. Run 2. Cl VS Alpha . . . . 93

28. Run 2. Cd vs Alpha. . . . . . 94

29. Run . ., Cm vs Alpha 95 ...... . . . . . . 3(). Run ··~ ..::.. Crm vs Alpha 96

31. Run 3 .. Cl vs Alpha . 97

~5~!. Run 3. Cd VS Alpha . . 9E3

:~;3. Run 3 .. Cm VS Alpha . . 9''Y

:34-. Run ::; It Crm VS Alpha . . . 100

....... Run 4 • Cl vs Alpha. 101 . .::.,J .. . 36. RLtn 4. Cd vs Alpha . . . 102

~$7. Run 4. Cm vs Alpha . to:~:

~$8. F~Ltn 4. Crm vs ?Upha 104

:~9. Run a::. .. Cl vs Alpha 105 '-~• . 40. Run C'

...J. Cd vs Alpha . . . 106

4l.. Run .:::· Cm vs Alpha 107 ....~. . 42. Run 5. Crm vs Alpha 108

4~$. Run 6. Cl vs Alpha. . 109

44. Run 6. Cd vs Alpha . 11 (l

4""' ...Jo Run 6. Cm vs Alpha. . . . 1 1 1

46. Run 6. Cr·m vs Alpha :l12

viii

Figure p,'::\qe

47. Hun 7. Cl vs Alpha . . . . . . . . . 11:3

48. F~un 7. Cd VS Alpha . . . . . . . 114

49. Run 7. Cm vs /Upha . . . . . . . . . . . . . 115

50. Run 7. Cr·m vs Alpha . . . . . . . . . . 116

51. Run 8. Cl vs Alpha . . . . . . . . . . . . . 117

5:2. Run 8. Cd VS Alpha . . . . . 118

5~5. Run 8. Cm VS Alpha . . . . . • . . . . . . . . . 119

54. Run 8. Crm VS Alpha . . . . . . . . . . . . . . 120

55. Run 9. Cl vs Alpha . . . . . . . . . . . 121

56. Run 9. Cd vs Alpha . . . . . . . . . . . 122

57. Run 9. Cm vs Alpha . . . . . . . 1 ::~~..)

58. Run 9. Crm vs Alpha . . . . . . . . . . . . 124

59. Run 10. Cl VS f'-ll. ph a . . . . 1 r"\1::: ..::.'-'

60. Run 10. Cd vs Alpha . . . . . . . . . . . . 126

61. Run 10. Cm vs Alpha . . . . . . 127

6''~ ..:: .. Run 10. Crm vs Alpha . . . . . . . . 128

6~:;. Run 11. Cl vs Alpha . . . . . . 129

64. Run 11. Cd VS Alpha . . . . . . . . 1:::;o

6""" J. Run 11. Cm VS Alpha . . . . . . . . 131

66. Run 11. Crm vs Alpha . . . . . . . . 132

67. Run 12. Cl vs Alpha . . . . . . 133

68. Run 12. Cd vs Alpha . . . . . . . . . . . . 134

6CJ • R1 . .1n 1:2. Cm vs {Upha . . . . . . . . . . . . 135

70. R1..m 1'"·' ...... Crm vs Alpha. . . . . . . . . . . 136

71. Run 13. Cl VS Alpha . . . . . . . . . . 137

.. ,2. R1..1n 13 • C!::l vs Alpha. . • . . . . . . 1~.::;8

Fi gun? P;age

T~:. Run 13. Cm vs Alpha . . . . . . 1:39

74. F~un 1''~ ..:: ... Cr·ni vs Alpha . . . . . . . . . . . . :1. /.f.(l

7~i. Run 14. C:l vs Alpha . . . . . . . . . . . . . . 141

76. Run 14. Cd vs Alpha . . . . . . . . . . . . 142

'77. Rt.tn 14. Cm vs Alpha . . . . . . . . . . . . 143

7€-3. Run 14. Crm VS Alpha . . . . . . . . . . 144

79. Run 15. Cl vs Alpha . . . . . . . . . . 145

80. Run 15. Cd vs Alpha . . . . 146

81. Run 15. Cm vs Alpha • . . . . . . . • . . . . . 147

El2. Run 15. C:r·m vs Alpha . . . . . . . . 148

8 .... .,.::. Ill Run 16. Cl vs Alpha . . . . . . . • . . . . . . 149

84. Run 16. Cd vs Alpha . . . . . . . . . . . . 150

8""' ;.J. Run 16. Cm vs Alpha . . . . . . . . 151

86. Run 16. Crm VS Alpha . . . . . . . . . . 152

87. Run 17. Cl vs Alpha . . . . . . . . 1 0::'':!' ;;J~·

88. Run 17. Cd vs Alpha . . . . . . . . . . . . 154

89. Run 17. Cm vs Alpha . 0 . . . . . . . . . . 155

90. Run 17. Crm vs Alpha . . . . . . . . . . . . 156

91. Run 1.8. Cl vs Alpha . . . . . . . . . . . . 157

9~:.• ~ .. Run 18. Cd vs fUpha . . . . . . 158

(t~: .. RLln 18o Cm vs Alpha . . . . . . . . . . 159

94. 1::;:un 1EL Cnn VS Alpha . . . . . . . . . . . . 160

9C:' ,:J • Run 19. Cl VS Alpha . . . 0 . . . . . . . . 161

96. F~un 19. Cd vs Alpha . . . . . . . 0 . . . . 1.62

97. Rt..m 19. Cm vs Alpha . . . . 0 . . . . . . . 163

98. Run 1 (~ 0 Crm vs Alpha . • . . . . • . 164

Figure

99. Run 2Cl.

100. Run 2().

101. Run 20.

102. F~un 20.

Cl vs Alpha

Cd vs iUpha

Cm vs Alpha

Crm vs Alpha

. .

.

. . . . . . . . .

. . . . . . .

Page

165

166

167

• • • 168

NOMENCLATURE

Alpha. Angle of attack

AR Aspect ratio

b Wing span

Cd Coefficient of drag

Cd. C1::>eff i c i f:H1t of induced drag

Cd Coefficient of par·asi te drag

Cl Coefficient of lift

Cm Coefficient of moment

D drag

e Oswald wing efficiency factor

f Equivalent parasite area

L Lift

L/D Lift to drag ratio

q Dynamic: pressure

S Reference area

V Airspeed

W Weight

_,P Air densi"t:y

CHAPTEF~ I

INTRODUCTION

Adverse yaw is a phenomenon that occurs when ailerons

are deflected to turn an airplane in one direction while

the nose of the aircraft tends to yaw in the opposite

It is an undesirable control response and must

be countered. Many methods have been developed but all

have shortcomings. The classical solution requires the

pilot to overcome the adverse yaw by appropriate rudder

deflections. The objective of this research is to examine

a means of eliminating the adverse yaw characteristic from

Eliminating adverse yaw would make aircraft control an

easier task for the pilot. However there are more

compelling reasons for eliminating this undesirable control

char ;;act er· i s t i c. One of the lesser reasons is that failure

to counter adverse yaw results in uncoordinated flight and

unnatural inertial forces which cause discomfort to

Enhancement of flying safety alone l. "" • :::> adequii~te tc:>

justify determined efforts to eliminate adverse yaw.

The stall/spin accident still remains one of the leading

types of aviation accident. The pilot inadvertently stalls

1

2

the aircraft and fails to recover before striking the

ground. In some cases the stall progresses to a spin from

which recovery is difficult or impossible. Excessive

adverse yaw may result in a spin while the plane is in a

stalled condition. Many aircraft roll when entering a

stall. Efforts by the pilot to recover level flight by

large aileron inputs without coordinated rudder can

aggravate the situation and lead to the spin. As will be

explained later, an aircraft entering a stall is in the set

of conditions which result in the greatest adverse yaw and,

at the same time, the pilot is least able to counter it by

rudder inputs. Intentional spins performed during flight

training are usually entered by a different technique.

However this method of spin entry is easy to demonstrate.

Eliminating adverse yaw from an aircraft•s natural behavior

will enhance safety.

Another safety problem involves the common situation

of a pilot without instrument flight skills flying into

poor weather conditions and losing visual reference. When

a person in an aircraft does not have visual reference his

sense of orientation is totally unreliable. The human

sense of balance functions properly only when in a

stationary reference frame. The inertial accelerations

of an aircraft in flight result in incorrect attitude

sensations. Disorientation on the part of a pilot is

referred to as vertigo. If ihe pilot is disoriented he

will make incorrect control inputs in an effort to maintain

-~· ·-·

level flight. If the pilot is not trained for instrument

flight~ he may easily panic and make incorrect decisions.

The typical accident sequence is one in which the

noninstrument pilot inadvertently gets caught in bad

weather, becomes disoriented, loses control of the

aircraft, and crashes. Such accidents are usually fatal.

Uncoordinated flight resulting from adverse yaw is a prime

contributor to vertigo. Th~ pilot would have a much better

chance of maintaining level flight after losing visual

reference if his aircraft was incapable of producing

The elimination of adverse yaw would therefore be

desirable from a safety standpoint. A stall proc:Jf

aircraft without adverse yaw would be a major improvement

ever most current general aviation aircraft.

Before being able to counter adverse yaw it is

necessary to understand the mechanism that causes it.

Quantitative analysis requires the use of rather complex

·f 1 ui d t.he!Clr" i es. Qualitative study of the mechanism of

adverse yaw can be accomplished by a much more straight

forward process using physical laws.

An aircraft is supported in level flight by

symmetrical distribution of lift en the wings. It i !S

necessary to bank the aircraft to effect a coordinated

"l:l .. II'"Tl • <Coordinated flight is flight in which all inertial

accelerations are perpendicular to the floor of the

iain::r·<aft.) The bank tilts the lift vector in the desired

direction of turn. The corresponding horizontal component

of the lift force provides the horizontal acceleration

necessary to accomplish the turn. This is illustrated in

Figure 1. Some means are necessary to cause the aircraft

to roll about the lateral axis. This is done by altering

the lift distribution of the wing to a nonsymmetrical

state. This creates unbalanced moments about the

longitudinal axis and the aircraft rotates about that axis

until symmetry is restored. The Wright brothers initially

4

accomplished this by twisting the wings in flight. Warping

the wing structure had many disadvantages. The primary

disadvantage was the difficulty of building adequate

flexibility for in flight wing warping and at the same time

obtaining adequate structural strength and resistance to

warping under load. At the higher speeds at which aircraft

now operate, wing flutter would be a serious problem.

Moveable surfaces <ailerons) were developed to alter the

lifting characteristics of the wing. The movement of the

ailerons is a differential type of motion to increase lift

on one wing while decreasing lift on the other. The

ailerons are displaced in opposite directions from their

neutral position.

An aircraft in flight experiences drag forces

resulting from its movement through the atmosphere. For

steady flight to be maintained the drag forces must be

maintained in a state of symmetry. Unfortunately, when

symmetrical lift is disturbed to roll the aircraft the

Lift equal to weight

...

Vertical lift

0

Lift

b

Weight

..--- Horizontal force required for turn

equal to weight:;__ __ Lift ---=-=

Weight

Figure 1. Forces Acting on an Aircraft in Level Flight and in a turn

5

6

symmetrical drag is also disturbed opposite to the desired

direction of turn. This is the mechanism of adverse yaw.

Figure 2 illustrates the lift and drag forces for an

aircraft with the controls displaced to cause a roll to the

right. Maintaining drag symmetry during a roll would

eliminate adverse yaw.

Drag experienced by an aircraft is categorized into

parasite drag and induced drag. Parasite drag is the drag

that results from resistance to motion through the

atmosphere and increases with increasing speed. Induced

drag is drag caused by the production of lift. Induced

drag is high at low speed and low at high speed. It is not

intuitively obvious that induced drag should be inversely

proportional to velocity. Further explanation of the

functional relationship of drag and velocity is provided in

chapter two. Figure 3 illustrates the parasite~ induced~

and total drag of an aircraft as a function of velocity.

Since induced drag plays a key role in the mechanism of

adverse yaw it can be seen that adverse yaw will be more

pronounced under conditions of high lift and slow speed~

the conditions that create high induced drag. This is why

an aircraft approaching a stall is in a condition where

adverse yaw forces are high. Since dynamic pressures are

low at the low speed associated with the stall the aircraft

control surfaces have the least capability to counter

adverse yaw when the adverse yaw forces are greatest.

Induced drag is a function of lift on the wing.

Lift

c,

Adverse yaw moment

i d

Rolling moment

I 0 Lift

$ t b

Desired direction

of turn

Figure 2. Lift and Drag Forces on an Aircraft with Ailerons Displaced for a Turn to the Right

7

Total Drag:: I -+ \Jt

Total drag ___ ____,

Velocity

Induced drag

Figure J. Drag of an Aircraft for Low Subsonic Velocities

8

9

Therefore any change in lift will result in a change in the

induced drag of the wing. The differential aileron action

effects the roll necessary for a turn by modifying the lift

distribution of the wing and the drag distribution is

altered as a consequence. Unfortunately the drag

distribution is altered in such a fashion that it is not

symmetrical and the aircraft tends to yaw. If the ~aw were

in the desired direction of turn it would give a favorable

yaw roll coupling. Unfortunately it is in the opposite

direction and opposes the turn. This is the situation

illustrated in Figure 2. In some situations it can be the

dominant factor and result in the aircraft turning in the

opposite direction. Ultralight aircraft sometimes

experience this condition and they must be turned by the

use of wingtip drag devices instead of ailerons. The

man powered aircraft recently developed by Paul McCready

experienced this difficulty (6). Being able to turn was

one of the greatest challenges to the development of the

man powered aircraft. The early experimental aircraft

built by the Wright brothers also proved difficult to turn

because of the effects of adverse yaw. The problem was so

great that the aircraft would turn in the direction

opposite of the attempted turn du• to an unfavorable

yaw-roll coupling. It is a mark of their insight and

genius that they were able to identify the cause of the

problem and provide a solution <4>.

Aircraft of the size generally used by individuals for

10

private flying are not dominated by adverse yaw and can

provide satisfactory control by the use of ailerons.

However the adverse yaw is still present and a significant

consideration for aircraft flown by novice pilots.

Adverse yaw is not considered a problem for large

aircraft, such as those used for commercial operations.

Large aircraft generally fly at higher speeds where the

higher dynamic pressures on the stabilizing surfaces give

mere resistance to yaw. These aircraft are typically flown

by highly experienced and well trained pilots who are not

likely to have difficulty controlling the aircraft when

adverse yaw forces exist. Large aircraft have autopilots

and control augmentation systems capable of overcoming

undesirable control characteristics. However, designers of

commercial aircraft built for V/STOL operations must still

be concerned about the problems of adverse yaw.

The result is that efforts to eliminate adverse yaw

from an aircraft,s control response are directed at lower

speed aircraft piloted by less skilled pilots, i.e., light

general aviation aircraft. The economics of these aircraft

preclude the use of automated control augmentation systems

such as usmd on large aircraft.

An effective and economical method of eliminating

adverse yaw will improve the safety of light aircraft being

operated by novica pilots. The design developed in this

1 1

research is intended to meet this objective. As will be

explained in the following chapters, this design shows the

potential to meet the objective of improving flight safety

through improved aircraft control.

CHAPTEI:~ I I

DEVELOPMENT OF THEORY

Classical aerodynamic analysis methods employ

the use of coefficients for force measurement. The

coefficient is a nondimensional number chosen to fit the

following equation:

F ::::: C q S

F is the force being measured.

C is the coefficient for the force.

q is the dynamic pressure of the airstream.

S is a reference area.

The.type of aircraft considered for this design proposal

fly at low subsonic speeds where compressibility is not a

·factor. For incompressible flow dynamic pressure can be a

c.::d. cul.c::~ted by q == .f V /2 when~ f is the air· density and V

is the speed of the undisturbed airstream relative to the

For drag measurements the equation ~ecomes

Dr·.-ag ::::: Cd q S

where Cd is the coefficient of drag. For lift measurements

the equation becomes

Lift :::: Cl q S

where Cl is the coefficient of lift.

1 ') -

A similar system can be used fer moments by the addition of

a length term to the equation.

Moment = Cm q S 1

where Cm is the coefficient of moment and 1 is a reference

1 f:i!ngth"

Drag is usually analyzed iR two categories. Induced

drag is the drag resulting from the production of lift.

Parasite drag includes all other sources of drag.

Parasite drag is proportional to the square of

velocity and can be expressed as

Parasite Drag = Cdr q S

where Cdf is the coefficient of parasite drag.

desirable to express induced drag in the same way.

Induced Dra~J •= Ccij q S

where C~ is the coefficient of induced drag.

Whereas Cc~remains essentially constant for different

angles of attack~ C~ is totally a function of angle

As indicated in an earlier discussion induced

drag is inversely proportional to velocity. Cd; is f.~

function of angle of attack and angle of attack for an

aircraft in flight is related inversely to the velocity.

Classical aerodynamic theory shows that

Cda = Cl1 I 'tr' AR e

where AR is the wing aspect ratio and e is as an efficiency

factor to relate lift distribution to the

ideal elliptic lift distribution. ·rhen:.~f c:we cis;; Cl

increases the induced drag coefficient increases

14

f?.:·: ponent i ;ally. Since the function of an aileron is

to change the Cl of the wing actuation of the aileron will

change the induced drag of the wing. This will have the

greatest effect when Cl is high, or in other words at high

angle of attack or low speed~.

Shevell <12> provides a good development of the lift

and drag relationships. The total drag <parasite plus

i nduc:f:~d) is given by the following equation:

D - f.R_ v~ + ~··~-e.-···(\;) I ~ -~~·

when:? ·f:::: eqL.ti valent parasite area

L== lift

IJ::!! wing span.

Collecting constant terms together the equation becomes:

o == c, v2 + c. ; va

<an cl

This is the equation illustrated in Figure 3.

The principle of this proposal is an aileron design

intended to counter adverse yaw. For conventional aileron

design one aileron is displaced downward to roll the

Lift is increased and the total drag increases

due to increased induced

The other aileron is displaced upward to

reduce lift and the total drag decreases. Th:L s unbalances

the drag symmetry of the wing and results in a yaw moment.

Figure 2 shows the effect for a roll to the right. The

figure also shows that the yaw is to the left, opposite or

adverse to the desired direction of turn.

Another type of aileron is proposed to overcome this

result of aileron displacement. It is designed to have an

upper and lower surface that can be moved independently.

The two 5urfaces move together when displaced downward.

This action is identical to that of a conventional aileron.

When moved upward, however, only the upper surface is

displaced, leaving the lower surface in the neutral

position. This results in a split action and the aileron

is referred to as a split aileron. Figure 4 illustrates

tht:~ conc:E.'~pt.

Up

Cc)nvf:?nt i onal Df?!;i gn

c--·-~

c

Figur··e 4.

Split Design

.... c: ___ ~

c

Aileron Profiles.

16

The split aileron is used on the wing on which

the aileron is displaced upward. The deflection of

the split aileron increases the parasite drag to balance

the loss of induced drag and the increase of drag on the

opposite wing. Since induced drag decreases as the square

of velocity and parasite drag increases as the square of

velocity it is expected that the design must be optimized

for a specific velocity. This situation is typical of

design activity. Wind tunnel testing of this design has

shown that this split aileron design can actually be

optimized over a much broader range than might be expected.

This is explained in Chapter Six.

CHAPTER III

OTHER METHODS OF COUNTERING ADVERSE YAW

Numerous ether methods have been used in the past to

deal with adverse yaw. The purpose here is to review other

methods and point out their deficiencies, thereby

substantiating the need for an improved design.

The most common method has been to train the pilot to

input appropriate rudder displacements at the same time

ailerons are displaced. This can work fairly well if the

pilot is sufficiently skilled. It has the obvious

disadvantage of increasing pilot workload. It does not

work if the pilot fails to make appropriate inputs. A

pilot with low experience cannot be depended upon at times

of high stress to make all appropriate inputs consistently.

It was pointed out in the introduction that a pilot who is

spatially disoriented is at high risk. Another

disadvantage is that some situations require the pilot to

use the rudder for other purposes. The inputs to counter

adverse yaw must be superimposed on the inputs required.

Some aircraft have been designed with the aileron and

rudder controls interconnected in such a way that they move

at the same time to counter adverse yaw. The pilot

displaces the ailerons to affect a turn and the rudder

17

18

moves automatically with the ailerons. While this method

has some distinct advantages it has been poorly received by

pilots. Pilots have been unwilling to give up control of

the rudder because they cannot use it for other purposes.

This design can only be optimized for a narrow range of

speed. The aircraft will be operating off design at any

<:Jther speed. Interconnecting rudder and ailerons

eliminates the capability of making crosswind landings by

conventional methods.

Wing spoilers have been used with some success for

roll control. Ailerons may not be installed on the

aircraft. An upper surface spoiler is displaced to effect

a roll by reducing lift on that wing, causing it to sink.

The spoiler also increases drag on that wing. One

disadvantage is that it alters lift on one wing instead of

both. Another problem is that overall wing lift of the

aircraft is reduced during the roll since there is no

increase of lift on the other wing. The most serious

problem with using spoilers is the highly nonlinear control

characteristics of spoilers. Without automatic control

augmentation systems it is difficult to obtain acceptable

flying qualities using spoilers. For small displacements

the spiJiler gives little control response, but the response

can become very pronounced at higher spoiler angles. This

would give the pilot a "stuck drawE-!r" effect.

One very interesting design can be found on the Helio

aircraft. This aircraft is designed for short field

19

operation and operates under the conditions which give the

greatest adverse yaw forces. The wing has flaps and

leading edge slats which enable it to fly at high Cl or low

airspeeas. The wings have a large span and the ailerons

are on the outer portions of the wing to permit essentially

full span flaps. The yaw forces have a large moment arm to

create yaw moments. The fact that adverse yaw is a problem

on this aircraft is evidenced by the manufacturer adding a

device to counter it. A spoiler type device is deflected

on the upper surface in front of the aileron. This system

may even increase roll rate with positive G loading. This

solution would require structural design to carry the loads

across the cutouts for the spoilers. Since most aircraft

use stressed wing skins for structural integrity these

slots would weaken the wing. It would be difficult to

retrofit existing aircraft since a major wing redesign

would be required. If adequate drag is provided at low

speed it would predictably be excessive at high speed.

This system will not work under negative G loading. The

extra mechanical complexity might also be viewed as a

drawback.

Another method is the use of differential displacement

of the ailerons. The aileron displaced up is displaced

more than the aileron displaced down. This is usually

referred to as differential aileron action. The extra

displacement of the upward aileron is intended to suffer

parasite drag due to the departure from the basic airfoil

shape. This method has proved to be ineffective.

Another method of dealing with adverse yaw is with

automated control systems or yaw dampener systems. The

cost~ weight, and complexity of such systems preclude use

en small aircraft where they would be needed most.

20

CHAPTER IV

WIND TUNNEL TESTING

It was desirable to conduct wind tunnel tests to

carefully measure the split aileron performance compared tc

conventional aileron performance. Since the aileron

functions by modifying the wing airfoil it would at first

seem appropriate to use two dimensional airfoil testing

techniques. Airfoil data is usually determined by two

dimensional testing techniques to eliminate wing tip

effects. The test wing completely spans the space between

two parallel walls oriented perpendicular to the span.

Lift and drag values are determined with pitot static probe

measurements and application of momentum theory. Rae and

Pope <10) point out that this method is invalid far any

flow with separation. An aileron operates in the area of

the wing where same separation exists and the split aileron

concept is intended to cause additional separation.

Therefore two dimensional testing will not provide proper

data.

Direct force measurement in two dimensional testing is

difficult because the wing must touch the end plate walls.

Even very small gaps at the wing tips cause large

measurement errors (10). It is difficult to separate

21

forces acting on the wing from forces acting on the end

plates.

Two dimensional testing of this concept is not

desirable in any case. The reason for the difference

between two dimensional and three dimensional airfoil

performance is the alteration of flow characteristics

caused by the wingtips. The flow in the region of the

22

wingtips is drastically different from two dimensional

behavior and this is the very region in which the ailerons

operate. Additionally each aileron has two wingtips of its

own. Therefore meaningful testing must be done in a three

dimensional mode with proportions similar to the type of

aircraft intended to incorporate this concept.

Lift and drag performance is also sensitive to

Reynolds number CRN). Previous wind tunnel experience has

shewn that results obtained at RN of one million are valid

up to about twenty five million (10>. Since the aircraft

of interest for this design operate at Reynolds numbers

between two and seven million, it is important that the

testing be accomplished at RN of one million or higher.

This rules out the use of small slow speed wind tunnels.

Fer this test the Walter H. Beech Memorial Wind Tunnel

at Wichita State University was chosen. This tunnel has a

7 X 10 foot test section and can operate at Reynolds number

in excess of one million at the test section. This tunnel

is highly automated and computerized to provide corrections

for wind tunnel errors, i.e., blockage, buoyancy, etc. The

data produced by this tunnel of the highest quality and

well suited to the requirements of this test. A complete

description of the test facility can be found in reference

two.

A model simulating a small general aviation aircraft

was constructed to obtain realistic results. A NACA 23015

airfoil section was chosen for the model. Although the

actual airfoil section selected is not relevant to the

purpose of the test, it is important that the airfoil

chosen have no unusual characteristics. This eliminates

the possibility of obtaining unusual test results due to

some unique feature of the airfoil. The NACA 23015 airfoil

is typical of that used on general aviation aircraft.

Although airflow over a fuselage is not normally

influenced by ailerons, a fuselage does cause interference

on the wing flow characteristics and separation at the

fuselage wing junction. Therefore the test model wing was

equipped with a fuselage to simulate the real world

environment. An empennage was not used because it has no

bearing on wing performance and is difficult to simulate

realistically in the wind tunnel. Since the tunnel walls

alter the downwash behind the wing, the tunnel environment

is not the same as the free stream flow case unless the

model is small for the test section. This in turn will

result in a low Reynolds number. Use of an empennage for

wing testing can therefore cause distortion of the results.

Since a wind tunnel is bounded flow there are many

24

limitations to the use of wind tunnel data. Wind tunnel

data must be corrected for many sources of error and these

sources must be minimized by careful model design and

testing procedure. Reference nine gives a good development

for low speed wind tunnel corrections. One of the

advantages of the Walter H. Beech memorial wind tunnel is

the capability to remove the errors and produce corrected

output for the experimenter. This data output is produced

on line by computer and the experimenter can observe

results as the test is in progress.

The experimenter provides the tunnel operator with a

list of correction parameters before the test is started.

These parameters are entered into the computer data

reduction program. The details of data reduction and error

removal can be found in references five and nine. The

parameters calculated for this test are found on the first

page of Appendix B. The details of the test model

ccmstruction can be found in Appendix A.

Ideally the test sequence might be set up with the

wing at a certain angle of attack and then test data could

be taken for all possible control surface configurations.

Since this data is needed for many angle of attack settings

it would require that this sequence be repe~ted for each

angle of attack. Since the wind tunnel must be shut down

to change model configuration this would be very time

consuming.

The actual test was run by setting a specific control

25

surface configuration and than cycling the test model

through all angles of attack of interest. Since the angle

of attack could be altered by remote control with the

tunnel in operation this sequence resulted in considerable

time savings.

Since performance at a negative angle of attack is of

interest for a control surface some testing at negative

angle of attack is required. Also performance in the stall

mode must be tested to determine any unusual behavior that

may occur. A range of minus six degrees to stall

(approximately 18 degrees) was chosen for the test to meet

these objectives. The control surfaces were built for

deflections from fifteen degrees down to eighteen degrees

up in three degree increments. The upward deflection was

available in both normal and split mode.

Run 1 was made with the ailerons in normal position to

obtain basic model characteristics.

both ailerons down fifteen degrees.

Run 2 was made with

Subsequent runs were

made with the ailerons elevated at increments of three

degrees until full deflection of eighteen degrees was

obtained. Run 13 was made with the left aileron full down

and the right aileron full up. This measured the maximum

rolling moment capability in normal mode. On run 14 the

left aileron was remained down but the right aileron was

now split to full open position. This measured maximum

rolling moment capability in split mode operation. Runs 15

through 20 were then made with increments of aplit opening

on both from three degrees to eighteen degrees. This

provided a data base for comparing conventional versus

split mode lift and drag characteristics. Table I shows

the testing sequence.

The test data output is in numerical and graphical

26

form. Appendix B contains the numerical output which shows

force and moment coefficients as a function of angle of

attack. Dynamic pressure, mach number, and effective

Reynolds number are listed for each run. Appendix C

contains computer generated plots for coefficients of lift,

drag, pitching moment, and rolling moment as a function of

angle of attack.

Although care was taken to set angles of attack to

consistent values, i.e., 2,4,6, etc., it can be seen in the

numerical output that the actual angles of attack vary from

one run to another. Angle of attack is one of the

parameters in wind tunnel testing that requires correction.

The values shown in the numerical output are corrected

values.

This creates a slight inconvenience in the use of

the data. Since the data points in the listings are not at

a consistent set of angles of attack interpolations must be

made. Linear interpolation is not suitable because the

plots show substantial curvature at places.

A computer program was written to interpolate using

the cubic spline technique of numerical methods. This is a

classical mathematical approach which gives good results

TABLE I

WIND TUNNEL TESTING SEQUENCE

60 PSF dynamic pressure

RUN NUMBER CONFIGURATION

1 Clean 2 15 down 'T ..... 12 down 4 9 down .:::· ...J 6 down 6 . .,..

·-' down 7 .. :!" ·-· up 8 6 up 9 9 up

10 12 up 11 15 up 1 ~,

"'- 18 up 13 Left 15 down

Right 18 up 14 Left 15 clown

Right 18 split 15 .. :r ·-· split 16 6 split 17 9 split 18 12 split 19 15 split 20 18 split

Note: On runs 1-12 and 15-20, both ailerons have identical deflections.

28

when the explicit functional form for a data set is not

known. The interpolation program was used to produce

coefficients for nonvarying angles of attack. Due to

length and complexity the computer routine is not published

with this thesis.

author.

It is, however, available from the

The interpolated data sets are used primarily for

making comparisons of multiple test runs. The graphical

plots are used for qualitative interpretations on specific

runs.

CHAPTER V

RESULTS OF TESTING

The wind tunnel tests produced some very useful data.

Examination of the graphical coefficient plots <found in

the appendices) shows the characteristics of the model and

analysis of the numerical data reveals the information

sought from the test.

Table I of chapter four shows the wind tunnel testing

sequence. Run number one is the basic model performance

with no control deflections. The Cl versus angle of attack

curve is typical for wing airfoils. The curve is

essentially linear until near the stall. A very slight

nonlinearity occurs around twelve degrees angle of attack.

The stall portion occurs over a fairly large range of angle

of attack. If this curve represented an actual aircraft it

would indicate excellent flying qualities because the stall

is gradual with sufficient warning to the pilot that the

stall is imminent.

The Cd versus angle of attack is typical with a

minimum drag at just under zero degrees angle of attack.

For the NACA 23015 airfoil used this is expected. A

noticeable discontinuity occurs at around twelve degrees

29

angle of attack and a small discontinuity in the stall

region.

30

The fuselage acts as a lifting body and contributes to

lift until it stalls. It is a streamline body with low

fineness ratio. One characteristic of such streamline

bodies is a positive pitching moment whereas the airfoil

has a very slight negative pitching moment. The pitching

moment curve shows that Cm is positive but takes a drop to

near zero at about twelve degrees angle of attack. This is

good evidence of a change of flow characteristics on the

lifting body. It is obvious that something is happening in

the region of twelve degrees angle of attack. Examination

of all of the plats shows that this occurred on all twenty

runs and at approximately the same angle of attack. The

curve discontinuity can be found on the plots far lift~

drag~ pitching moment~ and rolling moment. Very convincing

evidence can be found in the moment coefficient curves to

indicate that wing root separation is occurring here along

with separation on part of the fuselage body.

When the lifting body stalls it would be expected that

wing root separation would occur at the same time. This

would cause a slight change in wing lift and pitching

moment. The rolling moment changes at twelve degrees angle

of attack and the right wing root separation is greater

than the left wing root separation.

This accounts for a slight lass in the lift as the

angle of attack is increased through this region, for the

31

sudden increase in drag, and for moment curve fluctuations.

Since the expected results of wing root separation are

found in all of the curves, this is the likely explanation

for the behavior in the twelve degree angle of attack

region.

The cause of the fluctuation in the stall region can

be found in the rolling moment curves. It is impossible to

build an aircraft so that both wings stall precisely at the

same time. Rolling moments change significantly when a

nonsymmetrical stall occurs. The rolling moment curves

indicate this change. The first drop in the lift curve is

small and represents another indication that the stall

transition is gradual.

Plots of drag coefficient versus aileron deflection

were prepared for several wing angles of attack to examine

the results of the aileron performanc~ <Figures 4 through

10). These plots show the drag characteristics of the

split aileron along with the conventional aileron. At

small negative angles of attack there is no significant

difference. At small positive angles of attack there is a

small increase in drag with the split configuration. For

high angles of attack the drag increase is progressively

higher. Figure 9 shows some erratic behavior. This plot

is at twelve degrees angle of attack where drag

characteristics are changing due to lifting body and wing

root flow separation. A mutual interference is apparently

occurring where the upward aileron deflections are

influencing the fuselage and wing root separations.

Remnants of this behavior are still apparent at fifteen

degrees angle of attack <Figure 10).

32

The lift to drag ratio <LID>, is a useful parameter

for indicating aircraft performance and is the glide ratio

of the aircraft. A high L/D implies an aerodynamically

clean aircraft capable of a long glide at a shallow angle

and low L/D implies a steep glide angle. Figure 11 was

prepared to show the affect of the split ailerons on the

L/D ratio for the model.

' ' ! ,_ i

' i : ! ' I

-~o WING ANGL£ ___ u_ ____ bF_P:nACK-;---------- ---

1 l

I --- -!

:c. q•E;b p~f:' -----------1

l

.of:L ________ __J_ _________ - -+---____:- 1 ;=; ---'-------;

.o-s- ----- ---~ ------- ----

1

I l

------------;----j ______ t__ ___________ .f. .O't --~--------'--c---~---1 - • ! - - - ; ClR~CE'O- PolNT4 -

I

r l j . ! ' i . : J l , ! I

1-------------~-~-----: ____ L.~---i I 1

l ' !

I I ! , I i I i !

1-------- -.. --!.- ----- ------+------------- ---I I

- - I ·q . 6- ____ 3i ~5 . \.2. i : - - :

. I ' ;r· ___ : _____ :.At~_Efto_N ___ '[j)_~N- _____ :__

. : ! ' ! '

I . i . : I ;

______ /t4-~~~Jf.:\.I .. ~1G1JR'I\\_\ON

I ---------- ___ J. __ -------

!

:3

;

q i -·---1 \.

_ L A1\..flR9t'L\b1! _______ _ i - I ! i , I I i

! t t-- - -: !_:-~--~-- -_­~--~

1----- - : l!

li - h

~

--1

t

Figure 5. Drag Curves at -6 Degrees Angle of Attack

1...0 1...0

---------------------' ------- -----·-- -- - -·-- ------ -- ------------· . ------~ "l

i r--·--1 - : t·----------- --~------ . ' . I ' ' ... ,. 6D !'liE .

f'-------.--- ---+-- -- ~-~ f ···· : ~~ I · ~ I - ~, ~ 1 -- · 1

I j - ; --- ---+---- ! ' I l ~· . -'f W11\1G A!lllit~ . ' . CIRC~fp .,;iN~ AP;\o--_- ----r- ~ I OF: ATTACK i ~~ ..... .Ol\". ·- 'Of'LlT CONFIGU~Tlooi . i . i I i j . ~ ~+---- ' . ---------1

-r~~-L--~l·,_~O'\ . j ' I -------]-··----~---'- l

.:--:~~ I : 1 . :· - i , ~ r --~ ~=~ - ,_

l ' - - l

-- - - 1 ' . I .. . •. : I • -- - ----- ~--:--._ t -=---~-~-:.=:_.'_ =.:..___ ' . - . ~ .. - I -. - t·

! • . ; ; -r __ --~:=--,-- iz-:----

1iq--:- 6 -T- -~]; f - ~j i i -,- 9 l ,}!-! 1k:~: [Is

I - ; I - i : ! . I -1-- :

1 ~-------- _ AtL~RQ~ -~WI'(_~- ____ i ----- ._;AlLE~QN __ _up----- 1---· .:.~---t-~:____: _____ _ - - 1 ! i - ' ; I I i :-I I - I I I • -

t: __ ::_ .:. ___ ~_-__ :_~_.l_~: __ J :_ ~-- · :. i _ __ __________ __! _______ . _________ L , - i •---Figure 6. Drag Curves at -4 Degrees Angle of Attack

'vJ +=-

!

I l

.......... .j. -······ .......... .

l .( '

~ . ~' ' ! ~! . . •

..... '---~-------- ____________ _; ___ ..

111 1 tT !

I

i . J .. ------ ....

j

I

-'·-··------~ . I

Ira a •

35

:-; -··----------- --- ·--- ---- "1

! q•60PSF

i

!

1 -- -

' l

~ - I : - i --- - , - : l ______ _,_ ____________ _! ____ • _____ j__±~ Wtritt ANGLE_ OF ~TIACK ____ _l__ _j

. . l ! i i C• . ' ' t

' ' tl4 ' ' ' I ! j I i ! 1

! - ; --·----- -- - - -- ~05. -- ----- - f - --- - '- -- ~ -- ---- -~ ---- ---l- -- - - --- -1 l . ! ! ! ~,

I ' !

! . : 1 . ' ! 1-----------__l____ L-+--__ _;____SP!JL__ ___ ;__ -----1 • ! . • """'---: , I' I ; . ~ : .. - -~ I - - -- -

f · ·····~ ·· ·~·-··· ·- ~I~· ··· + .~.:... , . ; ... , . . . , ·- · r·-__ -·. ! --~1 f I --~- • ~ i I - ___ , _____ ;: • I

I .... .,.,__ . I • l

~ . / ! - --. j C~\IE.N't\ON.P..\.! i ji ~---------- l ; --.----;--,--- i ! j 1: !' ' ' I I I ' '

, 1 , r - -i 1·

1--~---- ·- -- _)~--- - --------L------; - .or ____ _:__~--~-_[_-----~------- -l---------~ ---- ---~----~~---1~ 1-. I I : I ! I . .. : .. - - '

' ' I I I r I ! i J ____ . }

1----l 1-

I I I '

;1z Lq : -6 : -- I - •-

: .... , . i. . ' _ _:__nt~- I::CWtL

l

q I . ! ' 3 ! r_

i l '0 I .

' . ~n,' --L-----~--JAU.~~ __ Ufl __ _

i . I , i : •

-+-

~ ~

i !

Figure 8. Drag Curves at 4 Degrees Angle of Attack

i

I J

'u> ()'.

! i i

Ca.

---~--- ---- ----t-- ·------.----~---~~-L~-------~-------· ----

.10 cp~ ~ '~F

--l-~- .. t-- --

i ! . . i ' I : I

_.o_q_ ________ j:-j[ __ \NJM_~AN~L£ of" ATT1c.K i ' .. l .. ·- t j : j I

·- f- ... ·------ __ , ---------- -~-- --· -.08 i I

! --1 i

' ' -----!

1 ! ~ l I ! l I

~~--- ------ -----t- .... D-"l ... ~. ----~-----t--------'----: -------1 i ! SPi\T I . ' ·"' ' ' I J--- · ·- --;----~--- ·-··· -·--·-·· ·----------·---~----- - -- ·r· ----- ---------t---------~------ ______ .._ _________ ---~ i ! ; I 1 1 I

, : . . l ! I ! I , , , l . ~ 1 i i , I i------- ___ __1 ___ ~----- _i _____ · ~- --~ _____ L_ ______ ~------L--~-----'-

1 i f i ' l • l i l l I ' ' I ; i ' r I . i

l! ·---- -]----- j --------j , _ ·----- - -----t_--l:o_-~~~o.l~- --- ---j ! I ! i ' I ' ' IF ,, '

' 'I ! ' 'I •. . ·, ' I,. ·,t '. :• I I'' ·'2. i ~ 6 i 3 ~ 3 ! 6 q I IZ ~ jg I

f---- -__ A ILE~9N Q9.:Wf'l( .. - - -- -· ; ...... AtlfR9-~ -\.l\? l- ------ ·-'·-- --- --+---- -- --- ... I I i i ' I

~ l 1 i ~ ! I 1 . . , ' ' • 1 1-- ... ------------ --~-----~--~-----1----'-~--...J._ ______ , ____ .l__ ---------------! . --~--~

Figure 9· Drag Curves at 8 Degrees Angle of Attack

VJ -...:1

; r-------- ---- .

I I-I

.I£"

c .. q•:GcfPSr i

__ _c ______ --IJ~---~---- -~--l ···---------i------------·-1 ! ! I

--- -- ... - -.......--- - ----·-- -:-------- -----, ~-·fl !

.tz ±lt'_W.\N§.AH ~1£_QL~II~_ck_ -------t-----.. --- . ;

_......_,__ __ ~~lJT i i

··]----~-----

I. I : : ••v , ~----j--·-. ------T .,. . ~ I · 1

I . I ! . . r----------~ ....... --+- ___________ ..._ ________ _ I . ! I : ,.

I· I

i ' t

--+-!

' I I

I

12 ! q : 6 I

.. .At~RON. QO'WN I

I· . .. L~---·--·

I •

!

3 ; . ' '3 i 6 : q

. AtLERON~ UP i I

Figure 10. Drag curves at 12 Degrees Angle of Attack

:1~.

I I

I I

-I I

\...0 (X)

I f. - ---

' : ~----- --------·-----!------------ --··---

i r--- --! I

' I

. -·--· -t--·--· ~-!

i

' q:. 60 p~

.. --~ ~--m- ;-i:l~irr!a~ ANGLE ~~---~-_ 1

' ' . 1 . j . i ' .. ID ; · ! ' .

-- ____ .. J.1 __ - - . ' - l . I --- -------- . --- ----- ---t------~· ---·-j ' ' . l I , ~ I . .

----------- _j ______ . --- _j

I I ·----~-! t--1-- ! --r -. ~ 1 - -- . 1 . _ ! I I ! - . : I ; -l

[---- ---- -~-- ------+-------~-~ · -- f ---~- "+~- ---- -- ---T -- --+---- -- ; ---- -~---_:--+-- ~-- 1

I l i : C;(!)NVEtilTIONAL I I

~I : i ' i . I :. : ! i I'

- . i ·r-----. ' . . . . i . . ·-· ______ ! . - !

r=---~ --~- ---- -l------~-:---------~-- ---~------- :--------r---- -- _i --_ -- -,

l i i i i I f il : ' I I' I ' ! ' f i • ' I I ,s ! ll I . 'C\ . ' : . 3 ; i : 3 ; 6 q ! \2 ~5' : I& ! L-----1~~~~~r--o~~~ _ .J ______ !_ _ -~'uao~ ~~--- t----- _ _c ________ J _________ • ____ 1

Figure 11. Drag Curves at 15 Degrees Angle of Attack

I....V \,()

f:tj 1-'• Oti c 'i ([)

1--' (\)

t-1 1-'· 1--!J c+

~ 'i Ill

Oti

"'d 1--' 0 c+ (/l

17

16

15

14

13

12

0 11 •.4 +' ~ 10

bO ro ~ 9

0 8 +'

+' '+-< •.4 7 ..:!

6

5

4

J

Clean configuration

Split on both Ailerons

t_ • - --• --~ -- -- ~ ~

1 2 J 4 5 6 7 8 9 10 n12- -f314 -is -i6--i? Angle of Attack

+ 0

CHAPTER VI

DISCUSSION OF RESULTS

The purpose cf the split aileron is to add parasite

drag at the same time induced drag is being lost. Parasite

drag addition ideally should be equal to the induced drag

reduction plus the parasite and induced drag increase of

the opposite wing. <See Figures 4 through 10).

As was mentioned earlier, a problem of design

optimization could be expected. Since parasite drag and

induced drag vary in an opposite manner for changes in

flight airspeed it would be expected that a balance can

only be found for one airspeed.

When aircraft airspeed is changed the angle of attack

increased or decreased to change the lift coefficient to

maintain a balance between lift and weight of the aircraft.

At low speed a high angle of attack results to achieve high

Cl. Under this condition the forces contributing to

adverse yaw are great. Since parasite drag is lower at low

speed, intuition would indicate that a large aileron split

angle would be required. An aircraft operating at high

speed is operating at low Cl and the forces contributing to

adverse yaw are small. Since parasite drag is high at high

41

speed it would seem that a small split angle would be

required.

These factors imply that the adverse yaw can be

countered at only one airspeed and/or one split aileron

deflection. Inspection of Figures 4 through 10 show that

this is not the situation. At low angle of attack where

42

little drag increase is desired, there is little difference

between the conventional aileron and split aileron

performance. At moderate angle of attack, there is a

slight increase in drag for the split concept. At high

angle of attack, where high drag increase is desired,

there is a larger increase in drag for the split aileron

over the conventional aileron. The split aileron design

would appear to modulate the drag contribution according to

demand. It is contrary to expectation but is the behavior

desired for such a control device. The implication is that

an optimization can be performed over a much broader range

of operating conditions than originally anticipated. Even

if the design could only be optimized aver a narrow range,

it would be an improvement aver conventional control design

where no optimization was possible. In fact, conventional

design does little, if anything, to counter adverse yaw.

The fact that optimization is possible over a broad range

increases the utility of this concept significantly.

At negative angle of attack another interesting result

is noted. A device designed to counter adverse yaw for

positive angle of attack would be adding to adverse yaw for

4-~

negative angle of attack if it continued to function. The

figures for negative angle of attack show identical drag

characteristics for split and conventional aileron designs.

The maximum difference of the curves is .001 which is the

level of round off. No difference is measurable, and is a

convenient result. Test data was not taken at deflections

below negative six degrees angle of attack so operation

below negative six degrees cannot be predicted. However

only an experienced pilot with an aerobatic type airplane

would purposely operate at such negative angles.

One area of concern for a control surface is the stall

region. One must insure that no unusual characteristics

are present because a pilot must have good control response

for stall recovery. The data shows that the lift and drag

curves are of the same form for both the split and the

conventional aileron mode. Deep stall operation was not

tested in this experiment so that region remains unknown.

Only T-tail aircraft experience the deep stall and

airplanes are not normally flown in the deep stall region.

Thus it does not represent an area of concern.

Runs 13 and 14 were accomplished to compare the

rolling moment capabilities of the model for the two

modes of operation. The rolling moment curves have

approximately the same form. However the split mode curve

shows a loss of rolling moment of 20 percent at low angle

of attack and about 50 percent at stall. Since these

curves were made for identical aileron deflection angles

44

the implication is that the designer would have to provide

greater deflection angles to maintain the s~me rolling

moment. Large rudder deflection is required to counter the

large adverse yaw during a stall with a conventional

aileron configuration. Adverse yaw induces a rolling moment

in the opposite direction to the aileron induced moment.

This counters part, or perhaps all, of the aileron moment.

The split mode aileron could be depended on to maintain

rolling moment in the desired direction. It is necessary

that the downward deflection angle be matched with the

proper split mode angle to achieve a proper balance of drag

forces. This functional relationship could vary for

different angles of attack. The proper match of angles for

a given angle of attack can be found from plots such as

those in Figures 4 through 10. A horizontal line can be

made intersecting the downward deflection curve on the left

and the split mode curve on the right. The intersection

indicates the appropriate angles. Inspection of the curves

indicates that the split mode opening angle on the up going

aileron must be significantly larger than the down going

aileron. Since the split mode curves tend to be

essentially flat and then start to increase, it can be seen

that it would have been useful to have data points beyond

the eighteen degree maximum tested. Most of the

effectiveness is realized at higher split aileron

deflectiqns.

The match between angles for control deflection

45

appears to be essentially independent of angle of attack of

the wing. This is significant from the design standpoint

in that optimization can be achieved ever a broad operating

range.

The L/D curves (see Figure 11) are fairly typical for

an aircraft. The significant feature is the displacement

between the curves. Even though the ailerons span only a

portion of the wing, deployment of eighteen degrees of

split aileron lowers the overall aircraft L/D

significantly. This implies that the split control could

be used as a glide path control device.

Some aircraft have ailerons that can be lowered

simultaneously to augment flap action. Differential

deflections far rolling moments are superimposed on this

aileron droop. In a similar manner a symmetrical split

could be applied to both ailerons to provide additional

drag to act as a speed brake or to provide a steep landing

approach without gaining speed. The differential action

for roll control can be superimposed on this symmetrical

split. Such a symmetrical split is not limited to raising

the upper surface alone while leaving the lower surface in

normal position. It could be accomplished by moving both

surfaces apart from neutral or by displacing only the lower

surfaces in an action similar to that of a split flap.

This would increase the mechanical complexity of the

control system. This is merely a design problem, however,

and similar systems are already in use on STOL aircraft.

46

As an example, an aircraft flying on approach without

flaps would have a Cl of 0.8.

the L/D is approximately 14.5.

In the clean configuration

At the same lift

coefficient with eighteen degree split deployment the L/D

drops to about 8.5, a reduction of over forty percent.

This is sufficient for glide path control. As another

example, an aircraft at high speed would have a Cl of 0.3.

For this situation L/D drops f~om 12.5 to 7.5, again a loss

of forty percent. This would serve nicely for a high speed

penetration maneuver such as performed in some instrument

flying procedures or when descending through a small

opening in a cloud layer.

Spoilers have been used successfully as glide path

control devices in the past. However few small aircraft

have them and a major wing redesign is required to install

them. The split aileron design maintains the advantage of

serving as a control surface replacement item.

• CH?-IF'TER SEVEN

CONCLUSIONS AND RECOMMENDATIONS

FOR FURTHER STUDY

The wind tunnel testing and data analysis show that

the split aileron concept provides a simple device to

counter adverse yaw. It has the unexpected property of a

modulated drag affect that permits it to be optimized over

a broad range of operating conditions. One of th~?

strengths of this design is that it could be easily

retrofitted to existing aircraft designs without changing

th~:·~ w:i. ng d€-?~~:i. gn. It would amount to a control surface

change with some modification of existing actuator

linkagf?s.

The test data showed that it can be used as an

effective glide path control device. Although this was not

the original objective, it is a bonus.

Since this was the first test of this idea~ the wind

tunnel tests were designed to establish the validity and

any unusu;:..l bl~h.-avi Dr Df the CDncept.

areas were ne~t investigated.

Consequently some

Additional testing could be done to experiment with

greater control surface deflection angles. From this

testing a detailed optimization profile for downward and

47

upward aileron settings could be established. This would

provide design of an operating linkage concept.

48

Another area for further research would be to

determine why the drag increment is modulated with angle of

attack. This result was not expected and cannot be

explained. Flow visualization techniques will probably be

required here.

Further research should be accomplished with other

airfoil sections to determine if the performance of the

split aileron is independent of the airfoil section.

Modern airfoils such as the GAW series should be

investigated.

Investigations of deep stall and extreme negative wing

angle of attack should be performed to discover any

undesirable or limiting characteristics. This research

could be performed with the original test model.

Hinge moment testing should be accomplished. This

will insure that a hinge design can be made to permit the

pilot to easily operate the ailerons with unpowered

controls. It must be insured that there are no sudden

changes or reversals of hinge moment which would create

control problems for the pilot.

A design should be made for retrofitting an existing

aircraft with this concept. The cost of producing and

marketing the design should then be determined. This will

permit assessing the financial feasibility of the concept.

A SELECTED BIBLIOGRAPHY

(1) Abbott, Ira H. and Albert E. von Doenhoff. !h~QCY gf_~ing_§!£tiQD!• New York, McGraw-Hill 1949.

<2> Davidson, M. L. E~£!!!t~_Q!~£C!ei!gn_gf_tb~ Z_!_lQ Eggt-~2lt~c-~~-~!~£b_~!mQc!~l-~!n2 I~no~l· Aeronautical Engineering Department, Wichita State University, Wichita, Kansas, 1979.

<3> Dole, Charles E. El!9ht_!h~QCY-2QQ_B~C9QYO!ffii£a· John Wiley & Sons, 1981.

(4) Garber' Paul E. "Warped Wings. II s~Ql!::!i!QQ __ gf B!c'c~ft_~!ng_Q~~!gn, American Institute of Aeronautics and Astronautics, 1980.

<5> Habluetzel, Tracy. IhC~!=Q!m!D!!QD~l_Egc;~-Q~t~ 8'9Y!§!t!gn_AQQ_B!Q!::!£t!gn_Bgyt!n!_Egc_!b! ~llt!c_H~-~!~£h_Z_K_lQ_Eggt_bg~_§Q!!~-~!n~ IYOD~!· Aeronautical Engineering Department, Wichita State University, Wichita, Kansas, 1980.

(6) Lissaman, P. B. S. "Wings For Human Powered Flight." 5~9!!::!!!gn_gf_B!c;c!ft_~!D9_Q!!!9D, American Institute of Aeronautics and Astronautics, 1980.

<7> McCormick, Barnes W. B!CQQ~D~ffi!£!L_6!CQDIY!!£aL and E!!9h!-~!£h1Di£!• John Wiley & Sons, 1979.

( f::1) Perkins, Courtland D. and Robert E. Hage. E!cfgcm!D£!_§t2P-ilitY_In2_Qgntcgl. 8< Sons, 1949.

6!CQlAD.~ John Wiley

(9) Pope, Alan and John J. Harper. bg~_§g_~gg_~!DQ I~DD!l I!!!!ng. John Wiley & Sons, 1966.

<10) Rae, William H. and Alan Pope. bg~_§Q!!Q_~!DQ !YDD!l I!!t!ng. John Wiley & Sons, 1981.

<11) Riegels, Friedrich W. B!~QfQ!!_§~;t!QQ§, London, Butterworths, 1961.

49

( 12) Shevell, Richard S. Prentice-Hall,

EYD~~ffi§O~~i§_gf_Ell9bt· Inc:. , 1983.

<13) Torenbeek, Egbert. §~ntb§§i§_Qf_§y~§QO!~ eicQl~D~-Q~ai9D• Delft University Press, Delft,Holland and Martinus Nijhoff Publishers, The Hague, Holland, 1982.

<14> Warner, Edward P. Blce!~o~_Q§§ign_E§CfQCffi~Q£§· New York, Mc:Gr·aw-Hill, 1936.

50

APPENDIXES

!'51

APPENDIX A

MODEL CONSTRUCTION

Details of the wind tunnel test medal construction are

provided as a matter of record~ The wing section chord and

span ware chosen to be consistent with the wind tunnel test

section dimensions. Based on past experience the wind

tunnel operators recommend that wing span should not exceed

70% of the test section span. This limitation minimizes

errors from wall interference. This limited span to seven

feet or less due to the ten foot test section width of the

Walter H. Beech Memorial Low Speed Wind Tunnel. A

reasonably thick wing is desired to obtain structural

strength and provide space for actuating systems. A thick

wing requires a long chord for a given airfoil section.

For a given span, a long chord leads to a low aspect ratio.

A high aspect ratio is desired to reduce wingtip effects.

Thus airfoil thickness and aspect ratio require a

compromise. An aspect ratio of seven was chosen as a

reasonable value to represent light general aviation

aircraft. This results in a wing chord of one foot and a

wing area of seven square feat.

Dimensional tolerances must be very close for wind

tunnel models to achieve reliable test data . Structural

•

53

strength is important because the aerodynamic loads of the

wind tunnel can be large compared to the model size. If

the model deforms under load it is impossible to predict

the actual angle of attack and control settings for a given

set of data. For the data to be reliable the deformations

under load must be small. Therefore it is necessary to

build in rigidity. Dimensional accuracy and rigidity are

major considerations for model construction. The model

design was based en these considerations, available

material~ and equipment.

The main structural member of the wing -is a simple

vertical spar at 25% chord. The spar is a continuous seven

foot length of 0.132 inch stainless steel milled to 1.400

inch height. The stainless was chosen for its high

strength. Ribs were placed at one foot intervals along the

span. Aluminum alloy 6061-T651 of 0.375 inch thickness was

chosen for the ribs. This thickness permitted the wing

skin to be attached by rivets driven into the ribs without

the use of flanges. Rectangular sections were cut and

milled to high dimensional tolerance to construct the one

piece ribs. All interior cutouts and holes were made

before the ribs were milled further because it was easier

to accurately locate positions for the cuts while the rib

blanks were still rectangular. Two reference holes at two

and seven inch positions of 0.1875 inch diameter were cut

along the chord line. These holes were later used to hold

the ribs in a jig fixture by use of steel dowel pins.

54

After all interior cuts were machined a carefully made

wooden airfoil was attached to the rib blank with dowel

pins and the airfoil shape was scribed onto the rectangular

aluminum rib blank. An approximation of the airfoil was

then cut by use of straight line cuts made on a vertical

milling machine. This series of cuts was made to within a

few thousandths of an inch of the airfoil shape. This

airfoil approximation was then bolted into a wooden jig

illustrated in Figure 13. A metal cutting tool was

installed in a router and used to cut the airfoil shape.

By geometrical techniques developed by the author this

wooden jig was cut in a manner such that when the router

base was guided along the outer curves~ the cutting tool

edge scribed the airfoil shape on the inside. This left a

slightly rough surface which was then smoothed by hand by a

single cross cut file.

This procedure was experimental in nature and was

necessary due to a lack of profile milling equipment. The

greatest concern was dimensional consistency. Although it

was not necessary that the airfoil be a perfect

representation, it was necessary that all eight ribs

produce identical airfoils. The maximum deviation of any

airfoil from any other airfoil is approximately 0.002 inch.

Figure 14 shows all eight ribs bolted through the reference

holes and the good fit is evident.

Figure 14 also illustrates a circular cutout at the

front of each rib. This was finish cut on the vertical

Figure 13. Rib Profiling Fixture

Figure 14. Rib Sections

56

mill with all ribs bolted together to insure consistent

location. This cutout was used to install an aluminum rod

along the leading edge to establish the leading edge radius

curve when the sheet metal covering was installed.

Figure 15 illustrates the spring loaded mounting

mechanisms for the two wing mounting points. A second spar

is placed between the mounting fixtures to prevent the lift

loading from twisting the main spar. This secondary spar

was made of aluminum and.is not primary structure.

Actuation of the mounting device was provided through small

slots in the lower wing surface.

The ailerons were constructed for high torsional

strength so that a single point attachment could be used

for setting deflections. The upper and lower aileron parts

were milled from single pieces of 6061-T651 aluminum.

Figure 16 shows the aileron parts.

Some means to set aileron deflections was necessary.

An internal mechanism is more streamlined but the shallow

airfoil thickness at the aileron hinge point provides poor

mechanical advantage. "Blow down" of control surfaces is a

common problem of wind tunnel models. The aerodynamic

loads tend to cause the surfaces to deflect somewhat from

the static setting. External brackets were used with an

indexing arrangement for bolting the aileron surfaces

rigidly in place. The index permitted three degree

increments of aileron position. The indexing bracket was

attached externally to the wing structure and the upper

57

Figure 15. Tunnel Mounting Fixture

Figure 16. Milled Aileron Components

aileron surface was secured to the indexing bracket. A

single small bolt at the trailing edge attached the lower

58

aileron surface to the upper surface. When the split mode

operation was desired this small bolt was removed and two

small drive pins in the indexing bracket were driven

forward to capture the lower surface at the zero deflection

position. The axis of the drive pins was tapered slightly

to eliminate all slack from the ailerons.

for an illustration of this arrangement.

See Figure 17

The ailerons were sized to cover two feet of span on

each wing tip. The chord length of the ailerons was chosen

at 30% chord or 3.6 inches. This is a figure typical of

control surface design.

Small rods threaded on both ends were placed between

ribs and used for rigging purposes only. The structure was

rigged into alignment using a precision straight edge along

the leading edge <see Figure 18). Figure 19 shows the

structural assembly ready for the surface skin.

The surface skin was made from 0.040 inch thick

aluminum sheet metal. This thickness was used to provide

rigid surfaces between ribs to prevent deflection under

dynamic loading. A tail sting of 11/16 inch aluminum

tubing was attached securely to the main spar to provide

the third mounting point. Additional structural support of

the tail sting can be seen in Figure 20.

A fuselage was provided for the test model to simulate

the wing-fu5elage flow interference of an actual aircraft.

60

Figure 19. Complete Structure Prior to Covering

Figure 20. Tail Sting Components

This arrangement provided the best simulation of actual

aircraft flow conditions to enhance the applicability of

the results. The fuselage was constructed by bonding

planks of wood under pressure and turning to final

dimension on a lathe. The upper and lower parts of the

61

fuselage were provided with an internal slot to fit tightly

on the tail sting. This increased the rigidity of the tail

sting to prevent flexing of the tail sting and change of

the angle of attack under load.

Figure 21 shows the complete model installed in the

wind tunnel for testing with the ailerons in neutral

positions. Figure 22 shows the model in the tunnel with

maximum split on the ailerons.

Figure 21. Test Model Installed in Wind Tunnel with Ailerons in Neutral Positions

Figure Maximum Split on Ailerons

6~ L

APPENDIX B

WIND TUNNEL NUMERICAL DATA

64

f I c ~ ! L~

' I., l ~ l j

~ ~ 1<~:~:

r • .~ It:

n II~ ~ q

' I~ ~~ I~ ~ I~

1-!-"' u I I (] o:~ I~ I~ ~i ~~ I~ I~ ~~· f~ ·~ ! ~ ~ :r-: IJi ~~ ~ ~ I~

I I .... _ ... 'i lJ 1- -1. 11-I_,

~ c c c li: ' ~ I~

ol~ I~ :s I~ I~ ~~ :u 11]1nl I 1-: 1"1 I'' 1-s ll' lq ~ ~1:: ~~ I~ I~ I~ I~ ~~ I~ I~ ~~

_,... N"'

tt~ I~ 1-~ II : ....

'l'J 11'\o 1-: 11 I~ .air- Itt Ia- ~ ~ I~ I~ I~ I~ I~ I~ ~~ I~ I~ I~ I~ ~~~

~~- I~ 1-~ I~ b

I~

~ I I I ll ~ ..... ~

~ I I ~ -~ ~ _ ....

.. .II J ~ ~ -... - - ~

~ ~~~ ~ ~ ll ~

~~ .: ...

~J ~~ ~ 1t u ~

.J c ( 0 IJ~ J I~ ~ j IJ I~ ~ Oleo? ~ ~ I~ j 0 () 00 ~ \..! -.: - ..s ..S)

~ 1 j J -rl I~ 1-:~- f&.1 .JI t" I~ f' ~ ::: ·~ !!l I~ ·~ I~ I~ I~ :a.: ~ 11

rl -I~

65

WICHITA STATE UNIVERSITY 7 X ~0 FOOT LOW-SPEED WIND TUNNEL

c:onst. tllble 1 Static T11re 1000

MARCH 21, 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY Model upriQht

Stat.lc Tare Countl!l

ALPHA PSI lift dra9 pitch roll yaw t~ideforce

-eo 0 0 0 -3 1 0 0 -40 0 0 1 -4 0 0 0 0 0 0 1 0 -1 0 0 40 0 0 1 7 -2 0 0 80 0 0 1 19 -3 0 0 120 0 0 1 36 -4 0 0 160 0 0 1 S7 -s 0 0 200 0 0 1 82 -6 0 0

66

W :1: C H :1: T A B T A T 1::: U N :1: V 1::: I~ ~;:; :1: T Y '? X :1. 0 FDDT I ... DW····SP EI:::I> W :1: ND TUNN1:::1 ...

const. table 1 Static Tare 2000

MARCH 21 1 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY Model upright

S til t.i c Tare Counts

ALPHA PSI lift drag pitch roll yaw sideforce

-so 0 0 0 -3 2 0 0

-40 0 0 0 -3 1 0 0 0 0 0 0 0 0 0 0 40 0 0 0 7 -1 0 0 80 0 0 -1 19 -2 0 0 120 0 0 0 3S -4 0 0 160 0 0 1 56 -4 0 0 200 0 0 1 81 -6 0 0

W :1: C H :1: T A B TAT m: UN :1: VIE I~ S :1: T Y '7 X :1. 0 F D 0 T 1. .. 0 W ···· ~:; P 1::: 1::: X> ' W :1: N l) T U N N 1::: 1. •. •

const. table 1 DynoM.i.c: Tare 3000 Stat.ic Tare 2000

MARCH 21, 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY MOdel upr .i.gh t

w.i.nd axes data about trunnion

ALPHA PSI CL CD CH CRH CYM CY q <psf') (deg> <deg>

··8 0 .0012 .0305 -,0546 0 0 .002 S9.4S -4 0 .0012 . 029 -. 0528 0 -.0001 .002 59.44 0 0 .0013 '0277 -, OS21 -.0001 -.0001 .0022 59.34 4 0 .0013 . 026S -.OS1S -.0001 0 .0022 59.4:3 8 0 .0014 '0253 -.0501 -.0001 -.0001 .0022 59.33 12 0 .0014 . 0239 -. 0485 -.0002 0 .0022 59.33 16 0 .0014 . 0227 -. 0468 -.0001 0 .0023 S9.23 20 0 .0014 .021S -. 04S -.0002 0 .0023 59.42

REYNOLDS NUMBER/FOOT= 1.44S97E+06 MACH NUMBER= .2043

MODEL CONSTANT TABLE 1

7 1 7 7 0 .4 .6 .4167 0 . 0064 2.2 1.04 .as .9S .12 0 0 0 .a 0 . '7188

68

W :a: C H :1: "Y" A ~=~ 'T" A 'Y' m: l.J N :1: V m: I~ m :1: T Y '7 X ::1. 0 1::· C:J 0 T 1 •.• C) W ···· ~=~ 1::~~ m: 1::: l) W :1: N l) T U N N m: 1...

Run nul'lber 1 Stat .i.e: Tare 1000 DynaMic Tare ~~ 0 0 o const. table 1

MARCH 21 1 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY l'lodel upr.l.ght

W.i.nd AX !IS Data

ALPHA PSI CL CD CM CRM CYM CY q <psf'> < deg > (deg)

-6.27 0 -.3786 . 0308 -.0258 -.0026 -.0002 .0004 59,32 -4.15 0 -.2102 ,0238 -.0194 -.0029 -.0001 .0004 59.32 "'"~~I 02 0 -.0321 . 0203 -,0115 -.0028 -.0001 .0003 59.32 . 1 0 '148 '0206 -.0041 -.0025 -.0001 .0001 59.22 2.23 0 .3199 I 0244 .0034 -.0031 -.0002 .0002 59.21 4.35 0 .4948 '0321 .01 -.0032 -.0002 .0001 59.21 (). 47 0 .6643 .0434 .0168 -.0029 -.0001 -.0001 58.91 B.S9 0 .8361 • 058 . 0224 -.0028 -.0001 .0001 58.91 10.7 0 .988 .0757 .0283 -.0024 -.0001 0 58.62 12.76 0 1.0715 .1184 ,0099 -.0054 -.0001 .0008 S8.S2 1.3.79 0 1.1134 .1339 ,0083 -.0069 -.0002 .0016 58.35 14.82 0 1. 1SS6 .1492 .0051 -.008 -.0002 .0033 58.28 15.82 0 1.1655 .1?38 -. 004 -.0062 0 .0024 58.06 16.81 0 1.1457 .1969 -. 0152 -.0075 -.0021 -. 0035 57.94 i 7. Bi 0 1.1447 .2166 -.0212 -.0123 --.0031 -. 003? 57.6 18.83 0 1.1751 .2364 -. 02S -.0063 -.003 -.0048 57,65 19.83 0 1.1693 .2696 -.0439 .OO?S -. 002 -. 005 S7.SS 20.82 0 1.1521 .2871 -. 0524 .0186 -.0002 -.ooss 57.51

REYNOLDS NUMBER/FOOT• 1.40282E+06 MACH NUMBER= .201

WICHITA STATE UNIVERSITY '7 X ::1. 0 F C) 0 T 1... C) W ···· S P E t::: l> W :1: N l) T U N N 1::: 1...

l~un nuMber 2 Stet .i.e: Tare 1000 DynaMic T11re 3000 const. to.ble 1

MARCH 21, 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSlTY folodel upr.i.ght

Wind' A us Dr.ttQ

ALPHA PSI CL CD CM CRM CYM CY q <psT> < deg > (deg>

-6. i1 0 -.1619 .0419 -. 0811 .0024 -.0006 .0004 59.28 ·-4 0 -.0045 . 0398 -. 0726 .002 -,0006 .ooos 59.28 ··1. 88 0 .1645 . 0413 -.0627 .0026 -.0008 .0009 59,08 . 24 0 .3377 . 0462 -.0543 .0023 -. 001 .oooa 59.28

;.: .36 0 .5078 . 0546 -.0464 .0022 -.0011 .0006 59.18 4.48 0 .6795 . 0664 -.0397 .0021 -.0012 .ooos S9.08 6.6 0 .8471 . 081S -.0321 .0019 -.0013 .ooos sa.ae 8.72 0 1. 0107 .1002 -. 0266 ,0019 -.0013 .0004 58.79 10.83 0 1.1676 .1225 -' 0208 .0015 -.001S .0001 58.59 12.8!3 0 1.2441 .1727 -.041 -.0029 -.0018 ,0021 58.22 13. 9j, 0 1.2923 .1895 -.0419 -.0034 -.0017 .0037 S8.4S 14.94 0 1.33S .2083 -. 0451 -,0039 -.0016 .OOS7 sa. o9 15.93 0 1.3191 .2367 -, OS6 .0008 -. 002 .0025 5?.88 16.91 0 1.2865 .2632 -.0683 -. 0118 -. 006 -.0025 57.88

REYNOLDS NUMBER/FOOT= 1.39808E+06 MACH NUMBER= .201?

70

W :1: C H :1: T A ~:) T A T m: UN :1: V a::: a~ ~;; :1: T Y '7 X ::1. 0 F Cl Cl T 1... D W ···· B P a::: a::: X> W :t: N l) T U N N 1::: 1...

Run nuMber 3 Static Tr.1re 1000 Dyn<JMic T<Jre 3000 c:onst. t<Jble 1

MARCH 21 1 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY MOdel upright

W.Lnd A us Data

ALPHA PSI CL CD CM CRM CYM CY q <psf> < deg > ( d eg)

-··6' 14 0 -. 199S . 0382 -. 0736 .001S -.0003 .0007 S9.S6 -4.03 0 -.0381 . 03SS -. 0649 .0013 -.0003 .0006 59.46 -· 1. 91 0 .1251 .0363 -. OSS3 .002 -.ooos .0008 59' 17 .21 0 .3012 .0407 -. 0467 .0018 -.0006 .0008 59.36 ~~. 34 0 . 4'741 . 0483 -. 0394 .0014 -.0008 .0008 59.36 4.46 0 .6467 .OS99 -. 0328 .0013 -.0009 '0 0 0'? S9.27 6.58 0 .8132 .0743 -.0254 .0009 -.0009 .0009 59.07 8.69 0 .9786 '0922 -' 0189 ,0008 -.0008 .001 59.07 10.8 0 12.86 0

1. 131,8 '1137 -.0132 .0006 -.0009 .001 58.88 1.2086 .1619 -.0318 -.0035 -.0013 .0022 S8.6

13.8'? 0 1.2625 .1792 -. 0351 -,0045 -.0013 .0035 S9.S3 14.92 0 1.304 '197 -. 0369 -.OOS4 -.0013 .0055 S8.S6 15.94 0 1.3235 .2194 -.0438 -.0039 -.001 .0056 S8.42 t6.94 0 1.3321 .2443 -.OS -.0023 -.0009 .OOS3 S8.3 17.9 0 1.2697 .2721 -.0623 -.0111 -.0049 -.0021 58.21

REYNOLDS NUMBER/FOOT= 1.3BS90E+06 MACH NUMBER= .2023

71

WICHITA STATE UNIVERSITY '7 X :1. 0 F 0 Cl T 1... 0 W ···· ~:; P E 1::: I) W :1: N X> T l.J N N 1::: 1 ...

Run nuMber 4 Stat .i.e Tare tOIIO DynaMic Tare 3000 c:onst. table 1

MARCH 21, 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY MOdel upright

Wind Axes D.ata

ALPHA PSI CL CD CH CRM CYM CY q <psf> (dtQ) <deg>

-6.17 0 -.2368 . 035 -. 0648 .0003 -.0002 .0003 59.55 ·-4. OS 0 -. 0758 .0318 -. 0562 .0001 -.0001 .0004 59.5S ·-1. 94 0 .0883 . 0319 -. 0468 .0003 -.0002 .0006 59.25 .19 0 .267 . 0353 -. 038 .001 -.ooos .ooos 59.25 2.31 0 .4394 .0424 -.0307 .0007 -.0006 .0006 59.15 4.43 0 .6111 .0531 -. 0239 .0003 -.0007 .0006 59.05 6.SS 0 .783 . 0673 -.0168 .0003 -.0006 .0006 58.96 8.67 0 .9452 . 0845 -.0102 .0004 -.0006 .0009 58.86 i 0 . 7!3 0 1.1 .1053 -' 004 .0006 -.0008 .0009 58.67 12.83 0 1.1773 '1534 -. 0228 -.0033 -.0011 .0025 58.39 13.87 0 1.2233 '1703 -. 0252 -. 0049 -.0011 ,0036 58.22 14.83 0 1.1766 .1962 -. 039S -. 0073 -. 0044 -.0003 sa .12 15.85 0 1.2048 .2158 -. 0437 -.0108 -.0049 -.0015 sa. o7 16.86 0 1.218 .2403 -.0512 -. 0122 -. 0052 -.0013 57.84 17. 8!3 0 1.2368 .2596 -. 0528 -.0111 -.0053 -.001 57.79

REYNOLDS NUMBER/FOOT= 1.37593E+06 MACH NUMBER= .2015

72.

WICHITA STATE UNIVERSITY 7 X ~0 FOOT LOW-SPEED WIND TUNNEL

Run nuMber S Static: Tare 1000 DynaMic Tare 3000 const. table 1

MARCH 21 1 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY Model upright

W.i.nd Axes Data

ALPHA PSI CL CD CM CRM CYM CY q < psf> <de g > <de g >

·-6' 2 0 -.2769 '0304 -.0551 0 -.0001 0 59.43 -4' 08 0 -' 1138 '0263 -. 0471 -.001 -.0001 0 59.33 ··1 '96 0 .0508 '0258 -. 0377 -.0006 -,0001 .0002 59.14 .16 0 .2264 . 0283 -.0281 -.0004 -.0003 .0002 59.23 2,28 0 .3987 .0343 -. 02 -.0001 -.0005 .0003 59.33 4.4 0 .5706 '0441 -.0127 -.0001 -.ooos .0002 59.13 6.52 0 .7406 . 0575 -. 0055 .0001 -.ooo5 .0002 58.84 8.64 0 .9069 .0739 .0004 -.0004 -.0005 .0001 59' 04 10 '75 0 1.0576 • 0941 .0059 -.0004 -. 0 0 04 -.0004 58.85 12 '81 0 1.1404 '1408 -.0124 -.0051 -.ooos .0014 58.56 13.83 0 1.1755 .1587 -. 0179 .0032 -.0004 -.0001 58.5 14.84 0 1.1815 .1843 -. 0273 .0005 -.0003 .0016 58.48 15.83 0 1.1713 .2015 -.0346 -' 0113 -' 0042 -.0029 58.44 16 '8:"3 0 1.1779 .2247 -.0399 -.0126 -.0045 -.0027 58.31 j.7 .84 0 1.1853 .255 -' 0532 -.009 -.0013 -.0017 58.1

REYNOLDS NUMBER/FOOT= 1, 36739E+06 MACH NUMBER= .2021

73

WICHITA STATE UNXVERBXTY '7 X j, 0 1::· DDT 1... CJ W ···· ~:; 1:) m: 1::: I) W :1: N X> T l.J N N 1::: 1...

Run nuMber 6 Static Tare 1000 DynaMic Tare 3000 const. table 1

MARCH 21 1 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY Model upright

W.i.nd Axes Data

ALPHA PSI CL CD CM CRM CYM CY q <psf'> < deg) <deg)

-·6' 23 0 -.3205 '0291 -.0421 -' 0009 0 .0001 59.43 -4. u. 0 -.1534 '0237 -' 0352 -' 0011 0 .0003 59.32 -1 '98 0 .0218 '022 -.0267 -.0011 0 .0003 59.22 .14 0 .1961 .0239 -.0182 -' 0008 -.0001 .0003 59.32 ;z ,26 0 .368 . 0291 -.0102 -.0012 -.0002 .. 0004 59.32

* 4 '38 0 .5376 . 038 -.0026 -. 0012 -.0002 .0004 59' 12 -1-6,5 0 .7076 .osos ,0047 -.0009 -.0002 .0006 59' 02 8.62 0 .8698 '0661 '0113 -.0008 -' 0001 .0007 58.83 10 '72 0 1.0187 .0872 .0159 -.0009 -.0001 -' 0001 58.64 :t;~ '78 0 1. 0977 .1283 -.0005 .0036 -.0003 .0015 58.34 13.8 0 1.1259 .1494 -.0072 .0022 -.0002 .0015 58.3 14.81 0 1.1375 .1734 -.0147 -.0004 .0001 .0028 57.97 15.82 0 1.1585 '1902 -. 0193 -.0025 -.0013 .0002 57.81 16.81 0 1.1471 .2135 -. 0289 -.012 -' 0039 -.0015 57.59 17.81 0 1.144 .2439 -. 0449 -.0083 -.0004 -.000!3 57.48 18.B2 0 1.1575 .2634 -. 0487 -.DOSS .0002 -.0006 57.34 i9.8S 0 1.1992 .2889 -. 0579 -.0009 .0004 .0037 57.3 20 .as 0 1.1989 .3057 -. 05?8 -.0003 .0009 .0081 57.25

REYNOLDS NUMBER/FOOT• 1.36872£+06 MACH NUMBER= .2006

74

W :1: C H :1: T A ~:; T A T 1::: l.J N :1: V m: I~ $ :1: T Y "7 X ::1. () 1::· 0 D T 1. .. C) W ···· ~!) 1:> 1::: 1::: I> W :1: N X> T U N N m: 1. ..

Run nuMber 7 Static Tare 1000 DynaMic Tore 3000 const. table i

MARCH 21 1 :1.985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY Model upright

W.i.nd Axe• Datu

ALPHA PSI CL CD CM CRM CYM CY q <psf> < cltg) (deg >

-6.3 0 -.4267 '0345 -. 0128 -.0023 -.0002 .0002 59.43 ·-4.18 0 -.2574 . 026 -. 0048 -.0031 -.0001 .0004 S9.42 -2.06 0 -.oass '0213 ,0037 -.0031 0 .0004 59.22 .07 0 . 0963 . 0199 . 0115 -.0028 0 .0004 59.22 2.2 0 .2777 . 0227 .0186 -.0026 -.0001 .ooos 59.21 4.32 0 .4551 . 0293 ,0248 -. 0028 -.0001 .0007 59.21 6.4S 0 .6291 '0399 .0309 -. ooa? 0 .0009 59.21 8,56 0 .7949 .0541 .0362 -' 0025 . 0001 .0012 59.21 10.66 0 .9342 . 0?54 '0362 -.0008 -.0001 .0026 59' 13 12.72 0 1.0238 .1122 .0219 .0025 -.0001 .0022 58.92 13.76 0 1.0778 .1265 .0194 .0038 0 .0015 58,94 14.76 0 1.0745 .1541 .0092 -. 003 -.0001 .0041 58.53 :1.5.8 0 1.1264 .1715 .0056 .0023 -.0003 .0029 58.46 16.76 0 1.0735 .1966 -,0093 -.0123 -.0018 .0003 58.46 17.76 0 1.0697 .223 -' 0179 -.009 .0001 .0007 58.34

REYNOLDS NUMBER/FOOT= 1.36201E+06 MACH NUMBER .. . 2025

75

W :1: C H :1: T A m T A 'l" 1::: l.J N :1: V 1::: I~ ~=~ :1: T Y '7 X ::1. () 1::· (;) (;) T 1. •• 0 W ···· ~=~ P m: 1::: l) W :1: N X> T U N N 1::: 1...

Run nuMber 8 Static Tare 1000 DynaMic Tare 3000 const. tab 1 e 1

MARCH 21 t 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY Modlltl upriQht

W.ind Axes Data

ALPHA PSI CL CD CM CRM CYH CY q <psf) <deg> (deg>

-6.32 0 -.4577 .0391 .0009 -. 0019 -' 0003 . 0008 59.44 -4.21 0 -.2952 '0303 .0085 -.0024 -.0002 .0008 59.23 -2.09 0 -' 1251 . 0241 .0179 -.0026 -.0002 .0008 59.23 .03 0 .0443 .0211 .0266 -.0023 -.0003 ,0006 59.32 2.16 0 .22 ,0217 .0349 -.0027 -.0004 .0007 59.41 4.28 0 .3962 . 0269 .041 -.003 -,0004 .0008 59.21 6.4 0 .5698 .0358 . 04'73 -.0032 -.0003 .0009 59.31 8.52 0 .'7367 .0487 • 0523 -.0031 -.0002 . 0011 59.31 10.62 0 .8?3 . 0683 .053 -.002 -,0003 .002S 59.23 12.69 0 .9735 .103 .0383 .0014 -.0003 .0022 59.1 13.72 0 1.0153 .1201 .0327 . (}0 U:t -.0003 .0017 59.24 14.74 0 1.0502 .1439 . 0242 0 -. 0003 .0001 59 15.78 0 1.095? .1597 .0209 . 0013 -.ooos -.0001 58.94 16.75 0 1. 0644 .1795 .0109 -. 0116 -.0024 -.0014 58.81 17.77 0 1.0846 .1964 .0054 -. 0105 -.0027 -.0017 58.75 18.76 0 1.0694 .2182 -.0047 -.0022 -.0012 -.0016 58.63

REYNOLDS NUMBER/FOOT• 1.35343E+06 MACH NUMBER• .203

?6

WICHITA STATE UNIVERSITY '7 X :1. 0 1::· Cl (:) T 1... Cl W ···· ~;; P 1::: E X> W :1: N D T l.J N N 1::: 1 ...

Run nuMber 9 Static Tare 1000 DynaMic Tare 3000 c:onst. tab 1 e: 1

MARCH 21 1 :1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY Model upright

loHnd AU15 Data

ALPHA PSI CL CD CM CRM CYM CY q < p sf'> < deg > ( deg >

-6.35 0 -.5009 . 0454 .0094 -.0041 -.ooo5 .0008 59.45 ·-4. 24 0 -.3323 . 0352 .0184 -.0041 -.ooos . 0011 59.34 -2 I j,2 0 -.1669 . 0283 .0279 -.0037 -.0005 .001 59.54 0 0 .0041 .0243 .0376 -. 003 -.ooos . 0011 59.33 2.12 0 .1702 . 0234 .0466 -. 0032 -.0006 . 0011 59.42 4.25 0 .3478 . 0269 .054 -.0033 -.0007 .0012 59.42 6.37 0 .5165 .. 0341 .0604 -.0039 -.ooo5 . 0011 59.31 !3.48 0 .6813 .045 .0661 -' 0049 -,0003 .0006 59.21 10.6 0 .841 .0593 .0708 -.0053 -.0002 .0006 59.31 12.67 0 .947 '0942 .0566 -.0079 -. 0001 .0001 59.28 13.7 0 .983 .11 ,051 0 -.0002 .0015 59.22 14.74 0 1.0394 '1251 .0483 .0012 -.0003 -.0004 59.25 15.73 0 1.0351 .1532 .0316 -.0053 .0001 .0018 59.14 16.74 0 1.0389 .1702 .0261 -.0071 -.0006 .0002 58.89 17.73 0 1. 03 '1978 . 0114 -.0035 .0012 .0008 58.78

REYNOLDS NUMBER/FOOT= 1.34750E+06 HACH NUMBER= .2032

77

WICHITA STATE UNIVERSITY '7 X j, () F 0 D T 1 ... C) W ···· =::;: 1=> 1::: 1::: I> W :1: N X) T l.J N N E 1 ...

Run nuMber 10 Stat .i.e: T11re 1000 DynaMic Tr.1re 3000 c:onst. tr.lble 1

MARCH 21, 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY Model upright

W.Lnd Axes Data

ALPHA PSI CL CD CM CRH CYM CY q < psf> < deg > < deg >

-6.38 0 -.5356 . 0509 .0187 -.oo55 -.0007 .0007 59.26 -4.26 0 -.3646 . 0399 .0277 -.0051 -.0006 .001 59o35 -2 014 0 -0 1959 0 0327 o0373 -o0048 -o0006 o0012 59o25 -0 02 0 -.0267 o0283 .0484 -o0031 -o0006 o001 59.34 2 0 1 0 .1356 0 027 .o58S -. 0026 -.0008 o0009 511'o34 4.22 0 .308 0 0299 .0657 -.003 -,0008 o0007 S9o43 6o33 0 .4717 o0358 '0725 -.0037 -.0007 o0007 59o33 8o45 0 .6361 o04417' o0787 -.004 -.0006 o0009 59o32 10 0 56 0 .791 0 0574 '0837 -.004 -.ooo5 o0008 59o32 12o62 0 .8744 0 0928 o0657 -.ooo5 -.0007 .0022 59o2 13o67 0 o9457 01053 .065 .0003 -o0004 oOOOB 59 0 02 14o66 0 o9394 01322 o0504 -.0041 .0001 o0008 59.01 iS o 6°7 0 o9522 .1524 .0411 -. 0022 .002 .00:1.9 58o77 16o7 0 . 98;:!8 01731 o034 -.0004 o0038 .0046 58o52 i7o7i 0 1o0067 o1931 0 0224 -o0008 o0027 .0048 S8o48

REYNOLDS NUMBER/FOOT= 1o33873E+06 MACH NUMBER= .2027

78

W :1: t::: H :1: T A ~=~ T A T 1::: l.J N :1: V 1::: I~ !:~ :1: T Y "7 X :1. () 1::· () l:l T 1 ... 0 W -·· !:~ P 1::: 1::: I) W :1: N I> T U N N 1::: 1...

Run nuMber 11 Static Tare 1000 DynaMic Tare 3000 const, table 1

MARCH 21, 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY Model upright

Wind Axes DatQ

ALPHA PSI CL CD CM CRM CYM CY q Cpsf') (deg> (deg >

·-6 ,41 0 -.5783 .058 '0287 -,0059 -' 0008 .0012 59.57 -4.29 0 -.4111 .0468 '0368 -.0052 -.0006 .0013 59.37 -2.16 0 -.2321 '0382 '0474 -.0048 -,0005 .0015 59.26 -. 04 0 -.0615 .0333 . 0581 -' 0.035 -.0005 .0014 59,26 2. 06 0 .0905 .0316 .0678 -.0025 -.0006 .0014 59,35 4.19 0 .2677 .0335 '0769 -,0021 -.0008 .0017 59.44 6.3 0 .4266 '0386 .0838 -.0021 -.0007 .0017 59,34 8.42 0 .5901 .04?2 .09 -.0024 -.0008 .0018 59.24 10,53 0 .746 '0591 .0959 -.0029 -.0008 .0019 59.24 11.58 0 .821 .066? .0977 -' 0032 -.001 .002 59.44 12.6 0 .8517 .0891 .0844 -.0061 -.0007 .0012 59.3 13.64 0 .9019 .1052 .0787 ,-,0083 -.0009 .0021 59.13 14.64 0 .9011 '1263 .0681 -.0002 -.001 .0004 58,9 iS.64 0 .9041 .1511 .ossa -,0012 .0005 .0021 58.88 16.65 0 .9224 '1705 . 0456 -.0001 .0017 .0029 58.73 17.67 0 .946 .192 .0336 -.0031 .0003 .0034 58.59 18.66 0 .9285 .2106 .0249 -' 008 -.0013 .0004 58.36

REYNOLDS NUMBER/FOOT= 1.32449E+06 MACH NUMBER= .2025

79

W :1: C H :1: T A B TATE UN :1: V E I~ B :a: T Y '7 X ::1. 0 F Cl 0 T 1. •. 0 W ···· B P E a::: I) W :1: N l) T U N N a::: L

Run nuMber 12 Stat .i.e Tare 1000 DynaMic Tare ~5 () 0 0 const. table 1

MARCH 21, 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY MOdel upr.i.Qht

W.i.nd Axes Datl.l

ALPHA PSI CL CD CM CRH CYM CY q <psf> < deg) (de g)

-6.42 0 -.5963 . 0646 .037 -.0076 -.001 .0012 59.59 ·-4. 3l 0 -.4383 . 0533 .0449 -.0068 -.0008 .0014 59.48 --;.~. 19 0 -.2632 . 0443 . 0547 -.0067 -.0006 .0013 59.38 ·-. 06 0 -.0909 . 039 .065 -.0056 -.0006 . 0011 59.37 2. OS 0 .0712 . 0374 .0751 -.0044 -.0007 .0011 59.27 4. 17 0 .2351 . 0392 .0834 -.0037 -.0009 .0013 59.37 6.28 0 .3917 . 0435 ,0913 -.0028 -.001 .0013 59.36 8.39 0 .5502 . 0511 ,0979 -.0029 -.0009 .0013 59.16 i 0. 5 ' 0 .7094 .0622 ,1039 -.0029 -.001 .0015 59.36 12.57 0 .8111 .0929 . 0899 -.0065 -.0008 .0004 59.22 :1.3.61 0 .8582 .1 061 .0862 -.0079 -.0008 .0013 59. OS 14.61 0 .8666 .1273 .0754 -.0002 -.0013 .0006 58.92 :1.5.61 0 .8655 .1504 • 0645 -.0009 .0003 .0015 58.79 16.63 0 .8882 .1714 .0538 -.0012 .0005 .0019 58.65 17.64 0 .9058 .1905 .0442 -.0038 .0001 .0022 58.4 18.65 0 .9192 .2094 • 0347 -.0041 ·-.0009 .001 58.36

REYNOLDS NUMBER/FOOT= 1 .32114£+06 MACH NUMBER= .2025

80

WICHITA STATE U.NIVERSITY '7 X ::1. 0 1::· 0 D T 1 ... 0 W ···· B P t::: EX> W :1: N X> T U N N 1::: 1...

Run nuMber 13 Statlc Tare 1000 DynaMic Tart:! ::so 0 0 c onst, table 1

MARCH 21 1 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY ,.,odel upright

W.i.nd Axes DatiJ

ALPHA PSI CL CD CH CRH CYM CY q (psf> < deg > < deg)

-6.27 0 -.379 '0546 -' 0254 '073 .0034 .0026 59.5 ·-4' 1 s 0 -.2157 '0481 -' 0171 .0726 .0016 .0027 59.49 ·-2' 03 0 -' 038 '0447 -.0069 '0725 -.0006 .0032 59.49 .1 0 .1361 '0452 '0022 .0742 -.0028 .0036 59.49 ;; '21 0 .3018 .049 .0121 '0766 -' oos .004 59.29 4.34 0 .4773 .0566 .0199 .0776 -.0073 .0042 59.29 6.46 0 .6441 '0676 .0272 .0784 -.0096 .0042 59.19 8.57 0 .8031 '0812 .033 .0784 -.0116 .0042 59.29 10.68 0 .9591 '0987 .0375 .0777 -.0138 .004 59.29 12.74 0 1.0395 '1386 .02 .0722 -. 0157 .0031 58.99 13.77 0 1.0866 '1539 .0154 .0698 -.0165 .0041 58.82 14.76 0 1.0667 '1822 .0038 . 073 -.0175 .0043 58.62 15.73 0 1.03SS .2057 -. 0097 .0649 -. 0216 .ooos 58.6 16.73 0 1.0373 .2283 -.0194 .0621 -.0217 -.0009 58.37 l. 7. 74 0 1.0498 .2461 -. 0236 .0607 -.0219 -.001 58.32 18.75 0 1.0614 .2626 -' 0282 .0591 -' 0225 -.0021 sa. 17

REYNOLDS NUMBER/FOOT= 1,31333E+06 MACH NUMBER= .2022

81

W :1: G H :1: T A !:;: T A T m: l.J N :1: VIE I~ !:3: :1: T Y 7 X ~0. FOOT LOW-SPEED WIND TUNNEL

Run nuMber 14 Static. Tare 1000 DynaMic Tare 3000 const. table 1

MARCH 21 1 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY Model upright

lUnd Axes Data

ALPHA PSI CL CD CM CRM CYM CY q <psf> <deg> < deg)

-6.26 0 -.363 . 0538 -' 0284 .0703 .0031 .0023 59.5 ·-4' 14 0 -.194 '0469 -. 0211 . 0672 .0012 .0026 59.59 -2.02 0 -. 0288 . 0439 -. 0143 .0655 -.0007 .0031 59.49 .11 0 .16 .0442 -. 0068 • 0638 -.0027 .0034 59.39 ;.~. 24 0 .3435 .049 .00()1 .0618 ·-.0046 .0036 59.29 4 '~~7 0 .5252 . 058 .006 .0605 -,0062 .0037 59.29 6. 4'7 0 .6972 .0704 '0114 .0585 -' 0078 .0033 59.29 8.61 0 .8681 • 0862 ,0167 .0565 -.0091 .003 ·59. 29 10.73 0 1.0257 .1058 .0202 .0538 -.0104 .01123 59.19 12.79 0 1.1166 .1491 -.0009 .0474 -. 0118 .0014 58.99 13.8;;~ 0 1.1582 .1657 -.0034 .0443 -. 0126 .0017 58.83 14.82 0 1.1516 .1943 -. 0175 ,049 -.0132 .0006 58.62 . 15.84 0 1.1877 .2133 -, 0239 .0516 -.0136 -. 0005 58.66 16.79 0 1; 1151 .243 -. 0399 . 0372 -.0161 -. 0027 58.39 17.8 0· 1.1301 .2635 -. 0456 .0348 -. 0162 -. 0 033 58.34 18.81 0 1.1436 .28 -.0491 .0302 -. 0162 -. 0 037 58.19

REYNOLDS NUMBER/FOOT= 1.30688E+06 MACH NUMBER= .2022

82

W :1: C H :1: T A S T A T a::: l.J N :a: V a::: a~ m :a: T Y "7 X ::1. 0 1::· 0 C) T 1. •• CJ W -·· s; P a::: 1::: l) W :1: N l) T l.J N N a::: 1...

Run nuMber iS Static Tare 1000 DynaMic Tare 3000 c:onst. table i

MARCH 21 1 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY Model upriQht

Wind Axes Data

ALPHA PSI CL CD CH CRM CYH CY q <psf'> < deg) ( deg)

-6.29 0 -.4126 '0348 -.0141 -.0022 -.0003 -.0002 59.43 -4' 17 0 -.2431 '0269 -.0089 -.0027 -.0002 .0001 59.43 -2.04 0 -. 0562 '0223 -. 0025 -.003 -.0003 0 59.42 .1 0 .135 .0221 .0032 -.0032 -' 0003 .0001 59.42 2.23 0 .3242 . 0264 ,0087 -,0033 -, 0 0 OS .0002 59.32 4.35 0 .5009 '0342 .0148 -.003 -.0005 .0002 59.22 6.47 0 .6'709 '046 '0206 -.0024 -.0005 .0006 59.22 8.59 0 .835 .061 '0256 -.0018 -.ooos .0014 59.32 10.7 0 .9869 . 0797 .03 -. 003 -.0005 .0002 59.23 12.76 0 1 '0'736 '1211 . 0115 -.0066 -.0006 -.ooos 58.92 l.3 .8 0 1.1248 '1367 .0099 -.0077 -.0005 .0004 S8.9S 14.8 0 1.1367 '1605 -' 0023 -.0033 .0018 .0017 58.72 1.5.8 0 1.123S '1859 -.0131 -,0009 .ooos .0006 58.71 16.8 0 1' 1345 .2116 -, 025 .0036 .004 .0028 SB.SB 17.77 0 1.0896 .2328 -' 0299 -.0111 .ooos -.0024 58.37

REYNOLDS NUMBER/FOOT= 1,30797E+06 MACH NUMBER= .2025

83

W :1: C H :1: T A ~;) T A T m: U N :t: V 1::: I~ ~:; :1: T Y '7 X j, 0 F 0 0 T 1 ... D W ···· m P E 1::: l) W :1: N l) T l.J N N 1::: 1...

Run nuMber 16 Stat .i.e Tr.u·e 1000 DynaMic Tare ;30 0 0 const. table j_

MARCH 21' 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY Model upr.i.ght

W.i.nd Axes Data

ALPHA PSI CL CD CH CRH CYM CY q <psf) (cteg) <deg)

-6.31 0 -.442 . 0385 -.0039 -.0017 -.0004 .0003 59.44 -4, i 11' 0 -.2699 . 0301 .0014 -.0022 -.0003 .0003 59.43 -2' 06 0 -.0886 '0254 .007 -.0024 -.0004 .0002 59.33 .07 0 .1017 . 0243 .0123 -.0026 -.ooo5 .0003 59.43 2 I 21. 0 .2927 .0274 .0175 -' 003 -.0006 ,0005 59.32 4' 3:3 0 .4706 . 0344 .0221 -.0032 -.0006 .0007 59.42 6.46 0 .6544 . 0459 . 0263 -.0029 -. 0006 .0012 59' 12 8.S8 0 .8209 . 0604 .0301 -.0033 -.0006 .0011 59.12 j_O. 69 0 .9753 . 0789 .0335 -.0033 -.000'7 .0004 59.03 12.72 0 1.0218 '1256 '0111 . 0011 -.001 .ooos 58.76 13.75 0 1.0557 .1464 .003 ,0006 -.0011 -,0004 58.71 14.78 0 1.096 .1633 -.0022 .001 -.0011 -.0008 58.55 1S.8t 0 1.1442 .1802 -.0045 .0024 -.0012 -.0014 58.58 16.75 0 1.0623 .2103 -. 0232 -.0108 -. 0006 -.0029 58.31 17.76 0 1.0785 .2302 -. 0276 -.0112 -. 0 0 03 -. 0029 58.26

REYNOLDS NUMBER/FOOT= 1.30079E+06 MACH NUMBER= .2023

84

WZCHXTA STATE UNIVERSITY 7 X ~0 FOOT LOW-SPEED WIND TUNNEL

I~ un nuMber 17 Static Tare 1000 Dynaf''lic Tare 3000 const. tabla 1

MARCH 21, 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY "odel upr.ight

Wind Axes Data

AL.PHA PSI CL CD CM CRM CYM CY q <psf> < deg > < deg >

-6.34 0 -.4809 . 0442 .0043 -.0032 -.0006 .0004 S9.SS -4.21 0 -.2986 .0341 .0099 -.003? -.0006 .0004 S9.2S -2.08 0 -.1143 .0284 . 0163 -.0039 -.ooos .ooos 59.34 .OS 0 .0?63 .0267 . 0222 -.0039 -. 0005 .0004 59.24 2.18 0 .256 .0293 . 02?4 -.0045 -.0006 .0006 59.33 4.31 0 .4378 .0361 .032 -.0045 -,0007 .0006 59.23 6.44 0 .6183 .0465 .0355 -.0048 -.0006 .0008 59.23 8.56 0 .?942 .0608 .0385 -.004? -.ooos .0006 59.23 10.67 0 .9425 .0806 .0393 -.0061 -.0004 -.0007 S9. OS 12.71 0 .9995 .1301 .0104 -.0045 -.ooos 0 58.68 13.73 0 1.0357 .146 ,0089 -.0036 -.0007 -.0002 58.62 14.77 0 1.089 .160.9 .0061 -.0003 -.0009 -.001? 58.45 15.8 0 1.1283 .1792 -.000? .0025 -.0004 -.0024 58.49 16.73 0 1.037 .2094 -. 0159 -.0111 -.0006 -.0025 58.32 17.75 0 1.0608 .229 -.0212 -.0116 -.0006 -.0038 sa .11

REYNOLDS NUMBER/FOOT= 1.2969SE+06 MACH NUMBER= .2022

85

W~CHITA STATE UN~VERS~TY 7 X ~0 FOOT LOW-SPEED WIND TUNNEL

Run nuMber 18 Static Tare 1000 DynaMic Tare ~5000

const. table 1 MARCH 21, 1985

GARY KREPS - THESIS PROJECT OKLAHOMA STATE UNIVERSITY

l'lodel upright

Wind Axes Data

ALPHA PSI CL CD CH CRH CYM CY q <psf'> < deg > <deg>

-6.36 0 -.509 . 0494 .0133 -.0041 -.0007 .0006 59.46 -4.24 0 -.3373 . 0392 .0179 -.0042 -.0006 .0008 59.55 -2.11 0 -.1507 . 0323 .0245 -. 0044 -.ooos .0007 S9.3S .03 0 .0421 . 0295 .0309 -.0044 -.ooos .ooos S9,3S 2.16 0 .2281 . 0314 .036 -.0048 -.ooos .ooos 59.24 4 .2·~ 0 .4038 . 0374 .0413 -.0051 -.0006 .0005 59,34 6.41 0 .5836 .0472 .0454 -.oosa -.0006 .0001 59.34 8.54 0 .7613 .0608 . 0485 -.0061 -.0004 .0001 59.14 10.63 0 .8889 . 0904 .03'73 -.0036 -.0006 .0016 58.99 1.2 .69 0 .9804 .125 .0233 -.001 -.0006 .0004 58.97 13.72 0 1.0186 .1431 .017 -.0018 -. 001 -.0004 58.81 14.76 0 1.0684 .161 .0102 -.0019 -.0009 -.0008 58.65 15.·78 0 1.1053 .1799 .ooss .0017 -,0001 -.0011 58.5 16.73 0 1.0348 .2073 -. 0098 -.0067 0 -. 0002 58.31 17.74 0 1.051 .2255 -. 0137 -.0108 -.ooos -.0014 58.26

REYNOLDS NUMBER/FOOT= 1.29428E+06 MACH NUMBER= .2023

86

W :J: C H :1: T A ~::; T AT E l.J N :a: V a::: a~ S :t: T Y '7 X :1. () F D Cl T 1 ... CJ W ···· ~::;; P 1::: EX> W :1: N l> T U N N E 1...

I~ un nuMber 19 ~tatlc Tr.~.re 1000 DynaMic Tare ;:~ooo

const. table 1 MARCH 21 1 1985

GARY KREPS - THESIS PROJECT OKLAHOMA STATE UNIVERSITY

Model upright

W.Lnd Axes Data

ALPHA PSI CL CD CM CRH CYM CY q <psf> < deg) (deg>

-6.38 0 -.5368 '0562 .0213 -' 0041 -.0009 .0004 59,57 ·~4 '26 0 -.3685 .0454 '0263 -.0042 ,-.0007 .ooos 59.27 -;~' 13 0 -' 1799 .0372 .0327 -.0047 -.0007 .ooos 59.36 .01 0 .0104 '0338 .0386 -.0043 -.0007 .0004 59.36 2' 13 0 '1892 '035 .0442 -' 0045 -.0006 ,0004 59,36 4.26 0 .3718 '0404 .0493 -.0049 -.0007 .0004 S9.3S 6.39 0 .5503 '0499 ,0537 -.0051 -.0007 .0004 S9.3S 8.51 0 .7265 '0629 ,0571 -' 0048 -.0007 .0009 S9.3S 10,63 0 .883 '0806 .058 -,0054 -.0007 .0007 59.26 12.7 0 .9853 '1186 '0385 -. 0095 -.0006 .0012 59.25 13.69 0 .9744 '1436 .026 -.0024 -.0007 .0009 S9 .13 14.73 0 1.0319 '1591 .0224 -.0012 -.001 .0004 59' 06 15.76 0 1.0782 '1744 .0173 .0005 -.0009 -.ooo5 58.89 16.73 0 1.0266 .2058 -.0007 -.0047 .0013 .0008 58.81 17.73 0 1.025 .2229 -' 0068 -.0107 -.ooos -.0018 58.66

REYNOLDS NUMBER/FOOT= 1.29348E+06 MACH NUMBER• .203

87

W :1: C H :1: T A ~:; T A T m: U N :1: V 1::: I~ B :1: T Y '7 X :1. () F 0 0 T L D W ···· B 1:) E 1::: X> W :1: N l) T l.J N N 1::: 1. ••

Run nuMber 20 Stat i.e Tr.H·e 1000 DynaMic Tare ~~000 c onst. ttJble 1

MARCH 21 1 1985 GARY KREPS - THESIS PROJECT

OKLAHOMA STATE UNIVERSITY MOdel upright

Wind Axes Data

ALPHA PSI CL CD CM CRM CYM CY q <psf> < deg > <deg)

-6.4 0 -.5661 . 0645 • 0286 -.0042 -.001 .0012 59.3 -4.28 0 -.397 . 0529 . 0334 -.0042 -.0009 .0012 59.49 ... 2. 15 0 -.2184 '0441 .0392 -.0045 -.0007 .0009 59.38 -' 01 0 -.0206 .0392 . 0454 -.0045 -.0006 .0008 59.38 ;2. i1 0 .1602 '0394 . 0511 -.0047 -.0005 ,0006 59.37 4.24 0 .3379 '044 .0561 -.0052 -.0006 '0,007 59.27 6.37 0 .5187 .0525 .0609 -.0052 -.ooos .0003 59.27 8.49 0 .6869 '0652 .0646 -.0053 -.0004 -.0002 59' 17 10.57 0 .802 '0946 .0541 -.006 -.0007 .0002 59.23 12.63 0 .8894 '1285 .0392 -.003 -.0007 .0001 59.01 13.67 0 .9431 '1429 .0344 -.0018 -. 0008 -.ooos 58.94 14.7 0 .9925 .1572 .0321 -.0016 -.0008 -' 0014 58.87 15.74 0 1.046 .1742 .0264 .0001 -.0007 -.0019 58.7 16.75 0 1.0571 .2023 .0109 .0034 .0025 -.0003 sa.sa 17.74 0 1.043 .2249 -' 0009 -.003 .0013 -.0019 58.36

REYNOLDS NUMBER/FOOT= 1.28S2SE+06 MACH NUMBER= .2025

APPENDIX C

WIND TUNNEL COEFFICIENT PLOTS

88

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VITA~ Gary L.. ~:::repr:s

Candidate fer the Degree of

M~ster of Science

Thesis: AN AILERON DESIGN TO COUNTER ADVERSE YAW

Major Field: Mechanical Engineering

Personal Data= Born in Bristow~ Oklahoma, October 18~ 1948~ the son of Bill J. and Edna v. Kreps. Marri~d to Judy A. Charlson on October 20, 1973.

Education= Graduated from College High School, Bartlesville, Oklahoma in May 1966; received a Bachelor of Science degree in Mechanic~l and Aerospace Engineering from Oklahoma State University in July 1970; completed the requirements for the Master of Science degree at Oklahoma State University in July 1985.