/ •

7  ̧ ." .

/9.

l,ow-Speed Wind Tunnel Invesligafion of the Static Stability /vand ('onlrol Characteristics of an Advanced Turboprop

('onfiguration wilh the Propellers Placed Over The Tail

by

GRAIIAM SC()TT RIIODES

A thesis submitted to the Graduale Faculty of

North Carolina State University

in partial flnlfillment of the

requirements for Ihe Degree of

Master of Science

I)EPARTMENT ()F MECilANICAL AND AEROSPACE ENGINEERING

Raleigh

1990

APPROVED BY:

(NA_-&_L ___c,:5, ,Qi__ Lf'l_#-._Pt=7_ _IN [_ TUN:'ICL

[,_Vr aT].. TI_7 , C_ T>_" qI'ATIC L. TA';_IL[ TY A_:_"

;_90-_, S01. 1
Abstract

RHODES, GRAHAM SCOTT. Low-speed Wind Tunnel Investigation of the Static

Stability and Control Characteristics of an Advanced Turboprop Configuration with the

Propellers Placed Over The "Fail. (Under the direction of Dr. John N. Perkinsl

Aspart of a cooperative agreement between NASA Langley Research Center and

North Carolina State University, an exploratory wind-tunnel investigation was performed

in the 30x60 Foot Wind Tunnel to determine the low-speed static stability and control

ch,'u'acteristics into the deep-stall regime of an advanced turboprop aircraft with tile

propellers located over the horizontal tail. By this arrangement, the horizontal tail could

potentially provide acoustic shielding to reduce the high community noise caused by the

propeller blades. The current configuration was a generic turboprop model equipped with

1 foot diameter single-rotating eight-bladed propellers that were designed for efficient

cnfise operation at a Math Number of 0.8. The data presented here is static force data.

The effects of power on the configuration characteristics were generally favorable.

An arrangement with the propellers rotating with the outboard blades moving down was

found to have significantly higher installed thrust than an arrangement with the propellers

rotating with the inboard blades moving down. The primary unfavorable effect was a large

pitch trial change which occured with power, but the trim change could be minimized with

a proper configuration design.
Acknowledgements

I would like to thank my graduate advisor, Dr. John N. Perkins, for his support

throughout this quest. In addition, I would like to thank Joseph L. Johnson, Jr., and Paul

L. Coe, Jr.. both of the NASA Langley Research Center, for their support, interest, and

insight into the project. My appreciation goes to the tunnel technicians, who endured five

weeks of night shift, computer problems, and engineers. To the remaining people at the

Flight Dynamics Branch. I appreciate tile opportunity to spend a ye_ working with you.

The experience and information I acquired will prove to be a great asset during my future as

an Aerospace Engineer.

To my parents, thanks for the support and encouragement. Without your help

through the ye_s, I would never have made it this far.

ii
Table of C_ntenls

Acknowledgements ......................................................................... ii

Table of Contents ............................................................................ iii

List of Tables ................................................................................ v

List of Figures .......................................... :.................................... vi

Symbols ...................................................................................... viii

Abbreviations ................................................................................ x

Introduction .................................................................................. I

Model and Test Technique ................................................................. 6

Results and Discussion ..................................................................... 16

Characteristics of Wing Body ..................................................... 19

Basic Longitudinal Characteristics ...................................... 19

Effect of Trailing Edge Flaps ............................................ 22

Effect of Droop on Lateral and Directional Characteristics ........... 22

Effect of Flaps on Lateral/Directional Characteristics ................. 26

Longitudinal Characteristics of the Unpowered Configuration ............... 26

Effect of Elevator .......................................................... 26

Effect of Flaps ............................................................. 30

Lateral/Directional Characteristics of Unpowered Configuration ............. 30

Effect of Rudder ........................................................... 30

Effect of Ailerons .......................................................... 33

Effect of Power and the Direction of Propeller Rotation ....................... 33

Effect on Body/Pylon/Nacelle ........................................... 33

Effect on Longitudinal Characteristics of Full Configuration ........ 40

Effect on Lateral/Directional Ch,'u-acteristics of Full Configuration.. 43

Effect of Fillet Fairings on Propellers Rotating in Both Directions.. 46

Characteristics of Final Configuration ........................................... 46

Effect of Fillets on Full Configuration Without Power ............... 46

Effect of Power on Longitudinal Characteristics ...................... 50

Effect of Power on Lateral/Directional Ch,'uacteristics ................ 50

Effect of Flaps ............................................................ 54

Effect of Flaps on Lateral/Directional Characteristics ................. 54

Effect 54

Effect 57

Effect 60

of Rudder on Directional Characteristics.. ......................

of Elevator on Longitudinal Characteristics ....................

of Power on Trim Capability .....................................

iii
Tal)le of Contenls - continued

Effect of Ailerons on Lateral Characteristics ........................... 61

Engine-Ot|t Characteristics ............................................... 61

Comparison To Theoretical Results .............................................. 65

Comparing the Results .................................................... 65

Concluding Remarks ............................................................... 73

References ........................................................................... 75

iv
List of Tables

Table 1 - Model reference geometry ....................................................... 7

Table 2 - Coordinate For NLF(I)-0416 Airfoil Section ................................ 9

Table 3 - Strain-Gage Balance Characteristics ........................................... 17
List of Figures

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Fi gu re

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

l - Photograph of Over-The-Tail Advanced Turboprop Model ............... 10

2 - Three-view drawing of model ................................................. !I

3 - Model Mounted In 30x60 Foot Wind Tunnel ............................... 14

4 - Propeller Configurations Tested .............................................. 15

5 - Axis System For Forces and Moments ....................................... 18

6 - Effect of Droop on Lift of Wing/Body ....................................... 20

7 - Effect of Droop on Pitching Moment of Wing/Body ....................... 20

8 Wing Stall Pattern ............................................................... 21

9 - Effect of Droops on Wing/Body Drag ........................................ 23

I0 - Effect of Flaps on I.ift of Wing/Body ....................................... 24

11 - Effect of Flaps on Pitching Moment of Wing/Body ....................... 24

12 - Effect of Droops o11 Wing/Body Rolling Moment ......................... 25

13 - Effect of Droops on Wing/Body Yawing Moment ........................ 25

14 - Effect of Droops on Lateral Stability ........................................ 27

15 - Effect of Flaps on Wing/Body Yawing Moment .......................... 28

16 - Effect of Flaps on Wing/Body Rolling Moment ........................... 28

17 - Effect of Flaps on Wing/Body Directional Stability ....................... 29

18 - Effect of Flaps on Wing/Body Lateral Stability ............................ 29

19 - Effect of Elevator on Configuration Pitching Moment .................... 31

20 - Inboard and Full Elevator Effecliveness .................................... 31

21 - Effect of Flaps on Configuration IAft ....................................... 32

22 - Effect of Flaps o71 Configuration Pitching Moment ....................... 32

23 - Effect of Full Rudder on Configuration Yawing Moment ................ 34

24 - Lower and Full Rudder Effectiveness ...................................... 34

25 - Effect of Ailerons on Configuration Rolling Moment ..................... 35

26 Aileron Effectiveness ......................................................... 35

27 - Effect of Power on Drag of Body/Pylons/Nacelles ....................... 37

28 - Illustration of a propeller at angle of attack (exaggerated) ................ 38

29 - Illustration of crossflow effect ............................................... 39

30 - Effect of Power on Lift of Body/Pylons/Nacelles ......................... 41

31 - Effect of Power on Pitching Moment of Body/Pylons/Nacelles ......... 41

32 - Effect of Power on Configuration Lift ...................................... 42

33 - Effect of Power on Configuration Pitching Moment ...................... 42

34 - Effect of Power on Configuration Drag .................................... 44

vi
Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

l.,ist or Figures - conlinued

35 - Effect of Power on Directional Stability .................................... 45

36 - Effect of Power on Lateral Stability ......................................... 45

37 - Effect of Fillets on Drag and Installed Thrust .............................. 47

38 - Effect of Fillets on Full Configuration Lift ................................. 48

39 - Effect of Fillets on Configuration Pitching Moment ...................... 48

40 - Effect of Fillets on Configuration Drag ..................................... 49

41 - Effect of Power on Final Configuration Lift ............................... 51

42 - Effect of Power on Final Configuration Pitching Moment ............... 51

43 - Effect of Power on Final Configuration Drag .............................. 52

44 - Effect of Power on Yawing Moment ........................................ 53

45 - Effect of Power on Directional Stability. ................................... 53

46 - Effect of Flaps on Final Configuration Lift ................................ 55

47 - Effect of Flaps on Final Configuration Pitching Moment ................ 55

48 - Effect of Flaps on Directional Stability ..................................... 56

49 - Effect of Flaps on Lateral Stability .......................................... 56

50 - Final Configuration Full Rudder Authority ................................ 58

5I - Effect of Power on Rudder Effectiveness .................................. 58

52 - Effect of Elevator on Pitching Moment ..................................... 59

53 - Effect of Power on Elevator Effectiveness ................................. 59

54 - Effect of Power on Trim Capability ......................................... 62

55 - Effect of Ailerons on Rolling Moment ...................................... 63

56 - Effect of Ailerons on Yawing Monlent ..................................... 63

57 - Effect of Power on Aileron Effectiveness .................................. 64

58 - Engine-out Yawing Moment ................................................. 66

59 Engine-out Rolling Moment .................................................. 66

60 - Orthographic View of QUADPAN Panel Model .......................... 67

61 Comparison to Potential Lift ............................. .................... 68

62 - Comparison to Potential Pitching Moment ................................. 68

63 - Comparison to Potential Directional Stability .............................. 70

64 - Comparison to Potential Lateral Stability ................................... 70

65 - Predicted Power Effects on Lift ............................................. 71

66 - Predicted Power Effects on Pitching Moment ............................. 71

67 - Predicted Power Effects on Directional Stability .......................... 72

68 - Predicted Power Effects on Lateral Stability ............................... 72

vii
Symbols

b

C

c

Co

C_

CI_

Cl_a

CL

CLmax

C_

Cm

Cmct

Cm_e

Cn

CNp

Cn_

Cnsr

d

J

It

L/D

N

q

Re

Wing span, inches

Local chord, inches

Geometric average wing chord based on trapezoidal reference area, inches

Drag coefficient, Drag/qS

Rolling moment coefficient, Rolling Mornent/qSb

Lateral stability, dCi/d[3,, per degree

Aileron control effectiveness, dCi/dSa, per degree

Lift coefficient, Lift/qS

Maximum lift coefficient

Slope of the lift curve, dCL/dOt, per degree

Pitching moment coefficient, Pitching Moment/qSc

Slope of pitching moment curve, dCJdct, per degree

Elevator control effectiveness, dCm/d_, per degree

Yawing moment coefficient, Yawing Moment/qSb

Propeller normal force coefficient, Prop Normal Force/qS

Directional stability, dCn/d]3, per degree

Rudder control effectiveness, dCn/dSr, per degree

Propeller diameter, feet

Propeller Advance Ratio, J=V/Nd

Distance from CG to MAC t

Lift to Drag Ratio, Ct.JCD

Propeller Speed, revolutions per second

Wind tunnel free stream dynamic pressure, pounds per squ:u'e foot

Radial station on propeller blade, feet

Reynold's Number based on MAC, (p*V*MACJ/I.t

viii
S

St

T'

C

V

Vii

v,

(1

(Ib

_r

8_

P

Wing reference area, S=cb

Area of horizontal stabilizer, square feet

Thrust Coefficient, Thrust/qS

Wing tunnel free stream velocity, feet per second

Horizontal Stabilizer Volume Coefficient,Stir/S'MAC

Resultant air velocity due to free stream and propeller rotation, fps

Non-dimensional chordwise location of propeller disk

Non-dimensional vertical location of thnlst ,axis

Angle of attack, degrees

Propeller blade angle of attack at radial station r, degrees

Sideslip angle, degrees

Propeller blade angle at radial station r, degrees

Aileron deflection angle, degrees, positive right wing TED

Elevator deflection angle, degrees, positive TED

Flap deflection angle, degrees, positive TED

Rudder deflection angle, degrees, positive TEL

Air density, slugs per cubic foot

Coefficient of viscosity for air. pound-second per square foot

ix
Abbreviations

AC

ATP

CG

IBD

LE

MAC

MACt

OBD

(YI'I"

psf

RPM

SM

TED

TEL

Aerodynamic Center

Advanced TurboProp

Center of Gravity

Propellers spinning with InBoard blades moving Down

Leading Edge

Mean Aerodynamic Chord, inches

Mean Aerodynamic Chord of ttorizonlal Stabilizer

Propellers spinning with OutBoard blades moving Down

Over-'Ille-Tail propeller

Pounds per Square Foot

Rotational speed of the propellers, Revolutions Per Minute

Static Margin, 100%*ldCm/dCL]

Trailing Edge Down

Trailing Edge [.,eft
Introduction

The NASA/Industry Advanced Turboprop Project (ATP) was begun in 1978 to

investigate the possibilities of using fuel efficient adv,'mced turboprop aircraft in the place of

current turbojet and tt,rbofan transport aircraft. The driving force for the project was a need

to overcome the rapidly increasing fuel costs which occurred since the 1973 Middle East oil

embargo. Existing turbojet and turbofan aircraft have the advantage of relatively high

cruise speeds compared to turboprop aircraft. Current turboprop propeller blades are fuel

efficient at Mach Numbers below approximately 0.6 but lose this efficiency rapidly at

higher Mach Numbers largely doe to compressibility effects. A major aspect of the ATP

since 1978 was the design of new propeller blades which operate efficiently at the higher

cruise Mach Numbers of turbojet and turbofan aircraft. Tile emerging propeller designs

included highly swept blade planforms, large blade angles, and a large number of blades

per engine for both single-rotating and counter-rotating configurations. The new propellers

operate with an installed propulsive efficiency that is increased by approximately 10 to 20%

from turbofans at cruise Mach Numbers around 0.8. Flight tests have shown this

increased propulsive efficiency from the new propellers alone could provide a fuel savings

as I,-u'ge as 30%. When considering additional improvements in the areas of mechanical

engine components and efficient engine design, the potential fuel savings for an advanced

turboprop aircraft is at or above 50% compared to existing turbojets and turbofans. 1-3

One of the major problems surrounding aircraft powered by advanced turboprops is

the large level of noise produced by the propeller blades. The noise from the new blades is

primarily aerodynamic pressure field noise from the blades themselves and from structural

vibrations induced by the acoustic effects of the blades on the wing and fuselage. Thin,

swept propeller blades have reduced noise compared to straight propeller blades at low

operating speeds, but at high cruise speeds, the supersonic helical tip speed blades
maintain a high overall noise level. For a conventional wing-mounted-engine

configuration, the propellers would be located close to the fuselage and beside the cabin

area resulting in a high degree of cabin noise. In addition, the level of noise on the ground,

the community noise level, could be quite high during takeoff and enroule. It was desirable

to reduce both the cabin and community noise levels without the need for additional aircraft

structure or weight which would reduce the benefits of the advanced turboprop concept. 1-5

Previous theoretical and experimental work has shown that the wing could provide

a significant degree of acoustic shielding to reduce the cabin noise level to a more

acceptable level when properly arranged with the turboprop blades. This shielding is due

to the wing both blocking and reflecting sound pressure waves away from the cabin area of

the fuselage. 4,5 The propellers and wing would be arranged to reflect the noise above the

wing for a high-motmted wing and below the wing for a low-mounted wing, For

conventional wing-mounted engines, with the propeller disk roughly centered above and

below the wing, the direction of propeller rotation is important. A low wing aircraft would

require the outboard blades to be moving down to reduce cabin noise and a high wing

aircraft would require the inboard blades to be moving down to reduce cabin noise. In

either case, the wing sweep and chordwise location of the propeller disk are important

because the chordwise position where shielding begins varies with Mach Number. For a

high wing with the inboard propeller blades moving down, it is reasonable to expect a

reduction in community noise as well since the sound waves would be reflected above the

wing and away from the community.

From another point of view, an aft-fuselage-mounted-engine configuration would

place the propeller blades far behind the cabin to reduce the cabin noise level somewhat,

though flight tests indicate acoustic treatment would still be required to achieve acceptable

cabin noise levels. A tail-mounted engine configuration would require significant

supporting structure. Past research indicates the anaount of additional acoustic treatment
required may be reduced by carefully positioning the fuselage/tail mounting structure and

turning the propellers with the outboard blades moving down 5,6.

Because of the highly loaded blades of advanced turboprop propellers, the effects

of the propulsion system can be more severe than conventional turboprops. The present

investigation represents a cooperative agreement between NASA and North Carolina State

University to determine the propulsion system installation effects on the static stability and

control characteristics of a generic advanced turboprop commuter/transport configuration at

low subsonic speeds.

Tile present investigation combines the concept of acoustic shielding and aftfuselage-mounted-engines by their application on an advanced turboprop configuration

which places the propeller blades over the horizontal stabilizer. This configur_ltion is called

the Over-The-Tail (O'lVl ") advanced turboprop. For this case, the horizontal stabilizer

would, potentially, provide the aforementioned acoustic shielding to reduce ground and

community noise levels. Since the entire propeller would be located above the horizontal

stabilizer, the direction of propeller rotation should not greatly effect the degree of

community noise shielding, though the sweep of the horizontal tail and position of the

propeller disk would remain important. The cabin noise level would be reduced as well

because the propellers are l¢n'ated near the rear of the filselage, far behind the cabin.

Several previous wind-tunnel investigations have been performed on aft-fllselagemounted-engine advanced turboprop aircraft. All of these showed generally favorable

effects of power on the static stability and control of the aircraft. In particular, the aftmounted propellers increased longitudinal and directional stability with no major

detrimental effects on control effectiveness. The configurations generally had enough

rudder authority to overcome and trim the large yawing moment with an engine-out

condition 7-9.

An OTT design has several complicated aspects which must be considered. Since

the propellers would be located above the horizontal tail, it is possible the thrust line would

3
be locatedabovethecenterof gravity. For poweron cases,this would resultin a large

decrementin pitchingmomentwhichmustbebalancedin orderto trim theaircraft.

A major aerodynamic concern is caused by the propellers being located close to the

horizontal and vertical surfaces. Past research has shown a jet blowing over one side of a

lifting wing increases the circulation about the wing resulting in an increased lift I0. A

propeller located close to a wing should have a simil,'u" effect. For example, the propellers

blowing over the upper surface of the horizontal stabilizer would potentially increase the lift

of the horizontal stabilizer. This increased stabilizer lift would have a corresponding

decrement in pitching moment, an additional trim change with power. The increased

stabilizer lift and decrement in pitching moment vary inversely with the gap between the

propeller blade tip and the surface of the stabilizer.

The trim changes which appear inherent to the OTT configuration would place a

burden on the elevator and may significantly reduce the available pitch control for high

power settings. Most large transport aircraft have all-moving horizontal stabilizers, and an

O'VI" design may reqt,ire one in order to trim at cruise while maintaining sufficient elevator

pitch control I1 • In addition, there would be a high degree of trim drag for all flight regimes

other than landing, where there is no appreciable thnlst. The problems of pitch control and

trim drag can be remedied to a large degree by a proper configuration design. Ideally, the

O"Iq" configuration would have a low-mounted horizontal stabilizer with anhedral to move

the thrust line close to the center of gravity. In addition, the propeller should be angled

downward slightly to direct the thrust line through the center of gravity to completely

eliminate a pitching moment change due to thrust. The pitch change due to increased

stabilizer lift with power could be minimized with a large gap between the propeller and the

stabilizer.

The O'Iq" arrangement has at least one potential performance advantage. Past

rese,'u-ch has shown that a jet blowing over the upper surface of a lifting wing reduces the

pressure drag on the wing significantly. Research at NASA-Langley Research Center has

4
shownthat a propellerblowing overa surfacehasa simil_ effect12-14 With the O'l"I"

configuration,thisdragreductionwouldbeseenatcruiseasa reductionin trim drag. This

dragreductionwouldat leastserveto balancethetrim dragincreasediscussedabove. For

the properly designed configuration, there is potential for lower trim drag than a

conventionaltwo-surfaceturbopropaircraft.
Model and Test Technique

The present investigation was performed in NASA-Langley's 30x60 Foot Wind

Tunnel 15 and utilized the approximately l/9th scale generic turboprop model shown in

figure I. The model is a conventional two-surface wing/tail configuration with a unique

engine/nacelle location that places the propellers above the horizontal tail. The horizontal

stabilizer has no dihedral and no effort was made to place the thrust line close to the center

of gravity. Figure 2 is a three-view drawing of the model and table 1 lists the geometric

properties of the model. The model construction was conventional, utilizing fiberglass,

balsa wood, and aluminum for most components. The model is approximately 10 feet long

with a wing span of 9.91 feet. The wing's MAC is 14.0 inches. The wing airfoil is an

NLF(I)-0416 section. The airfoil coordinates are listed in table 2. The wing was equipped

with a 26 inch span leading edge droop. The 8-bladed propeller installation included the

high speed Hamilton Standard SR-7 blades3,16 which were constructed from carbon-fiber

composites to maintain physical integrity despite the low thickness of the sections. The

propellers were located such that the blade tips were separated from the horizontal and

vertical stabilizers by a distance of only 1/Sth inch. The propellers have a d of 1 foot and

the same scale as the configuration. The blades were set at a nominal blade angle of 43

degrees at the 75% radius station. The two propellers turned in opposite directions to

effectively eliminate propwash "swirl" effects on the configuration lateral and directional

trim characteristics.

Small air turbines were used to drive the propellers. The air was supplied by

flexible tubing from external pressure domes. The maximum supplied pressure was 250

pounds per square inch. A copper tube "trombone" was placed inside the model between

the air source and the air turbines to remove the effects of the high pressure air on the force

data. Several checkout data runs were performed in the absence of the air turbines to verify

that the trombone was effective.

6
Table 1 - Model Reference Geometry

Fuselage:

Wing:

ltorizontal

Vertical

Tail:

Tail:

l.ength ...................................... 8.82 feet

Maximum diameter ................ 11.2 inches

Body Station of nose ............... 0 inches

Waterline of CG ....................... 20.488 inches

Body Station of CG ................... 52.12 inches

Airfoil ....................................... NLF(1 )-(1416

Thickness ................................. 16%

Area .......................................... 9.91 square feet

MAC ........................................... 14.0 inches

Span .......................................... 108.86 inches

Aspect Ratio ............................. 8.3

Taper Ralio .............................. 0.376

Qt, arter-chord sweep ............. 1.41 degrees

Dihedral ................................... 4.0 degrees

Root incidence ........................ 1.5 degrees

Tip incidence .......................... 0.5 degrees

Body Station of AC .................. 52.12 inches from

Waterline of AC ....................... 18.77 inches

IrE break chord ....................... 15.57 inches

nose

Airfoil ....................................... NACA 64A010

Thickness ................................. 10%

Tail volume, Vt! ....................... 0.825

Area .......................................... 3.25 square feet

MACt .......................................... 12.71 inches

Span .......................................... 40.0 inches

Aspect Ratio ............................. 3.42

Taper Ratio .............................. 0.326

Quarter-chord sweep ............. 33.0 degrees

Dihedral ................................... 0 degrees

Root incidence ........................ 0 degrees

rip incidence .......................... 0 degrees

Body Station of AC .................. 93.9 inches from nose

Waterline of AC ....................... 23.61 inches

Area of inner elevator .......... 66.2 square inches

Area of full elevator .............. 116.44 square inches

Span Station of elev. break.. 12.45 inches

Airfoil ....................................... NACA 64A010

Thickness ................................. 10%

Area .......................................... 1.69 square feet

Mean Aerodynamic Chord .... 15.09 inches

Span .......................................... 16.9 inches

Aspect Ratio ............................. !.17

Taper Ratio .............................. 0.454

Quarter-chord sweep ............. 33.3

Body Station of AC .................. 93.17

Waterline of AC ...................... 34.5

Inner rudder area .................. 36.34

degrees

inches

inches

square inches
Table 1 continued

Vertical "Fail:

(continued)

Pylons:

Propellors:

Full rudder area ...................... 61.52 square inches

Waterline of rudder break...36.33 inches

Airfoil ....................................... NACA 64A010

Thickness ................................. 10%

Diameter ................................... 12.0 inches

Number of blades ................... 8

Blade type ................................ SR-7 ltamiiton Standard

Arrangement .......................... Single Rotating

Blade Sweep Angle ................. 41 degrees

75% Radius Blade Angle ........ 43 degrees

Body Station of prop disk ...... 94.53 inches from nose
Table 2 Coordinates For NLF(I)-0416 Airfoil Section

Upper Surface x/c

0

0.00509

0.01393

0.02687

0.04383

0.06471

0.08936

0.11761

O. 14925

O. 18404

0.22169

0.26187

0.30422

0.34839

0.39438

0.44227

0.49172

0.54204

0.59256

0.64262

0.69155

0.73872

0.7835

0.8253

0.86357

0.89779

0.92749

0.95224

0.97197

0.98686

0.99656

1

Upper Surface _,/c

0

0.01446

0.02573

0.03729

0.0487

0.05964

0.06984

0.07904

0.08707

0.09374

0.09892

0.10247

O. ! 04 25

0.10405

0.10162

0.09729

0.09166

0.08515

0.07801

0.07047

0.06272

0.05493

0.04724

0.03977

0.03265

0.02594

0.01974

0.014

0.00862

0.O0398

0.00098

0

Lower SuiIace x/c

0

0.0(X)49

0.00073

0.00709

0.01956

0.03708

0.05933

0.08609

O. 11708

0.152

0.1905

0.23218

0.27659

0.32326

0.37167

0.42127

0.4715

0.52175

0.57122

0.62019

0.67014

0.72107

0.77156

0.82012

0.86536

0.90576

0.93978

0.96638

0.9852

(/.99633

1

Lower Surface _,/c

0

0.00403

-0.00439

-0.01154

-0.01883

-0.02594

-0.03254

-0.03847

-0.04361

-0.04787

-0.05121

-0.05357

-0.05494

-0.05529

-0.05462

-0.05291

-0.05009

-0.04614

-0.04063

-0.0325

-0.02231

-0.01221

-0.00364

0.00278

0.00667

0.00792

0.00696

0.00478

0.00242

0.00065

0
ORIGINAE PAGE

BLACK AND WHITE PHOTOGRAPH

ORIGINAL PAGE IS

OF POOR QUALITY

Figure 1 - Photograph of Over-The-Tail Advanced Turboprop Model

10
48.62

52.12

=14.0

4O

_ot Horizontal

Stabilizer

"_.8_-

I

"42.1

= 41.05_

OI Vertical _ _

Stabilizer / "T [ I t "

_,_.o2

42.41 __

105.81

/

•'" 108.86 '_

Figure 2 - Three-view drawing of model

11
Figure 3 shows the model mounted in the wind tunnel on a C-strut/sting assembly

and equipped with an internal six-component strain+gage balance for measuring

aerodynamic forces and moments. The model attitude was adjusted in angle of attack and

sideslip to approximately account for flow angularity in the tunnel. Weight tare corrections

were applied to all force data so that the representative data is purely aerodynamic. No

other corrections were applied to the data 17. In addition to force data, limited smoke flow

visualization was performed to visualize the flow field surrounding the propeller assembly.

Also, the wing and tail were tufted in order to determine their stall patterns. All data was

gathered at a q of 4 psf for a Re of 43(I,490 based on the MAC.

Angular brackets were used to set control surface deflections to measure control

effectiveness and authority. The wing trailing edge flaps were tested at 6f=0 and 35

degrees, and additionally at 8f---20 degrees during the power off model build-up runs. The

ailerons were tested at 8a=0, + 10, and +20 degrees. The rudder was split into an upper

rudder and lower rudder of approximately equal areas. The full rudder was tested at 6r =-

30,-20,-10, and 0 degrees. In addition, the lower rudder was tested independently at the

same angles. Similarly, the elevator was split into an inboard elevator and an outboard

elevator. The full elevator was tested at 5e=-30,-20,0,20, and 30 degrees while the

inboard elevator was tested independently at -20,0, and 20 degrees.

The propulsive effects were tested for propeller RPM's of 3500 rpm and 4500 rpm.

For the complete configuration, the 3500 rpm case provided an installed thrust coefficient,

T'c, of approximately 0.2 for a J of 0.995. The 4500 rpm case provided a T'c of

approximately 0.36 for a J of (}.774. The T'c=0.2 cases with 8r=35 degrees were

representative of a 3 degree takeoff climb and the Tc=0.36 cases with 8f=35 degrees were

representative of a missed approach full power climb. Variations in T'c due to configuration

changes represent part of the importance of the investigation. The propellers were tested

for cases with the propeller blades spinning IBD and OBD. In addition, fillet fairings were

+

tested to close in the inboard area between the tail surfaces and propellers as an attempt to

12
ducttheflow behindthe propellers and produce more thrust. Figure 4 shows the different

propeller configurations.

In addition to the wind-tunnel investigation, an analysis of the basic configuration

was performed utilizing the subsonic potential panelling code, QUADPAN 18. The

QUADPAN code is a doublet surface panel method which has the capability to model

propeller slipstream effects using a simple slipstream model. Such a panelling code is a

quick, relatively simple, means to calculate aerodynamic characteristics of complete

configurations for prelimin,'u-y design and analysis 19,20.

13
ORIGtNAE PAGE

BLACK AND WHITE PHOTOGRAPH

Figure 3 - Model Mounled In 30x60 Foot Wind Tunnel

ORIGINAL PAGE IS

OF POOR QUALITY

14
OBD Without Fillets OBD With Fillets

IBD Without Fillets IBD With Fillets

Figure 4 - Propeller Configurations Tested

15
Results and Discussion

All longitudinal forces and moments are referenced to the stability axes and all

lateral/directional forces and moments are referenced to the body axes. Figure 5 shows the

axis system graphically. The moment reference center was located 52.12 inches from the

nose of the model which corresponded to the 25% chord station of the wing's mean

aerodynamic chord.

Table 3 shows the characteristics and accuracy of the internal strain-gage balance

used to measure the aerodynamic forces and moments. "lq_e accuracy shown was calculated

for a steady static load, and is the minimum expected accuracy of the balance. In the wind

tunnel, the data acquired with the wind on must be time averaged or filtered because the

flow is somewhat unsteady due to free stream turbulence and vibrations which occur on the

C-strut model support. The longitudinal forces and moments are quite large and do not

vary much due to this unsteadiness; however, the lateral and directional forces and

moments are relatively small when the model is at a small or zero sideslip. The effect of

unsteadiness on these forces and moments is measurable. Due to limitations in the data

acquisition system at the 30x60 Foot Wind Tunnel, this unsteadiness is not completely

removed when the data is filtered. For this reason, the lateral and directional data presented

here will appear scattered. This inaccuracy is most obvious in the static stability

derivatives. For this reason, only large changes in these derivatives are considered

important.

Power effects were measured for the propellers spinning at 3500 rpm and 4500

rpm. The effective installed thrust varied somewhat depending on the configuration. For

this reason, all figures reference the speed of the propellers rather than the Tc. For cases

where Tc is important, T'_,is presented in the text.

16
Table 3 Strain-Gage Balance Characteristics

Componeni

Normal Force

Axial Force

Side Force

Pitching Moment

Yawing Moment

Rolling Momcnt

Minimum Load

-4(X) pounds

0 pounds

-2(X) pounds

-200(/inch-pounds

-2000 inch-pounds

- 1230 inch-pounds

Maxinlum Load

40i) pounds

200 pounds

20() pounds

20(1) inch-pounds

2000 inch-pounds

1230 inch-pounds

Accuracy

+/- 2 pounds

+/- 1 pound

+/- 1 pound

+/- I0 inch-pounds

+/- I0 inch-pounds

+/- 6.15 inchpound

Forces

CoeJficient

CL

q)

Cy

C,.

Cn

CI

and Moments in Coefficient

q=4 pounds per square foot

'Minimum

-lO.l

if)

-5.05

-3.604

-0.4635

-0.285

Maximum

10.1

5.05

5.05

3.604

0.4635

0.285

Form

A ccuracy

+/- 0.13505

+/- 0.0252

+/- 0.0252

+/- 0.018

+/- 0.00232

+/- 0.0() 143

17
CL

Wind v

Z/ ,.,n

Wind

Figure 5 - Axis System For Forces and Moments

18
Characteristics of Wing Body

Basic Longitudinal Characteristics. Figures 6 and 7 represent the longitudinal

characteristics of the wing/body with pylons and nacelles including the effect of a 26"

outboard leading edge wing droop. The lift curve of the wing and body without the droop

exhibits a slope, Cta of 0.085 per degree with a Ckrnax of 1.44 at an ot of 14 degrees. The

stall point is quite sharp and the loss of lift after stall is significant. A sharp drop-off in lift

at stall has been shown to be characteristic of aircraft with poor roll departure tendencies at

stall. The roll departure problems are due to non-symmetric stall patterns on the two wings

when the aircraft is in sidcslip, which results in one wing producing more lift than the other

causing the aircraft to roll into a spin or spiral. If the drop-off in lift at stall were small, the

difference in lift between the wings would not be large enough to cause violent roll

problems. However a large drop-off in lift indicates a possible large difference in lift

between wings with a correspondingly large rolling moment. A stall pattern which begins

at the wing tip would enhance this rolling moment due to a large lever arm for the

differential lift. The leading edge "droop" modification is intended to reduce this problem

by providing improved flow over the wing tips in two manners. First, the shape of the

droop improves the two-dimensional stall characteristics of the airfoil near tile tip,

postponing the tip stall. Second, the sharp droop edge produces a strong vortex which

sheds across the wing providing energy to the flow and acts as an _lerodynamic fence

which reduces spanwise flow, a contributor to tip stall and the roll departure problem 21.

With the droop installed, Ct_ does not change, while CLmax decreases slightly to 1.42 at

an ct of 14 degrees. In fact, the lift curve is flatter at the stall point. "lhere is still a large

drop-off in lift after stall, but an increase in lift into the deep post-stall ot range indicates the

droop is effective in improving the flow over the wing showing potential for some

reduction in rolling moment asymrnetries at higher o_'s. Figure 8 shows the wing planform

and stall patten_ with the leading edge droop mounted. It is seen that the flow at the wing

tip remains un-stalled until higher alphas, indicative that the droop is effective. The droop

19
1.75

EL

1.50

1.25

1.00

0.75

0.50

0.25

0.00

-5

"o I as'cw'n° Droop Attached I

I I " I I " I I I " =1

0 5 10 15 20 25 30 35 40

et (degrees)

Figure 6 - Effect of Droop on Lift of WingBody, 6f=O °

0.15

0.10

0.05

Cm 0.00

-0.05 -

-0.10 - -

0

-0.15

-5

Basic Wing

o

• I • I • I ' I ' I

5 10 15 20 25

ot (degrees)

Droop Attached I

l I ' i

30 35 40

Figure 7 - Effect of Droop on Pitching Moment, 6t=O °

2O
18

12 °

0

32 °

10 °

28

12 °

0

Figure 8 Wing Stall Pattern

21
is expected to improve roll tendencies at stall somewhat, but not to eliminate the problem

entirely since the initial lift drop-off is still large. Without the droop, the pitching moment

varies linearly with ct until stall. At stall, there is a stable break in Cma immediately

followed by an unstable break. The leading edge droop has little effect on pre-stall pitching

moment, but somewhat flattens out the breaks in Cma at stall. The effect of the droop in

the deep post-stall ct range is simply a shift in the curve. There are no major

improvements, but there are no detrimental effects either. Figure 9 shows is the drag curve

for the configuration with and without the droop. There is no significant drag penalty with

the droop attached. All remaining data will be presented with the leading edge droop

installed.

Effect of Trailing Edge Flaps. Figures 10 and 11 represent the effect of the wing

trailing edge flaps on the wing/body with pylon and nacelles. As expected, a flap

deflection causes an increment in the lift curve due to increased circulation about the wing.

The average increase in CL prior to stall is 0.39 due to a 35 degree flap deflection. The

maximum lift with 35 degrees of flaps increases to Ccmax=l.59 at an ot of I0 degrees. The

stall pattern is more abrupt with flaps. The pitching moment curve shows a decrement in

Cm due to a flap deflection. This result is expected and is caused by an increase in the aft

loading on the wing when the flap is deflected, behind the center of gravity. The average

change in Crn due to a flap deflection of 35 degrees is -0.038.

Effect of Droop on l_z_teral and Directional Characteristics. Figures 12 and 13 show

the effect of the leading edge droop on CI and Ca. At small angles of attack, both with and

without the droop, CI changes slightly with ft. This tendency is to be expected to some

extent with any configuration, due to characteristics of the configuration, irregular freestream flow and misalignment of the model with the free-stream. At high angles, without

the droop, there is a jump in CI at an cL of 20 degrees which indicates the possibility of roll

departure. With the droop, a jump in CI near st_dl still exists at higher alphas, but the jump

is smaller, showing the effectiveness of the droop. There is a slight shift in Cn which is

22
0.8"

0.6

CD 0.4

0.2

0.0

-,5 0

- I o With Droop op I

I I " ! " I I ' I I " !

5 10 15 20 25 30 35 40

_t (degrees)

Figure 9 - Effect of Droops on WingBody Drag, 6r=O °

23
1.75

1.50

1.25

1.00

EL 0.75

0.50

0.25

0.00

-5

O

°I

o

-----o---- No Flaps

G - 35 Deg Flaps

I " 1 " I " I " I

1 0 1 5 20 25 30 35

(_ (degrees)

i

'1

40

Figure I0 - Effect of Flaps on Lift of WingBody

Cm

0.15

0.10

0.05

000

-0.05 "

-0.1o - -

o

o 15 - d I ¢

-0.20

-5 0

35 Deg Flaps

I " I " I " I ' 1 ' 1 " I " |

5 10 15 20 25 30 35 40

o_(degTees)

Figure 11 - Effect of Flaps on Pitching Mome.t of WingBody

24
C!

0.08 -

0.06 -

0.04 -

0.02 -

0.00

l

-0.02 _-

i

b.

-0.04 -

-0.06 -

-o.o8i •

-5

No Droop

With Droop

• I " 1 " I ' I " I " ! " I " g

5 10 15 20 25 30 35 40

o_(degrees)

Figure 12 - Effect of Droops on WingBody Rolling Moment, 6f=O °

006 -

004 -

002 -

Cn 0.00- _"

-0.02 -

-0.04 -

-0.06 ' ,

-5

Figure 13

N°°'°°° Wilh Droop I

• I w I ' I • ! ' I ' I " I I

5 10 15 20 25 30 35 40

Alpha (degrees)

Effect of Droops on WingBody Yawing Moment, 6r=O °

25
most likely due to the droops being slightly asymmetric. Another indication of the

effectiveness of the droop is the increase in lateral stability (-Cl[_) shown in figure 14 in the

post-stall angle-of-attack region, ot above 14 degrees.

Effect of Flaps on LateralDirectional Characteristics. Figures 15 and 16 show the

variation in Cn and CI with and without flaps deflected. As expected, the flaps have little

effect on Cn. The largest contribution to the yawing moment variation is the body, pylons,

and nacelles, all of which have more vertical area ahead of and behind the CG. There is a

noticeable affect on Ci. In the pre-stall region, the flaps provide an increment in Ci

showing possible asymmetries in the model or free-stream flow. With flaps, the jump in

Ci in the post-stall region is increased from the no flaps case. This shows possible

asymmetric flap deflections and possibly a higher tendency to autorotate at stall with the

flaps deflected. This result was anticipated to some extent by the increase in loss of lift at

stall with the flaps deflected. Figures 17 and 18 show the effect of the flaps on directional

and lateral stability. As expected, the basic wing/body is directionally unstable (negative

Cnl 0 through most of the range of or. The flaps increase the directional stability slightly

through most of the ot range because there is a slight increase in the vertical area behind the

CG when the flaps are deflected. The flaps are shown to increase the configuration's

lateral stability.

Longitudinal Characierislics of the Unpowered Configuration

Effect of Elevator. The elevator for the configuration was split into two halves of

nearly equal area to deten'nine the effect of propeller blowing on the control surface

effectiveness. The resuhs presented in this section compare the full elevator to only the

inboard elevator without power as a reference to later results with power. Figure 19 shows

the configuration pitching moment curve with a variety of elevator settings. In the linear

range of alphas for no elevator deflection, Cma is -0.0921 which corresponds to a SM of

14.9%. For each case there is an increase in longitudinal static stability after stall because

of the loss of positive Cm from to the wing. The graph shows the full elevator is capable of

26
0.000

-0.001

-0.002

-0.003

C1_, -o.oo4

-0.005

-0.006 -

-0.007 -

-0.008

-5

• ! !

0 5 10

l___ N°°'°°p With Droop I

1 I " ! I l " I

15 20 25 30 35 40

(degrees)

Figure 14 Effect of Droops on WingBody Lateral Stability, 6r:O0

27
0.06

0.04

0.02

Cn 0.00

-0,02 -

-0.04 -

-0.06

-5

Jt'}_

• 1 I

0 5 10

35.oF,aps Deg Flaps 1

I j • I I " I • "1

15 20 25 30 35 40

o_ (de_ees)

Figure 15 - Effect of Flaps on WingBody Yawing Moment, 5r=O °

0.08

0.06

0,04

0.02

CI 0.00

-0.02 -

-0.04 -

-0.06 -

-0.08

-5

I _.___.__ NO35 Deg FlapsFlaps

' I ' | " ! " I " ! " I " | " I

0 5 10 15 20 25 30 35 40

ot (degrees)

Figure 16 - Effect of Flaps on Wing/Body Rolling Moment, _r=O °

28
C._

0.006

0.004

0.002

0.000

-0.002

-0.004

-5

_-----o--- 35o Deg FlapsFlaps 1

• ! " I " I " I I | " I • |

0 5 10 15 20 25 30 35 40

ot (degrees)

Figure 17 - Effect of Flaps on WingBody Directional Stability, 6r=0 °

O.OO2

0.000

o,,,,

-O.O02

-0.004 "

-0.006 -

-0.008

-5 0

• I !

5 10

[_ No35 Deg FlapsFlaps

I I I " I I " i

15 20 25 30 35 40

ot (degrees)

Figure 18 - Effect of Flaps on Wi.g/Body Lateral Stability, 6r=O °

29
trimming the aircr,'fft to approximately 22 degrees _ without power. The effect of elevator

is basically a vertical shift in Cm, with some loss of effectiveness at higher angles. Figure

20 shows the elevator effectiveness for the full elevator as well as the inboard elevator.

The effectiveness for both cases decreases smoothly ihrough the ot range with an anomaly

at stall. The inboard elevator provides an average of 57% of the effectiveness of the full

elevator. This percentage corresponds to the area ratio of the inboard elevator area to the

full elevator area, which is 0.569.

Effect of Flaps. Figure 21 shows the effect of the wing trailing edge flaps on the

configuration. As expected, the flaps cause an increment in lilt through the linear range of

t_'s. There is no change in the Ct_. The CLmax shifts from 1.52 at 14 degrees _ without

flaps to 1.62 at 10 degrees ¢t with 35 degrees of flaps. The stall pattern with flaps is

somewhat sharper than without flaps. This p,-uallels the results found for the wing/body

configuration. The flaps cause a nose-up pitching moment increment through most of the

o_range as indicated by figure 22. This effect is somewhat unexpected. Normally a nosedown change in pitching moment is expected because the flaps cause an increase in the aftloading, behind the CG, on the wing. In fact, referring to figure 11, this is the result for

the wing/body configuration of this model. The change in the effect of the flaps on

pitching moment is due to the downwash effects on the horizontal tail. With the tail on,

there is a positive shift in Cm due to the following: An increased downwash from the wing

with flaps increases the down-loading on the horizontal tail. This increased down-loading

is acting on a surface with a large VH, with a corresponding large shift in C m. For this

configuration, the positive Cm change due to down-loading of the horizontal tail is larger

than the negative Cm change due to the flaps, resulting in a net increment in Cm with flaps.

i_,ateral/Directional Characteristics of Unpowered Configuration

Effect of Rudder. The rudder for the configuration was split into two halves for the

same reason the elevator was split. Here, the lower half of the rudder would have an

increased effectiveness due to the propeller blowing. The results presented in this section

30
/

0.00

Figttre 20
CL

1,75

1.50

1.25

1.00

0.75

0.50

0.25

0.00

-5

d

-----o---- No Flaps

35 Deg Flaps

| 1 " 1 " | " ! 1 " |

5 10 15 20 25 30 35 40

ot (degrees)

Figure 21 - Effect of Flaps on Configuration Lift

Cm

02[

0.0 __1

-0.2 '

-0.4 -

-0.6 "

-0.8 -

-1.0

-5

[_ No 35 Deg FlapsFlaps

I | 1 " I " ! |

0 5 10 15 20 25 30

c_ (degrees)

a • i

35 4O

Figure 22 - Effect of Flaps on Configuration Pitching Moment

32
comparethe full rudderto only the lower rudder without power as a reference to later

results with power. Figure 23 shows the effect on Cn for a neutral rudder, full positive,

and fifll negative rudder. It appears the rudder authority is constant through the useful

range of or's, and beyond stall to approximately 16 degrees or. Above 16 degrees, the

authority drops off smoothly. Figure 24 shows a comparison between the full rudder

effectiveness and the lower rudder effectiveness. The lower rudder provides an average of

62% of the effectiveness of the full rudder, and this ratio corresponds approximately to the

to the ratio between the lower rudder ,area to the full rudder area which is 0.59. This plot as

well shows the rudder effectiveness is constant to approximately ! 6 degrees oz. The rudder

loses effectiveness above the stall ot because the vertical fin is embedded in the turbulent

wake behind the wing and filselage.

Effect of Ailerons. Figure 25 shows the rolling moment variation with neutral

ailerons, 10 degrees ailerons, and 20 degrees ailerons. Figure 26 shows the aileron

effectiveness through the angle-of-attack range. The ailerons are most effective at zero

degrees angle-of-attack, and the effectiveness decreases approximately linearly with or.

This reduction in aileron effectiveness with ot is likely due to trailing edge separation.

There is an anomaly at stall for both graphs which is probably due to model asymmetries.

The slight increase in aileron effectiveness at 35 degrees ot is another anomaly which was

repeated by several different runs.

Effect of Power and the Direction of Propeller Rotation

Effect on BodyPylon/Nacelle. Figure 27 shows the effect of power on the drag of

the body alone with the pylons and nacelles. As expected, with power, the drag shifts to

negative through most of the ot range. This shows that there is enough thrust to induce

acceleration. The thrust coefficient of the inst,'dled propellers at zero or, T'c, is calculated as

T'c=(CD_,Vith Power-(CDJProps Of[. At any angle of attack, a negative value of (CDJWith Power-

(CD)l'rops Off is an indication that the propellers are producing thrust. At a given alpha for

two test cases, the lower value of (CD)with Power-(CD)Props Off represents the case with the

33
0.06 -

0.04 -

0.02 -

Cn 0.00 -

-0.02

I

-0.04

O-

-----o---- Neutral Rudder

30 Deg Rudder

"_30 Oeg Rudder

-0.06

-5 0 5

! " I " I ' I ' I ' I " I

10 15 20 25 30 35 40

ot (degrees)

Figure 23 - Effect of Full Rudder on Configuration Yawing Moment, 6¢=35 °

0.0000 = ' '

z

-0.0005

I

!

_CnSr -0.OOLO

I

-0.0015

o-¢

-0.0020 • ( , • ,

-5 0 5 10

---.--a---- Full Rudder

.--.-o---- Lower Rudder

I ' I " I " I " I " I

15 20 25 30 35 40

(degrees)

Figure 24 - Lower and Full Rudder Effectiveness, 6¢=35 °

34
0.10

0.05

CI o.oo- -_

-0.05 -

-0.10 , '

-5 0

Aileron Deflection

(Degrees)

0

o 10

A 20

• I I I " I ' I ' I I

5 10 15 20 25 30 35

ot (degrees)

!

4O

Figure 25 - Effect of Ailerons on Configuration Rolling Moment, _f=35 °

0.000

-0.001

CiSa -0.002

-0.003

-0.004

-5

- I

I

I

i

- !

|

I

I

|

|

|

- !

I

-I

• !

0 5 10 15 20 25 30 35

ot (degrees)

Figure 26 - Aileron Effectiveness, 6r=35 °

35

4O
higher T'c. From the graph, it is seen that at higher angles of attack, the case with the

propellers rotating down outboard has much more installed thrust than the case with the

propellers rotating down inboard. This is due to a "crossflow" effect which occurs with

propellers at angle of attack. Figure 28 shows a propeller at a positive angle of attack. Due

to the free stream flow and the rotation of the propeller, the down-going blade operates at a

higher _ than the up-going blade 22. This means the down-going blade is more highly

loaded than the up-going blades, and there is a higher pressure change across the downgoing blades than across the up-going blades. This results in a "crossflow" behind the

propeller disk from the down-going side (high pressure) to the up-going side (low

pressure). Figure 29 shows that a diverging crossflow exists when the propellers rotate

IBD, and a converging crossflow exists when the propellers rotate OBD. This results in a

higher boat-tail drag for the IBD case compared to a lower boat-tail drag for the OBD case.

Evidence of this crossflow is provided by the following. Previous investigations present

results of a nacelle side force and yawing moment produced with a tractor propeller

configuration l6, 23. These forces and moments are caused by unequal pressure forces on

either side of the nacelle, a characteristic of the crossflow phenomenon. Additionally, one

report mentions a propeller tip free-vortex which moves away from the propeller centerline

on the side of the up-going blades 24. For a two-propeller installation, vortices on opposite

sides of the configuration would converge to the body with the propellers rotating OBD and

would diverge with the propellers rotating IBD. Since a free vortex follows the local free

stream velocity, this motion might well be evidence of the crossflow.

The difference in boat-tail drag between the two direction of rotation cases appears

as a large difference in installed T'c between the two cases. Theoretically, two free

propellers at zero angle of attack rotating at the same speed would have the same thrust,

regardless of direction of rotation. Looking back at figure 27, the T'¢ for the IBD case at

4500 RPM is 0.41 and the T'c for the OBD case is 0.45. The difference is due to slight

angularity in the propeller inflow which causes the propellers to be at incidence relative to

36
0.2 F

I

L

0.1 -

0.0

-0.1

CD

-0.2

-0.3

-0.4 -

-0.5

No Power

OBD 4500 RPM

_, IBD 4500 RPM

i

I I ' I I I | I " I

- 5 0 5 1 0 1 5 20 25 30 35 40

ot (degrees)

Figure 27 - Effect of Power on Drag of BodyPylonsNacelles

37
/

N------Pm_cUer Rotation Speed

(rsdians per second)

r_Radial Station [.ocation

_RF N_Velocity due to

propellerrotation

V_Fr_e-stn_sm Ve[0city

V r _L°cal Blade Air Velocity

t/sin (Ct},,_Compontmts of Free

Vcos {or)/stream into Prop Disk

{X-------Nacelle Angle of Attack

Otb_[.ocal BladeAngle of

Attack at Radial Station r

_b_Local Blade Angle at

Radial Station r

Vco_(a)

Down-Going Blade

Vcos(ct) I

°tb =[3b-tan-! {2nr_((l) !

Up-Going Blade

{ vcos(oO I

C(b "_3b" lanl (2nrN-Vsin(tx)/

(fib }Down_Goini > (Orb }Up_Golng

Figure 28- Illustration of a propeller at angle of attack (exaggerated)

38
Propeller Disk

v

Diverging Crossflow

ge Propeller Wake Area Resulting in High Drag

Propellers Rotating IBD

Propeller Disk 1

_onverging Crofsflow

_ Small Propeller Wake Area------_

Resulting In Low Drag

Propellers Rotating OBD

Figure 29 - Illustration of Crossflow Effect

39
the flow even when the configuration is at zero ct. This angular inflow is an effect of the

wake and downwash of the nacelles and pylons ahead of the propeller disk.

Figures 30 and 31 show the effect of power on CL and Cm for the body alone with

the pylons and nacelles. The body itself produces a small amount of lift which increases

with or. The addition of power causes an increase in the slope of the lift curve. This

increase in lift comes from the component of thrust acting in the direction of lift at t_ and

from the propeller normal force which occurs at ct. A small portion of the additional lift is

due to an improved flowfield over the pylons with power. Plotted also on the CL plot is the

basic body lift added with the lift due to the thrust vector alone, with T'c calculated from the

drag data. The remaining lift above this line is due to the improved flow over the pylon and

the propeller normal force. It is seen that these propellers are capable of producing a large

amount of normal force, a characteristic of the wide, highly loaded blades. The shift in CL

at zero ot is due to either the improved flow over the pylons or due to a slight angularity of

the flow entering the propeller disk due to the pylon. A look at the Cm curve shows the

props produce a negative shift in Cm. At zero degrees alpha, This is mostly due to the

thrust vector being located above the CG. Also shown on the graph, by the dotted line, is

the basic aircraft Cm plus the change due offset thrust, T'c, and to the additional lift provided

by the propeller normal force, CNp. The propeller normal force was assumed to be the total

lift with power less the basic body lift and component of T'c. This curve closely matches

the shift in Cm and change in SM. The improved flow over the nacelles contributes to this

stable change in Cma.

Effect on Longitudinal Characteristics of Full Configuration. Figures 32 and 33

show the effect of the direction of propeller rotation on the full configuration longitudinal

characteristics. There is a slight change in Ct.a from 0.098 to 0.105 per degree with full

power. The direction of rotation dGes not affect the slope change. The lift at zero cz is

increased with power. This result is primarily due to the effect of the propellers blowing

over the horizontal stabilizer since there can be no appreciable component of thrust or

40
CL

0 5 10 15 20 25 30 35 40

o_(degrees)

Figure 30 - Effect of Power on Lift of BodyPylonsNacelles

0.4 "

0.2 "

0.0

-0.2

Cm

-0.4

-0.6

-0.8

-1.0

Figure 31

•----u---- No Power

-----o--- OBD 4500 RPM

x IBD 4500 RPM

.... •.-..-

-5 0 5 10 15 20 25 30 35 40

ec (degrees)

Effect of Power on Pitching Moment of BodyPylonsNacelles

41
2.25

2.00

1.75

1.50

1,25

C_

1.00

0.75

0.50

0.25

0.00

-5

-_ No Power

450O RPM OBD

450O RPM IBD

........ CL+T'c* sin(o_)

0 5 10 15 20 25 30 35 40

c_ (degrees)

Figure 32 - Effect of Power on Configuration Lift, &r=0 °

Crn -1

-2 -

-3

-5

i

No Power

-----o---- 4500 RPM OBD

& , 4500 RPM IBD

oo •

_meb o

O°OI_

eOO0

• I " I " I " 1 " I " I " I " '1

5 10 15 20 25 30 35 40

o_(degrees)

Figure 33 - Effect of Power on Configuration Pitching Moment, _f=O °

42
propeller normal force in the lift direction at zero or. A look at the CL curve at stall shows

that there is no distinct configuration stall with power. In fact, the wing is stalling, but the

propeller thrust and normal forces have components acting in the direction of lift. The

basic power-off lift plus the component of thrust acting in the direction of lift, T'c*sin(o0, is

also plotted on the graph. This is meant to show any post-stall lift increase which is not

due to the thrust vector. In fact, more than half of the additional lift after stall with power is

not due to the thrust vector, but rather due to the propeller normal force and increased

stabilizer lift. This is again a characteristic of the highly loaded propeller blades. The

difference between the IBD and OBD cases is not as significant for the full configuration.

A look at the pitching moment curve shows that power causes a large decrement in

Cm, due to the thrust acting above the CG producing a negative moment and also to the

increased lift on the horizontal tail which also produces a negative moment. With power,

the stable break at stall is more pronounced, due to the improved flow and increased lift on

the horizontal stabilizer and nacelles. Throughout most of the range of o_'s, the OBD case

shows a larger decrement in Cm. Also plotted is the basic Cm plus the decrement caused

by the thrust vector, CNp, and increased stabilizer lift. The reason for plotting this curve is

to show there is no significant change in Cm due to the wing or body.

Figure 34 shows the variation of drag with power. From the definition of T'c, there

is propeller thrust through most of the o_ range for both directions of rotation, but with the

propellers rotating OBD, there is thrust through the entire range of a's tested. This

increased thrust with the propellers rotating OBD parallels the result found with the body

alone, ttere, because the horizontal stabilizer directs the propeller inflow, there is no

difference in T'e between the two cases at zero angle of attack. The value of T'¢ is 0.34. The

thrust here is lower than the body alone cases because of unfavorable interference of the

wing wake with the propeller inflow.

Effect on the LateralDirectional Characteristics of the Full Configuration. Figures

35 and 36 show the effect of power on the directional and lateral stability of the full

43
CD 0.5

No Power

------e---- IBD 4500 RPM

--- OBD 45OO RPM

-05

-5 0 5 10 15 20 25 30 35 40

O_(degrees)

Figure 34 - Effect of Power on Configuration Drag, _f--O °

44
0.008

0.006 o....

0.004

0,,,

C.p

0.002

0.000

-0.002 .

-5

No Power

IBD 4500 RPM

• ! • ! I " I I I ' ! • ''|

0 5 10 15 20 25 30 35 40

ot (degrees)

Figure 35 - Effect of Power on Directional Stability, 6f=O °

0.000 - i

-0.002

--'-"a"--- No Power

o IBD 4500 RPM

-=- OBD 4500 RPM

Cip -0.004

-0.006

-0.008

-5 0 5 10 15 20 25 30 35 40

o_ (degrees)

Figure 36 - Effect of Power on Lateral Stability, _f=O °

45
configuration.Throughthe entire ot range, there is an increase in directional stability, Cnl3.

When the aircraft is in sideslip the propellers produce a side force, similar to the normal

force experienced with the propellers at or. This side force is behind the CG and therefore

causes a restoring moment which increases Cnl3. There is also an increase in lateral stability

(negative CII0 with power. This is a direct result of the thrust axis being located above the

CG. Again, when the propellers are at sideslip, a side force is produced. This side force

causes a restoring rolling moment which decreases Cll3.

Effect of Fillet Fairings on Propellers Rotating in Both Directions. Removeable

fillet fairings were built which served as a duct for the flow between the propeller disk and

the tail surfaces. The intention by ducting the flow was to confine the propeller slipstream

to create a thrust increment from the cases without fillets. Figure 37 shows a slight thrust

increment for the propellers rotating OBD with the fillets installed compared to the case

without the fillets. With the propellers rotating IBD, there is actually a reduction in thrust

for the case with fillets compared to the case without fillets. The fillets serve to enhance the

crossflow effect described earlier. With the propellers rotating OBD, the flow converges

into a smaller area yielding more thrust. With the propellers moving down inboard, the

flow diverges outward further reducing the thrust. The T'c for the IBD case is 0.32 and the

T'c for the OBD case is 0.37. The effect is similar when looking at the lateral and

directional characteristics.

Characteristics of Final Configuration

The final configuration was chosen to be the case for the propellers rotating OBD with the

fillets in place. This was chosen because of the greatly increased thrust at high alphas for

the OBD rotation and for the enhanced crossflow effects with the fillets in place.

Effect of Fillets on full Configuration without power. Figures 38, 39 and 40 show

the effect of the fillets on the longitudinal characteristics of the full configuration without

power. There is no significant change in lift or drag. There is a slight negative shift in Cm

through the alpha range.

46
1.25

1.00

0.75

-----o----- No Power

--.---o---- IBD No Fillets

,L tBD With Fillets

o OBD No Fillets

:. OBD With Fillets

CD

0.50

0.25

0.00

-0.25

-0.50

-5 0 5 10 15 20 25 30 35 40

ot (degrees)

Figure 37 - Effect of Fillets on Drag and Installed Thrust, 6r=O °

47
CI.

1.75

1.50

1.25

1.00

0.75

0.50

0.25

0.00

/

,- [_ w,,h°u, With FilletF,,,°, Fairings _a'r'nos 1

._...,____

-5 0 5 10 15 20 25 30 35 40

ot (degrees)

Figure 38 - Effect of Fillets on Full Configuration Lift, 6r=O °

I_ Without With FilletFillet Fairings, Fairings !

Cm -0.4

-5 0 5 10 15 20 25 30 35 40

ot (degrees)

Figure 39 - Effect of Fillets on Configuration Pitching Moment, 6t-=O °

48
1.2

1.0

0.8

Co 0.6

0.4

0.2

0.0

-5 0 5

[_ Without With FilletFillet Fairings Fairings

10 15 20 25 30 35 40

ot (degrees)

Figure 40 - Effect of Fillets on Full Configuration Drag, 6r=O °

49
Effect of Power on Longitudinal Characteristics. Figures 41, 42 and 43 show the

effect of power on the longitudinal characteristics of the final configuration. The basic

model with no power is seen to have a Ct_=0.098 per degree with a CLmax=1.52 at a stall

ot of 14 degrees. The propellers rotating at 3500 RPM corresponds to a T' =0.2. There is a

slight increase in CI_ to 0.103 per degree due to increased lift on the tail and propeller

normal force. The zero ot lift is increased from 0.385 to 0.493 due to increased lift over the

horizontal tail. The propellers rotating at 4500 RPM corresponds to a T'c=0.37. The value

of Ct_ is increased to 0.11 per degree. The zero ot lift increases to 0.568 and is larger than

the 3500 RPM case due of course to a larger blowing effect over the horizontal tail. For

each of the power settings, the configuration exhibited a large negative shift in Cm for the

entire range of c_'s. This shift was due to the thrust line being located above the CG,

increased lift on the horizontal stabilizer, and propeller normal force. This Cm change has a

corresponding trim alpha change of approximately 25 degrees from the unpowered to the

full power case. The large trim ot change between the power off and power on cases may

be severe enough to warrant zm all-moving stabilizer if a prototype flying aircraft were built;

however, a proper configuration design cot, ld minimize the trim change. There is an

increase in longitudinal stability with power. The basic configuration has a C,,,_=-0.0173

per degree corresponding to a SM of 17.6%. With "I_=0.2, Cmct decreases to -0.0263 per

degree and a SM of 25.5%. With _=0.37, Cm_t decreases further to -0.0305 per degree

and a SM of 27.7%. The large increase in SM with power could result in dynamic control

sluggishness in flight. The static trim problem will be discussed in more detail when

describing the elevator effectiveness below. The drag curve shows that the propellers

produce thrust through the entire range of ct's, and induce acceleration through

approximately 13 degrees for 3500 rpm and 18 degrees for 4500 rpm.

Effect of Power on Lateral/Directional Characteristics. Figures 44 and 45 show the

effect of power on the lateral and directional characteristics of the final configuration.

There is a positive increment in Cn for each of the power settings. The probable cause of

5O
CL

2.25

2.00

1.75

1.50

1.25

1.00

0.75

0.50

0.25

o.oo

-5

d/

No Power

3500 RPM

4500 RPM

0 5 10 15 20 25 30 35 40

O_(degrees)

Figure 41 - Effect of Power on Final Configuration Lift, _f=O °

0"00 ' __-_:i,_:}_ I_ NoPower

-o.25: -,% i---o-- 3500

-0.50 M

-0.75

-I .00

-125

Cm

-1.50

-1,75

-200

-2.25 -2.50 1

-2.75 • _ , i , w , , , , , ' , ' , • ,

-5 0 5 10 15 20 25 30 35 40

ot (degrees)

Figure 42 - Effect of Power on Final Configuration Pitching Moment, _r=O °

51
1.25

1.00

0,75

0.50

CD 0.25

0.00

-0.25

No Power

----o--- 3500 RPM

A 4500 RPM

-0.50

- 5 0 5 1 0 1 5 20 25 30 35 40

ot (degrees)

Figure 43 Effect of Power on Final Configuration Drag, 6f=O °

52
0.06

0.04

Cn

0.02

0.00

-0.02

-0.04

-0.06

-5

------a------ No Power

-----o--- 3500 RPM

_, " 4500 RPM

= • = w • l w w • [ '1

0 5 10 15 20 25 30 35 40

et (degrees)

Figure 44 - Effect of Power on Yawing Moment, 6f=0 °

0.008 F I _ No Power

t

I ------o------ 3500 RPM

PM

0.006 f

C,,I_ 0.004 f

0.002 f

-0.002/

0ooof L

t ' i • 1 I , , I , i ' I ' ,

- 5 0 5 1 0 1 5 20 25 30 35 40

oc (degrees)

Figure 45 - Effect of Power on Directional Stability, 6r:O °

53
this incrementcomes from one of the reasonsdescribedbelow. The propellerseach

consistedof a complexhub/pitchadjustmentassemblywhich allow thebladeangleto be

varied: however, the blades could only be locked down in certain positions. The

construction of the blades did not allow both the left and right assemblies to be set at

"exactly" the same blade angle. For this reason, there was a slight difference in the blade

angle between the two hub assemblies. In addition, the propellers were driven by separate

air turbines which were supplied pressure from different sources. In a situation such as

this, it is impossible to regulate the propellers so that each air motor was running identically

at the same speed. Thus, there was a difference in thrust from one side of the model to the

other which resulted in an increment in yawing moment. A similar effect occurs for higher

o_'s in rolling moment. As expected, power increases tile directional stability, C,,[_, due to

the propeller side force when the model was in sideslip. An increase was also seen in

lateral stability.

Effect of Flaps. As seen in figure 46, the flaps decrease CL_ from 0.098 per degree

to 0.086 per degree. There is a shift in CL due to flaps of 0.41 with 35 degrees flaps.

With 35 degrees of flaps, CLmax is increased to 1.64. There is only a slight change in the

shape of the CL curve at stall indicating the droop is still effective in tile landing

configuration. A look at the pitching moment curve in figure 47 shows the same effect as

figure 22 above, an increase in pitching moment through the entire range of cos with flaps

deflected. As described earlier, this increase is due to the increased down-loading on the

horizontal tail with a large Vrl. There is no large change in SM with flaps.

Effect of Flaps on Lateral/Directional Stability. Figure 48 shows the flaps slightly

increase the directional stability through the low and high ot's. This is because the flaps

slightly increase the vertical area behind the CG providing a larger restoring Cn. Figure 49

shows the lateral stability is not affected significantly.

Effect of rudder on Directional Characteristics. Figure 50 shows the rudder

authority with a full rudder deflection and the rudder effectiveness for the full rudder and

54
1.75

1.50

1.25

1.00

CL 0.75

0.50

0.25

0.00

-5

0

°I

0

-----o---- No 35 Deg FlapsFlaps I

I

I ' 1 " I " I " 1 I I "

5 10 15 20 25 30 35 40

et (degrees)

Figure 46 - Effect of Flaps on Final Configuration Lift

Cm -0.4

-1.0

-5 0

I_ No35 Deg FlapsFlaps

5 10 15 20 25 30 35 40

(z (degrees)

Figure 47 - Effect of Flaps on Final Configuration Pitching Moment

55
Cnl_

0.005 -

0.004

0.003

0.002

0.001

0.000

-0.001

-0.002

I I_ No Flaps ps

• I " I " I I " I " I ' I " I

-5 o 5 lo 15 2o 25 3o 35 4o

ot (degrees)

Figure 48 - Effect of Flaps on Directional Stability

Cip

0.000

-0.001

-0.002

-0.003

-0.004

-0.005

-0.006

-0.007

-0.008

-5

I

No Flaps I

o ' 35 Oeg Flaps 1

I ' I ' I " ! I ' I I ' |

5 10 15 20 25 30 35 40

ot (degrees)

Figure 49 - Effect of Flaps on Lateral Stability

56
lower rudderonly. The rudder maintains its full authority until just after stall and drops off

through the post-stall ot's due to the vertical fin being located in the stalled wake of the

wing and fuselage. The lower rudder provides an average of 61% of the effectiveness of

the full rudder which corresponds approximately to the ratio of the lower rudder area to the

full rudder area of 0.591. Figure 51 shows the effect of power on the rudder effectiveness.

A comparison is shown for the lower and full rudders. It is shown that power increases

the rudder effectiveness significantly. This is due to the propeller blowing over the vertical

tail which increases the circulation for a rudder deflection. Smoke flow visualization

verified that the flow in the prop-wash was blowing across the deflected rudder. It is also

seen that the lower rudder with full power is more effective than the full rudder without

power. From this result it is possible to speculate the lower rudder alone might be

sufficient for cruise flight.

Effect of Elevator on Longitudinal Characteristics. Figure 52 shows tile effect of

elevator on pitching moment for several deflections of the full elevator. In addition, it

shows the elevator effectiveness for the full elevator and for the inboard elevator alone. It

is shown that the full elevator without thrust can trim tile aircraft to approximately 22

degrees ot with a corresponding CL of 1.27. The elevator effectiveness varies through the

ot range but remains more or less constant for ot less than 10 degrees. Above I4 degrees,

the effectiveness drops off sharply due to the horizontal stabilizer beginning to stall. The

inboard elevator alone has slightly more than half of the full elevator effectiveness which

roughly corresponds to the ratio of the inboard elevator area to tile full elevator area.

Figure 53 shows the effect of power on the full and inboard elevator effectiveness. It is

shown that power increases the effectiveness noticeably for o_ below 10 degrees. This is

due to the improved flow over the horizontal stabilizer with power because of propeller

blowing. At 10 degrees c_, the effectiveness drops off sharply, approaching tile power-off

elevator effectiveness. In fact, the full elevator effectiveness is lower than the power-off

effectiveness from an _ of 12 to 26 degrees. This result is likely due to the combined

57
Cn

0.06

0.04

0.02

0.00

-0.02

-0.04

_ -30 Deg Rudder

_ Neutral Rudder

__ Ru__.dder

0".)...,.,.__ _,. _ . ,,, .

v - v_u...o..._ v

-0.06

-5 0 40

I I 1 " I I " I " I

5 10 15 20 25 30 35

o_ (degrees)

Figure 50 - Final Configuration Full Rudder Authority, _f=35 °

0.000 -

-0.001 "

C"_r.0.002 '

e,,,

-0.003

No Power

o 3500 RPM

4500 RPM

-0.004 ' ' , ' ' ' ' ' ' ' ' ' ' ' ' '

-5 0 5 10 15 20 25 30 35 40

_t (degrees)

Figure 51 Effect of Power on Rudder Effectiveness, &r:35 °

58
0.5

0.0

Cm

-0.5

-t .0

-I .5

-5

-30 Oeg Elev

-20 Deg Elev

* 0 Deg Elev

--

! ' " I I I I ! " I " !

0 5 10 15 20 25 30 35 40

ot (degrees)

Figure 52 - Effect of Elevator on Pitching Moment, 6r:35 °

0.00

-0.01

Cm6e-O.02

-0.03

-0.04

o.-,

or

------o---- Full Without Power

a, Inboard 4500 RPM

' - Full 4500 RPM

I ' 1 I I I " I " I

-5 0 5 10 15 20 25 30 35 40

ot (degrees)

Figure 53 - Effect of Power on Elevator Effectiveness, &¢:35 °

59
influenceof theincreasingwingdownwashandpropellerpropwashmoving awayfrom the

stabilizer and interferingwith the local flow over theelevator,eliminating the blowing

effect and distorting the local flow over the elevator. At ct's above 26 degrees, the

propwash is close enough to the stabilizer to affect the inboard elevator while the stabilizer

is below the wing wake such that the effectiveness is again increased from the power-off

case. Another interesting result is that the increase of effectiveness of the inboard elevator

is substantially more than the increase of effectiveness of the full elevator. In fact, most of

the full elevator effectiveness with power is provided by the inboard elevator. This is

because the propeller is blowing across the inboard portion of the horizontal stabilizer, so

the inboard elevator is seeing a large effect of the blowing while the outboard elevator sees

little. For the full power case, the inboard elevator is more effective than the full elevator

with no power through most of the ot range.

Effect of Power on Trim Capability. Figure 54 shows a series of full elevator

deflections for the full power climb, missed approach condition of full power with a _f of

35 degrees. The neutral elevator case shows a large negative trim ot due, of course, to

power effects on the pitching moment. It is shown that the configuration can trim at

approximately zero degrees a with -20 degrees of elevator and at approximately 8 degrees

ot with -30 degrees of elevator. A reasonable upper limit on allowable elevator deflection is

30 degrees. With a larger deflection, there would be no appreciable gain in elevator

authority. With a minimum of 20 degrees deflection required purely to trim, there remains

only 10 degrees of deflection for pitch control. However, it was shown that power

increases the elevator effectiveness significantly. Because of this, the elevator is probably

effective to a deflection higher than 30 degrees. With a reasonable climb angle of 3

degrees, this missed approach condition was accepted as being reasonable considering the

potential to trim up to 10 degrees oc for the flaps down full power climb case. It is expected

that the flaps up cruise condition would not have a trim problem because the cruise thrust

setting would be lower than the full power climb, resulting in a lower pitching moment due

60
to power. Thus, the powereffectson elevator effectivenessare significant enough to

overcometheunfavorable pitching moment change with power.

Effect of Ailerons on Lateral Characteristics. Figure 55 shows the variation in

aileron effectiveness with ot in addition to the aileron authority for a full deflection of + and

- 20 degrees. The effectiveness varies approximately linearly with angle-of-attack through

the entire range of ot's. This reduction in effectiveness is caused by trailing edge separation

which is shown the the wing stall pattern (figure 8). ttad the wing droop proved to be

more effective, the aileron effectiveness would have been expected to remain constant to

higher ct's.

Figure 56 shows the yawing moment with neutral ailerons and with a 20 degree

positive aileron deflection. A slight amount of positive yawing moment with tile 20 degree

deflection indicates adverse yaw through most of the angle-of-attack range.

Figure 57 shows the effect of power on tile aileron effectiveness. There appears

to be an increase in effectiveness until around 25 degrees ot due to some effects induced by

the airflow into the propeller disks.

Engine Out Characteristics. A major concern with multiple engine transport aircraft

is tile ability to trim and continue flying with one or more engines out. Results presented

above show the capability to trim with full power and with no power. Figure 58 shows the

yawing moment curve. The symmetric full power case produces negligible yawing

moment until near stall where the wing wake begins to interfere with the propeller disks.

With the left engine windmilling and the right engine at full power, the aircraft experiences

a sharp negative yawing moment. However, with full corrective rudder deflection in

addition to the engine out condition, the yawing moment is shifted positive. The rudder

provides enough yawing moment to nearly double the engine-out yawing moment, the

aircraft can trim with one engine out until approximately 22 degrees. When one propeller is

inoperative, a rolling moment is produced due to the increased lift on the side of the

operative propeller. This is shown in figure 59. The curve with full corrective ailerons

61
0

Cm - 1

-30 Deg Elev

-20 Deg Elev

0 Deg Elev

20 Deg Elev

30 Deg Elev

-5 0 5 10 15 20 25 30 35 40

o_ (degrees)

Figure 54 Effect of Power on Trim Capability, t_f=35 °

62
0.10

0.08

0.06

0.04

0.02

Ci o.oo a'J

-0.02

-0.04 -0.06 -

-0.08

-0.10

-5

_ w m

+10 Deg Ailerons

A +20 Deg Ailerons

• I I " I " I ! I " I " |

5 10 15 20 25 30 35 40

ct (degrees)

Figure 55 - Effect of Ailerons on Rolling Moment, 6f=35 °

0.06

0.05

0.04

0.03

0.02

0.01

C. 0.00

-0.01

-0.02

-0.03

-0.04

-0.05

-0.06

l_i,_t..r i. t w _ v

, ! • !

-5 0 5 10

I _.____ 020DegDegAileron Aileron I

I " I I I I " l

15 20 25 30 35 40

et (degrees)

Figure 56 - Effect of Ailerons on Yawing Moment, 6r=35 °

63
C18a

0.000

-0.001

-0.002

-0.003

-0.004

-5

' or

F 04500 RPM

" _ ' I ' I ' I ' I ' I ' I " I '

0 5 10 15 20 25 30 35

(degrees)

I

|

4O

Figure 57 - Effect of Power on Aileron Effectiveness, 6r=35 °

64
shows the ailerons have enough authority to overcome the engine-out "rolling moment

through the entire a range.

Comparison To Theoretical Resulls

In order to provide a good overall set of data for the aircraft, the configuration was

analyzed with a first-order potential flow "doublet" panel method to p,'u-allel the wind tunnel

data. Such methods have proven to be excellent sources of basic aerodynamic

characteristics of full configurations. The panelling method, QUADPAN (ref 18), was

used since it provides a full surface analysis with relatively simple geometry input. In

addition, QUADPAN includes the capability to approximate propeller effects on the

configuration using a simple slipstream model. An orthographic view of the completed

panel model is shown in figure 60 The model contains 2070 quadrilateral panel elements.

QUADPAN calculates the potential flow over the model by the use of doublet and source

elements on the surface of the model. In order to assure proper alignment and positioning

of the elements, which would assure the satisfaction of all boundary conditions, some

minor adjustments were made from the true geometry to the panel model geometry;

however, no dimension was changed more than approximately 1/4 inch which is much

smaller than any major dimension on the model. Thus, the model represented the full wind

tunnel model.

Comparing the Restdts. Figures 61 and 62 show a comparison between the wind

tunnel results and the QUADPAN results for the longitudinal characteristics of the cruise

configuration without power (Sf=0). There is a good comparison in the lift curve. Since

QUADPAN is a linear potential code, no stall is shown for the theoretical results. The

slope and zero ot lift are extremely close to the wind tunnel results. The pitching moment

comparison is quite good as well. The SM from QUADPAN is approximately 16.1%

compared to 14.9% from the wind tunnel. The trim ot is slightly low from QUADPAN,

but the difference shown in the wind tunnel data could be due to viscous effects, model

support effects, etc.

65
Crt

Left Engine Out, Rudder=-30 degrees

oo,; /

0.02 -

0.00

-0.02 -

o_

-0.04

-0.06 • Lell

I

-0.08 • _ • , - i I ' , • r , r • 7

-5 0 5 10 15_ 20 25 30 35 40

ot (degrees)

Figure 58 - Engine-out Yawing Moment, _f=35 o

0.08 -

0.06 - a"

C_ ooo '

-0.02

Left

-0.04 I

-0.06 t

-0.08 I

Left Engine Out, Full Corrective Aileron

/

ct grees)

Figure 59 - Engine-out Rolling Moment, Sf=35 o

66
NASA-LANGLEY ATP-OTT FULL CONFIGURATION

Z

Figure 60 - Orthographic View of QUADPAN Panel Model

67
q.

1.75 -

1.50

1.25

1.00

0.75

0.50

0.25

0.00

-5 0 5 10

I

---.--131---- Wind Tunnel I

I

.......... QUADPAN

• I " I I ' I I ' i

15 20 25 30 35 40

et (degrees)

Figure 61 - Comparison to Potential Lift, _f=O °

Cm

1.0"

0.5-

0.0 "-=-

-0.5

-1.0

-1.5

-5 0

Figure 62

_Q_e

QUADPAN

I I ' I " I " I I I " I

5 10 15 20 25 30 35 40

(degrees)

Comparison to Potential Pitching Moment, _¢=0 °

68
Figures63and64showacomparisonfor the lateral and directional characteristics

of the same configuration. The directional stability is slightly low from QUADPAN. A

variety of element arrangements were checked to see if this underestimate in Cn9 was due to

an underestimate in side force on the vertical tail with sideslip; however, the value of the

side force on the vertical tail, and consequently Cn_, never changed much. From this it is

possible to conclude that the yawing moment due to the body, pylons and nacelles was

underestimated. If an average Cif_ is taken in the linear range from the lift curve, the

QUADPAN value of C1_ corresponds nearly exactly.

Figures 65 and 66 show QUADPAN's estimate to lift and pitching moment due to

the propeller slipstream effects. These results ,are for a total _ (left + right) of 0.36. For

the plot, the vector component of thrust was added to both the lift and pitching moment.

For CL, there was negligible change due to power. This did not compare well to the wind

tunnel results which showed a large change in lift due to blowing over the horizontal tail

which was not well represented by QUADPAN. The pitching moment shift with power

was low as well. From the plot, the theoretical shift in Cm was approximately half of the

wind tunnel shift. In addition, the majority of this shift was due to the thrust line being

above the CG Because QUADPAN did not predict the increased lift on the horizontal tail.

it also did not show the pitching moment change due to the additional lift on the tail.

Figures 67 and 68 shows similar under-predictions of the effect of power on Cnl_ and CI[_.

I

Thus, QUADPAN does not well predict the aerodynamic effects due to propeller slipstream

blowing over the tail surfaces for this configuration.

69
0.006

0.004

0.002

C.13

0.000

-0.002

-0.004

.I

_ Wind Tunnel I

........ QUADPAN

• I " I " 1 i " I I I " II

-5 0 5 10 15 20 25 30 35 40

ot (degrees)

Figure 63 - Comparison to Potential Directional Stability, 6t-O °

0.000 -

-0.001

wm m=mlm = i_ma=

-0.002 "

-0.003

CII3 -0.004

-0.005

-0.006

-0.007

-0.008 ! ' I

-5 0 5 10

-----o---- ........ Wind QUADPANTunnel ]

I ' I " I I • ! ' _1

15 20 25 30 35 40

ot (degrees)

Figure 64 - Comparison to Potential Lateral Stability, 6f-'O °

7O
q_

2.25

2.00

1.75

1.50

1.25

1.00

0.75

0.50

0.25

0.00

-5 0 5

Figure 65-

Wind Tunnel No Power

QUADPAN No Power

Wind Tunnel 4500 RPM

QUADPAN 4500 RPM

10 15 20 25 30 35 40

cc (degrees)

Predicted l'ower Effects on Lift, _f=O °

0

Cm -1

Wind Tunnel No Power

QUADPAN No Power

Wind Tunnel 4500 RPM

QUADPAN 4500 RPM

PI I I Ill I I fill II ii i lliIii ii i I i i I

Figure

0 5 10

66- Predicted

15 20 25 30 35 40

o_(degrees)

Power Effects on Pitching Monlent, _f = 0 °

71
C.13

0.008-

l

0.004 i

0.002

0.000

-0.002 •

-5 0

_ No Power

'm'-e--- 4500 RPM

I ........ QUADPAN NO Power

- I .......... , ?UADPAN4500RPM

............................ :.-;-.;.;

! ' ! I ' I I I I " I

5 10 15 20 25 30 35 40

(:z(degrees)

Figure 67 - Predicted Power Effects on Directional Stability, &f=O °

0.000

-0.002

Cll3 -0.004

-0.006

-0.008

----.B---- Wind Tunnel No Power

I ........ QUADPAN No Power

_ Wind Tunnel 4500 RPM

I .......... OUADPAN 4500 RPM

..................................... ._..

-5

' I " | " I " I • I " I " I ' I " I

0 5 10 15 20 25 30 35 40

c_ (de grees)

Figure 68 - Predicted Power Effects on Lateral Stability, 6f=O °

72
(_oncluding Remarks

As part of a cooperative effort between NASA Langley Research Center and North

Carolina State University, an exploratory wind tunnel investigation was performed in the

30x60 Foot Wind Tunnel to determine the static stability and control characteristics into

deep stall of an advanced turboprop configuration with the propellers placed above the

horizontal tail. A list of the major discoveries is presented below.

1) The model had fair longitudinal stall characteristics. A leading edge droop

modification on the wing somewhat reduced a tendency to autorotate at stall.

2) A large horizontal and vertical tail provided both longitudinal and directional

stability for the unpowered configuration.

3) Power improved the static longitudinal and directional stability through the entire

angle of attack range and improved the lateral stability in the post-stall angle of attack range.

In addition, power greatly increased the elevator and rudder control effectiveness through

the entire angle of attack range tested.

4) The case for the propellers rotating with the outboard blades moving down was

shown to have more installed thrust than the case for the propellers rotating with the

inboard blades moving down. This result was due to a "crossflow" behind the propeller

disk which varied depending on the direction of propeller rotation.

5) A set of fillet fairings which ducted the flow behind the propeller disks was

shown to enhance the crossflow effect and enhance the difference in installed thrust

between the two directions of propeller rotation.

6) Power was shown to drastically change the longitudinal trim angle of attack,

primarily due to the thrust line being located above the center of gravity. It is believed that

a proper configuration design could minimize this trim change.

7) The model was shown to have sufficient rudder and aileron authority to trim with

one engine inoperative and the other engine at fidl power.

73
8) A potential flow doublet panelling code, QUADPAN, accurately predicted the lift

and pitching moment of the full configuration. In addition, directional and lateral stability

were predicted with a fair degree of accuracy. The code was insufficient in predicting

propeller induced effects

74
References

1. ttager, Roy D., and Vrabel, Deborah. Advanced Turboprop Project. NASA SP-495.

1988.

2. Ziemianski, Joseph A., and Whitlow, John B., Jr. NASAIndustry Advanced

Turboprop Technology Program. NASA TM- 1(X)929. August 1988.

3. Aljabri, A.S. IVind Tunnel Tests On A One-Foot Diameter SR-7L Propfan Model.

AIAA Paper 87-1892. June 1987.

4. Dittmar, James H. Feasibility of Wing Shielding of the Airplane Interior From The

Shock Noise Generated by Supersonic Tip Speed Propellers. NASA TM-79042.

December 1978.

5. Dittmar, James H. An Experbnental Investigation of Reducing Advanced Turboprop

Cabin Noise by Wing Shielding. NASA TM-87112. July 1986.

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7. Coe. Paul L., Jr., Turner, Steven G., and Owens, D. Bruce. Low-Speed Wind-Tunnel

Investigation of the Flight Dynamic Characteristics of an Advanced Turboprop

Business/Commuter Aircraft Configuration. NASA TP-2982. April 1990.

8. Dunham, Dana Morris., et al. Low-Speed Aerodynamic Characteristics of a Twin

Engine General Aviation Configuration With Aft-Flt_elage-Mounted Pusher

Propellers. NASA TP-2763. October 1987.

9. Applin, Zachary T., and Coe, Paul L., Jr. Low-Speed Stability and Control

Characteristics of a Transport Model With Aft-Fuselage-Mounted Advanced

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Large-Scale Supersonic Transport Model. NASA TN D-8296. November 1976.

11. Hale, Francis J. Introduction to Aircraft Performance, Selection, attd Design. John

Wiley and Sons, Inc. Copyright 1984.

12. Johnson, J.L., Jr., and White, E.R. Exploratory Low-Speed Wind-Tunnel

Investigation of Advanced Commuter Cot?/igurations lnchuting an Over-77ze-Wing

Propeller Design. AIAA Paper 83-2531. October 1983.

13. Johnson, Joseph L., Jr., Yip, Long P., and Jordon, Frank L., Jr. Preliminary

Aerodynamic Design Considerations For Advanced Laminar Flow Aircraft

Configurations. Presented at SAE/AIAA/NASA/FAA Laminar Flow Aircraft

Certification Workshop. April 1985.

14. Williams, L.J., Johnson, J.L., Jr., and Yip, L.P. Some Aerodynamic Considerations

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75
15.Baals,DonaldD., andCorliss,William R. Wind Tunnels of NASA. NASA SP-440

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John Wiley and Sons, Inc. Copyright 1984.

18. Youngren, Harold It., Bouchard, Eugene E., and Coopersmith, Robert M.

Quadrilateral Element Panel Metlugd (QUADPAN) User's Manual Version 3.2.

Lockheed-California Company LR-30563. Copyright 1984.

19. Benin, John J., and Smith, Michael L. Aerodynamics For Engineers. Prentice-Hall,

Inc. 1979.

20. Margason, Richard J., et al. Subsonic Panel Methods - A Comparison of Several

Production Codes. AIAA Paper 85-0280. 1985.

21. Staff of Langley Research Center: Exploratory Sttuty of the Effects of Wing-Leading

Edge Modifications on the Stall/Spin Behavior of a Light General Aviation

Airplane. NASA TP-1589. December 1979.

22. Stinton, Darrol. The Design of the Aeroplane. BSP Professional Books, Great Britain

23. Samuelsson, Ingemar. Low Speed Wind Tunnel Investigation of Propeller Slipstream

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and Moments and Pressure Distributions on NacelleWing at Different Angles of

Attack and Yaw and at D(fferent Thrzt_t Coefficients. The Aeronautical Rese,"uch

Institute of Sweden FFA TN 1987-22. 1987.

24. Johnston, R.T., Witkowski, D., and Sullivan, J.P. Propeller Wakes and Their

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76