AFFDL-TR.71-26.

VOL I

STOL HIGH-LIFT DESIGN STUDY

Volume I. State-of-the-Art Review of

STOL Aerodynamic Technology
Fred May
Colin A. Widdison

The Boeing Company

TECHNICAL REPORT AFFDL-TR-71-26.VOL I

April 1971

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STOL HIGH-LIFT DESIGN STUDY
Volume I.- State-of-the-Art Review of STOL Aerodynamic TechnologyI
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Final Report January-December 1970

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Colin A. Widdison
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Air Force Flight Dynamics Laboratory
Volume I of a 2-volume report Research and Technology Division
Air Force SysterkLS Command
13 ABSTRACT

The state of the art of STOL aerodynamic technology for selec-

ted lift/propulsion concepts has:.been surveyed to identify
the-available test data and prediction methods in the literature.
The report consists of two volumes.
In Volume I important aroeas of technology and information necessary
for the evaluation of STOL aircraft aerodynamics are listed; the
5..aerodynamic test data and prerliction methodology r'elevant to
the deflected slipstream and externally blown flap concepts are
assessed, with emphasis on the latter; an empirical method for
the prediction of the longitudinal aerodynamic characteristics
of externally blown flap configurations is presented; and high-
lift technology for five lift/propulsion concepts is asst*.ssed in
j application to a medium-sized STOL transport.
Volume II consists of a bibliography that resulted from a litera-
ture search for aerodynamic information related to seven lift/
I propulsion concepts suitable for STOL aircraft.
contains references to approximately 900 reports classified by
The bibliography
concept and by technological area.

BD I JAN044 1473 UNCLASSIFIED

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ROLE WT ROLE Wr ROLE Wr

STOL aerodynamic technology

Test data
Prediction methods
Deflected slipstream
Externally blown flaps
5 lift/propulsion concepts
STOL medium t.ansport

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F

I"I

STOL HIGH-LIFT DESIGN STUDY

Volume I. State-of-the-Art Review of

[STOL Aerodynamic Technology

Fred May
FColin A. Widdison

The Boeing Company

This document has been approved for public release

and sale; its distr;'jution is unlimited.

I
I
I
I
SI
 1
1
1
1

FOREWORD j
This report was prepared for the Air Force Fliht Dynamics Laboratory by The
Boeing Company. The study was conducted under USAF Contract F33615-70-C-
1277, "STOL High-Lift Design Study", with Captain Garland S. Oates as Project
Engineer for the Air Force.

The study was performed from January through December. 1970 by Mr. Colin
Widdison of the Vertol Division and Mr. Fred May of the Military Airplane Sys-
tems Division.

This technical report has been reviewed and is approved.

I
i

Lt. Col. USAF

Chief, V/STOL Technology Division I

ii .1J

A
[ABSTRACT
The state-of-the-art of STOL aerodynamic technology for selected lift/propulsion
concepts has been surveyed to identify the available test data and prediction methods
in the literature. The report consists of two volumes.

r In Volume I important technology areas and information necessary for the evalua-
tion of STOL aircraft aerodynamics are listed; the aerodynamic test data and pre-
diction methodology relevant to the deflected slipstream and externally blown flap
concepts are assessed with emphasis on the latter; an empirical method for the pre-
diction of the longitudinal aerodynamic characteristics of externally blown flap con-
figurations is presented; and high-lift technology for five lift/propulsion concepts is
assessed in application to a medium-sized STOL transport.

Volume II consists of a Bibliography that resulted from a literature search for

aerodynamic information related to seven lift/propulsion concepts suitable for
STOL aircraft. The Bibliography contains references to about 900 reports classi-

fied by concept and by technology arpa.

[
[
[

[
[
I
I
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II
 TABLE OF CONTENTS

F Page

FOREW OA D ...................................... 1i
IABSTRACT...........................................Uii

LIST OF ILLUSTRATIONS ............................ viii

LIST OF TABLES .......................... ...... xiii
SUM MARY ....................................... 1
1. INTRODUCTION . ........................... ........ 2
2. FACTORS IMPORTANT TO THE DESIGN AND OPERATION OF
STOL AIRCRAFT .................................. 3
2.1 INTRODUCTION ............................... 3
2.2 REQUIREMENTS FOR OBTAINING STOL PERFORMANCE .. 3
2.3 CRITERIA AND MARGINS ........................ 8
2..1.1 Performance Criteria ..... ...................... 8
2.3.2 Flying Qualities Criteria .... .................... 19
2.4 GROUND EFFECT ............................. 23
2.5 MODEL TESTING ............................. 27
2.5.1 Introduction ............................. 27
k.5.2 Simulation of Propulsion Systems ................ 31
2.5.3 Test Conditions .......................... 31
2.6 IMPORTANT TECHNOLOGY AREAS AND INFORMATION
REQUIREMENTS .............................. 39
2.6.1 Basic Acro/Propulsion Technologies ............... 39
2.6.2 Interacting Technologies .... .................... 40
2.6.3 Systems Requirements ..................... 42
3. LITERAL 'RE SEARCH
3.1 SCO PE .. ................................... 44
3.2 SOURCES ................................... 44
3.3 CLASSIFICATION OF BIBLIOGRAPHY .................... 45
4. EXTERNALLY BLOWN FLAPS ............................ 46
4.1 AN ENGINEERING METHOD FOR ESTIMATING THE TAIL-
OFF LONGITUDINAL AERODYNAMIC CHARACTERISTICS
OF AN AIRCRAFT WITH EXTERNALLY BLOWN FLAPS .... 46
4.1.1 Introduction ............................ 46
4.1.2 Required Information ........... ............ 51
4.1 .3 Lift .. .. .. . .. .. . ... . .. . ... . ... . . .. .. .. 59
4.1.4 Force Polars ............................ 64
4.1.5 Pitching Moments ................ ........ 69
4.1.6 Asymmetric Thrust ........................ 71
4.1.7 Application of the Theory. 77
4.1.8 Test-Theory Correlation ..................... 81
4.1.9 Conclusions ............................. 81
4.2 DEFNITION OF DEFICIENCIES AN) APS IN
KNOWI.EI;E. ..................................... 87

Iv
 Page

4.2.1 Introduction .................................. 87

4.2.2 Test Data ................................. 87
4.2.3 Analytical Methods ............................ 90
4.3 RECOMMENDED PRO IRAMS ........ ............ 94
4.3.1 Wind Tunnel Test Programs .... ............... 94
4.3.2 Analytical Development ....................... 95
4.4 CONCLUSIONS .............................. 95
5. THE DEFLECTED SLIPSTREAM CONCEPT .................. 96
5.1 INTRODUCTION ............................. 96
5.2 REVIEW OF PREDICTION METHODS .................. 104
5.2.1 General Considerations ....................... 104
5.2.2 Propeller Methods .......................... 105
5.2.3 Wing-Flap System Methods .... . ............ 108
5.2.4 Effects of Nacelles and Fuselage ... ............. 110
5.2.5 Combined Propeller-Wing-Flap System ........... 110
5.3 A DISCUSSION OF SELECTED METHODS FOR THE
DEFLECTED SLIPSTREAM CONCEPT ................ 114
5.3.1 General .................................. 114
5.3.2 Discussion of Specific Methods .................. 115
5.3.3 Summary Comments ........................ 126
5.4 CORRELATION OF TEST DATA WITH PREDICTION 127
5.4.1 Prediction Method of Reference 28 .............. 128
5.4.2 Flight Test Data and Calculations ............... 128
5.4.3 Wind Tunnel Test Data ....................... 133
5.4.4 Comments on the Predictions .................. 138
5.5 DEFINITION OF DEFICIENCIES AND GAPS IN
KNOWLEDGE ............................... 138
5.5.1 General .................................. 138
5.5.2 Test Data ................................ 138
5.5.3 Analytical Methods .......................... 141
5.6 RECOMMENDED PROGRAMS ....................... 142
5.6.1 Wind Tunnel Test Programs ................... 142
5.6.2 Development of Prediction Methods .............. 142
5.7 CONCLUSIONS ........ ........................ 144
6. A COMPARISON OF LIFT/PROPULSION CONCEPTS APPLIED
TO A MEDIUM STOL TRANSPORT ........................ 145
6.1 INTRODUCTION ............................. 145
6.2 DATA CREDIBILITY AND LIMITATIONS ................ 145
6.2.1 Externally Blown Flaps ........... .......... 145
6.2.2 Internally Blown Flaps ..... .................. 148
6.2.3 Augnentor Wing ... ............... 154
6.2.4 Direct Jet Lift ............................. 161
6.2.5 Mechanical High Lift Devices.166
6.3 CRITERIA FOR A COMPARATIVE EVALUATION ....... 170

Vi
 6.3.1 Internally Blown Flaps....................... 176
6.3.2 Augnientor Wing........................... 177
6.3.3 Direct Lift Engine Concept .................... 177

.4
......... 7
6.IO CUIN ......
6.3.4 Mechanical High Lift Devices ..................
CONCLUSIONS....................................184
178

REFERENCES .................................... 185

Vi
 LIST OF ILLUSTRATIONS

Figure Page

1 Variation of Takeoff Distance with Takeoff Speed and

Thrust/Weight Ratio .............................. 4 1
2 Variation of Landing Distance and Field Length with
Approach Speed .................................. 5

3 Possible Improvements in STOL Landing ............... 6

4 Lift Coefficient for level Flight .................. 7

5 Minimum Safe Speed .............................. 9

6 Gust Sensitivity of Rigid-Body Motion ................ 10

7 Vertical Gust Allev'-,tion .......................... 11

8 STOL Margins .................................. 13

9 Sensitivity of Ground Rules ....................... 15

10 Landing Parametric .......................... 16

11 Landing Parametric ............................. 17

12 Sensitivity of Ground Rules to Transition Time Delay .... 18

13 Effect of Criteria on Minimum Takeoff Speed for a

Typical STOL Airplane ............................ 21

14 Control Power Comparison ........................ 25

15 Theoretic.il Ground Effect .......................... 26 L

16 Effect r' Configuration on Ground Effect .............. 28

17 Effect of Bank Angle on Yawing Moment ............ 29

18 Moving Belt Testing Requirement for High-Lift W~. gs ... 30

19 Schematic Representation of the Condition for Undisturbed

Flow at the Wind Tunnel Walls ................... 34

viii
I Figure Page

20 Schematic Representation of the Condition for the

5 Disturbance of the Tunnel Boundary Layer ............ 35

21 Schematic Representation of the Condition for

5 Incipient Stagnation 36

22 Schematic Representation of the Condition for Vortex

Formation ................................ 37

23 Pressure Distribution Showing the Effects of the

Disturbance of the Tunnel Boundary Layer ........... 38

24 Pressure Distributions Indicating the Presence of

Incipient Stagnation .............................. 41

25 Water Tunnel Flow Visualization ........ 7........47

26 Comparison of Aerodynamic Characteristics with Spread

and Concentrated Blowing, a - 00 .................... 53

27 Nomenclature, Effective Jet Deflection Angle and

Static Turning Efficiency .......................... 54

28 Nomenclature, Effective Jet Deflection Angle ......... 55

29 Effect of Thrust Level on Static Turning Efficiency ..... 56

30 Flap Turning Efficiency Correlation Double-Slotted

Flaps.................................... 57

31 Flap Turning Efficiency Correlation Triple-Slotted

Flaps ......................................... 58

32 Jet Flap Effect on Lift Curve Slope ................... 61

33 Two-Dimensional Jet Flap Effectiveness ............... 63

34 Jet Flap Finite Aspect Ratio Factor .................. 65

35 Nomenclature, Developed Flap and Wing Chords and

Flapped Wing Area ............................... 66

3 36 Correlation of Maximum Lift Due to Power ............. 67

3 ix

I
Figure Page

37 Force Polar Build-up ............................ 70

38 Center of Pressure of Powered Lift .................. 72

39 Finite Aspect Ratio Effect on Chordwise Location of

Center of Pressure of Flap Lift ..................... 73

40 Effect of Asymmetric Thrust ..................... 75

41 Roll Arm of Failed Engine Lift ..................... 76

42 Theoretical and Test Data on Coefficient of Lift Versus

Angle of Attack, Drag, and Pitching Moment at the
I
Quarter-Chord ................................ 79

43 Effective Jet Angle Correlation ..................... 82

44 Lift Curve Slope Correlation ....................... 83

45 Correlation of Lift Increment due to Power at Zero

Angle of Attack ................................. 84

46 Power Induced Lift Center of Pressure Correlation .... 85

47 Theoretical and Test Data on Coefficient of Lift Versus

Angle of Attack, Drag, and Pitching Moment at the
Quarter-Chord ............................. 86

48 Ground Effect of Jet-Flapped Wing ................... 89

49 Comparison of Lifting Line, Slender Body and Lifting

Surface Methods of Reference 22 for a Wing with Ii
Sinusoidally Varying Angle of Attack in an Infinite
Plane-sided Jet ................................ 117

50 Comparison of Lifting Line and Slender Wing

Predictions of Reference 22 with Test Data from
Reference 34 for a Wing Extending Across a V2
Circular Jet ................................... 117

51 Comparison of Predictions Obtained Using the Method

of Reference 28 with Data from Reference 40 .......... 121
V

x
I Figure Page
52 Comparison of Prediction by Method of Reference 29
with Test Data from the Same Source ................ 122

53 Comparison of Predictions Using Method of Reference

20 with Test Data from Reference 36 ............... 123

1 54 Prediction of Onset of Stall ....................... 124

55 Estimated Downwash at the Tail of Breguet 941 in

Takeoff Configuration ............................ 132

56 Comparison Between Lift Coefficient and Test Values

from Reference 47 .............................. 134

57 Comparison Between Predicted Drag Coefficient and

Test Values from Reference 47 ................... 135

58 Comparison Between Predicted Force Polar and Test

Values from Reference 47 ......................... 136

59 Comparison Between Predicted Lift and Drag

Coefficients with data from Reference 48 ............ 137

60 Externally Blown Flaps Longitudinal Characteristics .... 146

61 Components of Lifting Force ....................... 147

62 Lift due to Thrust ................................ 149

63 Engine Out Rolling Moment ........................ 150

64 Internal BLC .................................. 151

65 BLC Flap Effectiveness ....................... .. 152

66 Aerodynamic Characteristics Full Span, Double Slotted

J Flap ........................................ 153

67 Augmentor Wing ............................... 155

68 Augmentor Wing Longitudinal Characteristics ........ 156

69 Augmentor Wirg Longitudinal Control - Conventional

Elevator Typical 40, 000 LB. Airplane .............. 158

I xi
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Figure Page

70 Augmentor Wing - Lateral Control (Unpui 'ished

NASA Test Data) ................................ 159

71 Augmentor Wing Lateral Control Typical 40, 000 LB.

Airplane .................................. 160

72 Augmentor Wing Ground Effect ...................... 162

73 NASA Lewis Noise Test Models .................... 163

74 Lift Engine Pitching Moments ...................... 165

75 Noise Levels - Bypass Ratio ....................... 167

76 Maximum Lift Capability of Mechanical Highlift Devices.. 168

77 Triple Slotted Flaps with 20 Degrees Slat Deflection ... 169

78 Geometry of Double and Triple Slotted Flap Model ...... 171

79 Double Slotted Flaps with 20 Degrees Slat Deflection 172

80 Mechanical High Lift Devices Zero Thrust Longitudinal

Characteristics of 4 Propeller Deflected Slipstream
STOL aircraft with Leading Edge Slat and Triple
Slotted Flap ......................... ...... 173

81 Lift Comparisons................................179

82 Noise Comparison .............................. 181

x.

xii
I
LIST OF TABLES

Table Page

I Summary of STOL Flight Experience ..................... 14

II Sampling of Flying Qualities Specifications .............. 20
III Comparison of Flying Qualities Parameters ............. 22
IV Comparison of Low Speed Flying Qualities Criteria ........ 24
V Characteristics of V/STOL Low Speed Wind Tunnels ........ 32
VI Relationship of Cj and Do for Lift Curve Slope Calculations... 62
VII Summary of Gaps and Deficiencies in Knowledge of Externally
Blown Flaps ................................... 91
VIII Deflected Slipstream Model Test Data ................... 97
IX Summary of Flight Test Data .................. ... 102
X Prediction Methods for the Deflected Slipstream Concept ..... 114
XI Availability of Methods and Data. . ....... ......... 183

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I xiii

I
 SUMMARY

The state of the art of STOL aerodynamics has been evaluated by assessing the
Jvalidity and applicability of available analytical and empirical prediction methods.

The important technological areas in STOL aircraft aerodynamic design are out-
lined and a list is presented of the information needed to permit STOL aero -
dynamic design, analysis, and evaluation.

IAn extensive literature search was made of aerodynamic prediction methods and
data applicable to seven distinct STOL concepts:

IExternally Blow Flaps

Deflected Slipstream
Jet Flap
Mechanical High Lift Devices (including boundary layer control)
Fan-in-Wing
Tilt Wing
Direct Jet Lift

The results of the literature search are summarized in a comprehensive bibliogra-

phy of nearly 900 references. The bibliography comprises Volume II of this report.

Primary emphasis was placed on researching prediction methods applicable to

externally blown flap and deflected sl!pstream configurations. The literature
search revealed a considerable amount of potentially useful test data and a
number of prediction methods related to the deflected slipstream configuration.
In contrast, it was discovered that relatively little test data exists for the ex-
ternally blown flap configuration and that there was a complete lack of analytical
methods applicable to that concept. As a result, the program activity was re-
directed at mid-*erm to Include the development of a method for the prediction
of the aerodynamic characteristics of externally blown flap configurations.

. The methods available for the deflected slipstream concept were reviewed and
analyzed and one method was evaluated against test data.

i The gaps and deficiencies in the methods and data related to both of the above
concepts have been defined and an outline has been given for the research pro-
grams that are needed to correct the voids.

A comparison of five lift/propulsion concepts is made in terms of their applica-

bility to a near term Medium STOL Transport airplane. The concepts studied
were Externally Blown Flaps, Internally Blown Flaps, Mechanical High Lift
Devices, Augmentor Wngs and Direct Jet Lift.

I
 1. INTRODUCTION

This survey of the state-of-the-art of STOL aerodynamics was initiated by the

necessity of providing information concerning the technical base for the design
of STOL aircraft. In particular, the aim was to assess the current state of the
art of aerodynamics as applicable to near-term STOL transport aircraft.

The approach taken in the survey was first to limit the survey to include only
those STOL concepts that were in a reasonably advanced state of development
and could therefore be considered as likely candidates for the near-term STOL
transport role. Seven concepts were chosen and are listed in the summary.
Next, the important technology areas of STOL aerodynamics and the information
required in evaluating the aerodynamic characteristics of a STOL aircraft design
were defined and used as guidelines in the conduct of a literature search. The
literature search resulted in a bibliography of about 900 references, most having
abstracts. The results of the literature search are classified into technology
areas and STOL concepts and are presented in Volume II of this report.

From the outset of the study the concepts of major interest were the externally
blown flap and deflected slipstream types. The literature search revealed
several prediction methods that were applicable to calculation of the charac-
teristics of deflected slipstream configurations, but none applicable to the ex-
ternally blown flap concept. Consequently, it was decided 'at the major
emphasis of the remainder of the study should be placed upon the development
of a suitable prediction method for such configurations. Such a method, based .
upon jet flap theory has been developed.

The analysis of methods and data related to the deflected slipstream concept
consists of a critical survey of a number of prediction methods and correlation
of a small amount of flight and wind tunnel test data with one of the methods.

The work related to the externally blown flap and the deflected slipstream con-
cepts is presented in two separate parts of the report. Each self-contained
portion (Sections 4 and 5 of the report) includes an assessment of the gaps and
deficiencies in methods and data, recommendations for corrective programs
and a list of related references.

2
I
S2. FACTORS IMPORTANT TO THE DESIGN
AND OPERATION OF STOL AIRCRAFT

'iT 2.1 INTRODUCTION

Short takeoff or landing performance requirements impose special problems that
j[ are unique to STOL aircraft. This section includes a review of the aerodynamic
factors that influence the design and operation of the airplane. These factors
include the basic aircraft requirements for obtaining STOL performance,
criteria and margins for safe operation, characteristics of the airplane in
ground effect, and testing requirements and techniques.

2.2 REQUIREMENTS FOR OBTAINING STOL PERFORIMANCE

In the context of this study the term STOL includes only those types of airplanes
that achieve short takeoff and landing performance without unduly sacrificing
cruise performance. In this sense, successful STOL transport aircraft are
characterized by good cruise speed capability and comfortable ride qualities.

To obtain short takeoff performance the airplane must either takeoff at a very
low speed or employ a sufficiently high thrust level to accelerate to the takeoff
speed in the required distance. In fact, both measures are taken, takeoff speed
being lower and thrust/weight ratio being higher than for a conventional aircraft.
Figure 1 illustrates a typical relationship between takeoff distance and a param-
eter combining takeoff speed and thrust/weight ratio for a given configuration.

The attainment of short air distance during landing results from a steep ap-
proach path and ground run is minimized by touching down at the lowest pos-
sible speed and obtaining the maximum deceleration.

Figure 2 indicates typical approach speeds required to achieve short field

landing performance using two different sets of ground rules for calculating
the distance. Figure 3 shows the results of a parametric study of landing per-
formance in which the major factors that affect landing distance were syste-
matically varied. Starting from a base point represerting an airplane designed
for conventional landing, a number of things can be &kncto reduce its landing
distance. However, large changes in parameters such as reverse thrust,
ground roll friction and approach angle, have relatively little effect, while re-
chction in stall speed is very powerful. The conclusion is that in order to
ichieve STOL landing performance, the airplane must be designed to fly slowly.

Figure 4 shows the combinations of wing loading and usable lift coefficient re-
qired to achieve the low speed operation required for short takeoff and landing.
it is seen that either low wing loadings - of the order of 65 psf or less - or

I powered lift systems are required to achieve landing distances of 2000 feet.

!3

I - , ,i i
 12 I
I
10~ X VTO/(T/W)
VT IN KTS

8
TAKEOFF
DISTANCE 6
-1000 FT
4~ o

0 2 4 6 Xe 10 12 14

FIGURE 1. VARIATION OF TAKEOFF DISTANCE WITH TAKEOFF SPEED

AND THRUST/WEIGHT RATIO

4 1f
I

* 160

* 120 LANDING DISTANCE

APP ROACH s
SPEED - KTS 80

- \
'~FIELD LENGTH
40 INCLUDING FACTOR
OF 0.6 FROM
]'"FAR PART 121

0
0 1000 2000 3000 4000
LANDING DISTANCE OVER 50-FT OBSTACLE - FT

1FIGURE 2. VARIATION OF LANDING DISTANCE AND FIELD

LENGTH WITH APPROACH SPEED

I5
I
I
I
i! S
 En

..110- I
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H HJ
44, ~
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to)
 8

100 2000' FIELD LENGTH

W/S=20 40 6
60 40
LIFTHIHLF SYTM
1 COEFF.4

2-

II0
II0 20 40 60 80
VELOCITY KNOTS
100 120 140

H FIGURE 4. LIFT COEFFICIENT FOR LEVEL FLIGHT

I '7
Airplanes relying on conventional aerodynamic control surfaces have a limit on the
minimum flight speed because of the low dynamic pressure achieved. Thus, to
attain speeds sufficiently low to perform STOL maneuvers, new techniques are
required for aircraft and flight path control. Figure 5 illustrates the minimum
usable speed for a variety of STOL aircraft expressed as a function of the wing
loading.

It is seen that to achieve the low speeds required for STOL performance while
using wing loadings in the 60 psf to 100 psf range, new control techniques
utilizing propulsive augmentation of the aerodynamic controls are required.

The effect of wing loading, aspect ratio and wing sweep on the vertical accelera-
tion an aircraft is subjected to when it flies into a vertical gust of one foot per
second is presented in Figure 6 for an aircraft flying at 300 knots at sea level.
it shows that low wing loading leads to high gust sensitivity and is the most
influential parameter of those shown. The Ag due to a given gust varies as
1/(W/s) for any given flight speed.

It is therefore seen that to meet the basic STOL objectives of short field IJ
operation - which requires low speed capability - and comfortable ride qual-
ities, one solution is to use high wing loading to minimize gust sensitivity and
a powered lift system to achieve low speeds.

Another solution would 1-e to use a lower wing loading and employ a gust allevia-
tion system to ensure necessary ride qualities. This solution reduces the thrust
requirement for powered lift STOL systems and provides a natural means for
achieving acceptable noise levels.

Figure 7 shows the reduction in g per fps of vertical gust velocity that can be
obtained by the use of a LAMS-Qaoad alleviation and mode stabilization) controller.

STOL aircraft designed for military application and inte, d for operation from
rough semi-prepared sites have additional design requirements. A high degree
of agility and maneuverability on the ground is required if the aircraft is to be
placed in revetments or other concealed areas and for movement in confined
spaces. Recirculation of flows due to use of reverse thrust on semi-prepared
strips can lead to significant foreign object damage to engines.

2.3 CRITERIA AND MARGINS

2.3. 1 Performance Criteria

It is necessary, for safety of operation, to ensure that the aircraft maintains a

sufficient margin away from any dangerous condition which could occur due to
possible deviations from the normal operating condition. These deviations in-
elude atmospheric gusts, critical engine failures, and wave-off or other required

8
 Jet
C
C\J 2 .0
Transports

120 Conventional~
Controls
Adequate

u~100 1D
II Mohawk KPCLt:5.0
1 80 Caribou
>l BLC-C130cot 0
Light

H ~60 P~

Brcguet 00 Minimum Safe Speed

W~~ 4091minimum Power On Speed
Power Augmente W/s
20 Controls Req'dV =1.69 CLp2

H XV5A
XC-142

U0 20 40 60 80 100 120

1] WING LOADING, W/S - LBS/Sq.Ft.

I] FIGURE 5. MINIMUM SAFE SPEED

1 9
 F
VERTICAL
ACCELERATION
g's PER fps H
V = 300 KTS S.L. STD.

-c/2 = 00°
---
.20 Acl 2 = 26.60

.16 ASPECT RATIO=10

ASPECT RATIO = 6
S1,ASPECT RATIO = 4

.12

.08H

.04 COMFORT
ZONE

p a I , a
0 20 40 60 80 100 120

WING LOADING, W/S PSF I

FIGURE 6. GUST SENSITIVITY OF RIGID-BODY MOTION I
I

I
I- -4)

u)

xI 0

Ln0

HlH
0

Ell 04 m
r34
tg Z

u H

-W U _ _ _
flight path corrections. Gusts effectively change the flight speed (longitudinal
gusts) or angle of attack (vertical gusts), imposing the requirement for speed
and angle of attack margins. Flight path adjustments include transient altitude
cuanges and turns. These impose a requirement for a margin on normal load
factor available. For conventional aircraft, normal load factor, angle of attack,
and speed are directly related. Therefore, these required margins are tra-
ditionally ensured by application of a specified margin on flight speed above
stall speed.

Since the lift of a STOL airplane can depend strongly on the power setting as well
as the flight speed and attitude, it is necessary to develop STOL criteria that
differ from those applied to conventional aircraft, Angle of attack, speed, load
factor, and power margins must be independently applied for the STOL aircraft.

Figure 8 shows the influence of STOL margins on the flight envelope of an air-
craft. While the aircraft represented is a four engine turboprop, the principles
it illustrates are valid for any TOL airplane. Presented in the top portion are
steady state rates of climb or sink for various power settings as a function of
speed. Also shown are lines of constant flight path angle. On the bottom portion
of the figure the maximum load factor attainable versus speed is shown for
1, 2,3 and 4 engine operation.

The criteria for this aircraft were that it be able to pull 1. 2g with all engines
operating, 1. 1g with one engine out and have a 3-degree climb-out angle with one
engine out. For this particular aircraft configuration and weight it is seen that
the 1. 2g requirement with all engines operating is least critical, with a mini-
mum speed of about 23 knots, the 1. Ig requirement with one engine out is more
critical and that the 3-degree climb-out angle is the most critical require- It
ment as it establishes the highest speed of about 40 knots in order to be satisfied.
A required margin on flight speed is not shown but could have been included in a
similar manner.

Compared with conventional aircraft flight experience, STOL flight experience is

relatively sparse. STOL experience in the approach maneuver is summarized U
in Table I from Reference 1. The table illustrates maximum descent angles,
and minimum flight speeds for eight STOL airplanes that have been flown by
NASA pilots. Also included are the approach speeds used, ratio of approach
speed to minimum speed and the margin of approach speed over the minimum.
The table also indicates the factors that determine the minimum speed. It should
be noted that the minimum speed Is n necessarily fixed by CLma x .

Factors such as those above are interpreted into ground rules which strongly
affect the stated or advertised STOL performance as it is calculated for any given
vehicle. These ground rules form the basis for standardizing the comparison
between competing concepts and include such factors as obstacie height, rate of
descent at touchdown, load factor in flare, if a flare is used, time delays between

121

tI
 I

S6000
Y30

R TE OF -100%
400 Y-60 PWR
CLIL B
(FPM} ]' =90 °75%
,

50%

2000
BUFFET
"-
RlATE 0r ,S
SINK

(FPM) 6 %H EAVy
- "V
40 00 - 600'
, 4000[ 7560 BUFFFT (ESTItATED)
-300
5;0 5

Li1MAxIMUll25
TO LA G
FGACTOU,,,S
i 1.0
ILOA 0 40 60 an loo 120
SPEED ALONG FLIGHIT PATH - KNOTS

B FIGURE 8_ STOL MARGINS

'13
 1.

touchdown and activation of brakes, spoilers, etc., deceleration on the ground and
relationship between field length and actual takeoff and landing distances.

TABLE I 1,
SUMMARY OF STOL FLIGHT EXPERIENCE

', Vmin Limited Va AVa is

Aircraft deg knots by Vmin KTS

-1 64 Stall 1.17 11
YC-134A -4 68 1.15 10
-9 77-1/2 1.24 19
-1 55 Stall 1.15 8
-2 56-1/2 1.15 10
NC-130B -4-1/2 61-1/2 1.14 9
-8 68 1.20 13
VZ-3RY -16 36 Lateral 1.11 4 1

UF-XS -3-1/2
-5
Control
Stall a
CV-48 -6 47 Stall 1.17 8
-7-1/2 51 1.18 9
-4 49 1.16 8
BR-941 -5 50 Vmin 1.15 8 i
-6 52 1.14 7
-7-1/2 53 1.13 7

367-80 -3 75 Stall 1.2 15

Figures 9 through 12 serve to illustrate the sensitivity of takeoff and landing per-
formance to ground rules applied to the execution of these maneuvers.

One of the most important relationships is that between the flare maneuver and an
acceptable sink speed at touchdown. It is possible that for STOL aircraft a flare
should not be used but instead that the allowable slnk speed be increased and the
landing gear designed, at a cost in weight, to absorb the higher landing loads.
The reason that this philosophy may apply Is because the distance covered in
the flare maneuver is so dependent on individual pilot technique. If a flare is to [I
be used special instrumentation may bG required to esable the pilot to start the
flare at very precise heights and speeds and to pull the correct g during the flare
in order to have consistent and repeatable landing distances.

14 (I
 1.lg FLARE, VSINK-- 3 FPS

SLICK

£4 HARD RUNWAY
U) WET SIT
z

UN~.P PEPARElD

I DRY SIT

1 0

( 0 .

2 0

C4

DIS~IPNCt ovrp 50 rT L'EFT

I FIGURE 9. SENSITIVITY OF GROUND RULES

115
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a.OWJ C:) 00 Wi
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I10 nC)WNC
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1000 1500 2000

DISTANCE OVER 50 FT ^-FEET

FIGURE 12. SENSITIVITY OF GROUND RULES TO

TRANSITION TIME DELAY

18,
I
I Figure 13 shows the results of a study of the permissible takeoff speed, for a
typical STOL airplane, with a variety of takeoff criteria considered. It indicates,
for any given combination of wing loading and thrust-weight ratio, the minimum
permissible takeoff speed and the criterion that dictated the minimum speed re-
quirement. In this study the following criteria were considered:

11. 2g normal load factor - all engines operating

11. 1g normal load factor

VSTALL + 10 knots
one engine inoperative
1.2 VSTALL

7 = 30 steady climb

For the range of wing loadings and thrust-to-weight ratios studied, the normal
load factor requirements did not dictate the critical speed. As seen on the
figure, three distinct regions appear, in each of which one of the remaining three
criteria dictated the minimum speed. As would be expected, for low values of
thrust-to-weight ratio, the climb angle criterion dictates speed, with the stall
speed margins being critical at higher T/W.

2.3.2 Flying Qualities Criteria

Establishment of STOL flying qualities criteria is urgently needed for several

Ireasons. Many of the criteria that have been developed for conventional aircraft
are not applicable to the STOL regime flight characteristics and could limit
STOL potential without ensuring safety. The criteria should provide safety at low
speeds and reduce the pilot's work load during STOL maneuvers. It is important
that the criteria should specify sufficier t c'ntrol power to allow precise and
1repeatable maneuvers at low speed.

The main problems involving the selection of criteria are the small amount of
flight.experience and the variety of lift/propulsion concepts employed to achieve
STOL performance. The criteria for conventional aircraft have been established
over a long period from experience gained with many airplanes. Many con-
and intelligent
~different scientiousprivate efforts to
and government establish Due
agencies. STOLto criteria have
the limited been made
amount by
of flight
experience, many criteria, by necessity, are based on personal opinion and as a
result there are conflicting requirements.

Presently, no single flying qualities specification applicable specifically to STOL

aircraft is available although several sets of suggested requirements have been
compiled. Table 11 lists eight documents containing flying qualities specifications

3 19
 for conventional airplanes and helicopters and proposed recommendations for
V/STOL or STOL airplanes.

In addition, Reference 2 represents a concerted attempt by the regulatory agencies

and industry to produce an acceptable set of specifications. While the philosophy
of a desired criteria appears well understood, quantification of criteria has been
difficult and is incomplete at present. The most critical area of STOL aircraft J
at low speeds has been that of lateral control. This is particularly true of
powered lift systems for the engine-out condition.

TABLE II
SAMPLING OF FLYING QUALITIES SPECIFICATIONS F
9 A proposed military specification for V/STOL flying qualities -
Cornell Aeronautical Laboratory
• TND-5594; airworthiness considerations for STOL aircraft
* MIL-F-8785B; military specification flying qualities of piloted
airplanes
* NASA TND-331; an examination of handling qualities criteria for
V/STOL aircraft
* MIL-H-8501A; general requirements for helicopter flying and
ground handling requirements
* MIL-H-8501B (proposed); general requirements for helicopter
flying and ground handling requirements
• Agard report 408; recomm2ndations for V/STOL handling
~~qualities °

* USAAML TR65-45; suggested requirements for V/STOL flying

qualities

Table III shows a comparison of some important parameters which should be

specified when flying qualities criteria are defined; this listing is taken from
Reference 1.

The parameters that define control response are control power, force, linearity,
cross coupling, and apparent damping. The parameters that specify stability
and damping are also shown.

Since handling qualities are judged by pilots' opinions, the pertinent parameters
were chosen so that they could be easily recognized, appreciated, and quantified,
readily evaluated for compliance, and would include the effect of factors that
influence tie response or behavior of the aircraft.

20
I

ii WING LOADING - PSF

100
90
30 CLIMB ANGLE
S90 CRITICAL

H80 80

H 70

H 70 60
1.2 Vstal 1
i CRITICAL
SPEED-
CRITICAL

KNOTS
Ho
60

CVstall + 10 KTS
50 CRITERIA CRITICAL

] 1.2gn
l.lgn
All engines

Vstall + 10 kt 1 engine out

1.2 Vstall ,

I 0.4 0.5 0.6 0.7 0.8 0.9 1.0

STATIC THRUST/WEIGHT RATIO

FIGURE 13. EFFECT OF CRITERIA ON MINIMUM TAKEOFF

SPEED FOR A TYPICAL STOL AIRPLANE

J 21
I
 TABLE III
COMPARISON OF FLYING QUALITIES PARAMETERS
(FROM REFERENCE 1)

CONTROL RESPONSE:

Parameters to be Measured in
Item Roll Axis Yaw Axis Pitch Axis

1. Control
Power
Time to 300
Bank Angle
Steady-State
Sideslip Angle
Time for 100
Attitude Change
11
Roll Accel. Time for 150 Pitch Accel. Ii
within 1/2 SEC change in heading within 1 SEC
Max. Control Max. Pedal De- Max. Control ru
Deflection flection Deflection

2. Force Max. Force to Force to Achieve Max. Force to

Achieve Item 1. Item 1. Achieve Item 1.

3. Linearity Roll Accel. per Variation of Side- Pitching Accel.

Unit Stick Deft. slip Angle with per Unit Stick
Pedal Deflection Displacement

4. Cross (AW/AO)MAX Effective Dihedral

Coupling A0 Response About
an/g Longitudinal Axis

5. Apparent Number of Con- Number of Con-u

Damping trol Reversals trol Reversals
to Stabilize to Stabilize

STABILITY AND DAMPING: L

Lateral- Directional Pitch-Axis
1. Stability Directional: Period of Oscillation Insufficient infor-
Spiral: Time to Double Amplitude mation available
Dihedral Effect: No Criteria, In-
sufficient Information
2. Damping Directional: Time to Half Ampli- Level of M a
tude should not be so
Lateral: Roll Time Constant low that short pe-
riod motion is
aperiodically
divergent

22
!
I It should be noted that this treatment is very preliminary and should be modified
as new concepts of STOL aircraft with more advanced control and stabilization
systems are developed and tested.

Table IV compares the control response and stability and damping criteria from
Reference 1 and the proposed military specification for V/STOL flying qualities
from Cornell Aeronautical Laboratory. This comparison is for the lateral axis.
As can be noted, there are many areas where the proposed criteria are similar,
Band other areas where they are dissimilar. A comparison of the criteria about
the directional and longitudinal axes would show other points of agreement and
Idisagreement.
Figure 14 compares the maneuver control power criteria from several sources
for the roll, yaw and pitch axes. These data were obtained from the sources
I which are referred to in the figure. Two facts should be noted. First, these
data apply to an attitude or rate type control system. Other types such as trans-
lation control have no requirements defined to date. Second, some of the proposed
requirements are independent of aircraft gross weight while others vary with
weight.

II Further work is necessary in this area to establish reasonable criteria for a wide
variety of STOL lift/propulsion concepts. Additional research must be per-
formed to define the gust, wind shear and crosswinds that are encountered in
STOL operation. A systematic study should be made to relate control system
operational characteristics to control power, sensitivity and damping especially
in the IFR condition. Finally, more flight experience is required to define
Imethods of operation in the terminal area and to define the guidance and dis-
plays needed.

2.4 GROUND EFFECT

Flying close to the ground causes modification of forces and moments due to
Idistortion of the flow field around the aircraft.
Figure 15 illustrates the theoretical variation of lift in ground effect of a straight
wing of aspect ratio 7. The prediction, based on a planar horseshoe vortex,
serves to show that the magnitude of the ground effect increases with the amount
of lift generated by the wing. At the high lift levels (CL 4) attained by STOL
aircraft the planar horseshoe vortex model is an inadequate basis for predictions
because of the large flow distortions involved and the complex interaction of the
lifting and propulsion systems. Nevertheless, Figure 15 indicates that large lift
variations in ground effect are possible for the STOL airplane. In addition to
significant effects on forces and moments, the recirculation of flows created by
the STOL airplane can cause hot gas reingestion, foreign object damage, and
erratic dynamic motion of the aircraft.

1 23

II
 TABLE IV [
COMPARISON OF LOW SPEED FLYING QUALITIES CRITERIA
(REFERENCE 1 AND CAL V/STOL SPEC)
LATERAL-AXIS
CONTROL RESPONSE

Parameter to Level for Safe Operation (PR = 3. 5)

Item be Measured NASA TND-5594 CAL V/STOL Spec.
Control Time to 300 No More Than No More Than 1.8 SEC
Power Bank Angle 2.4 SEC
Roll Acceleration More than 0.4 More than ,_. 36 RAD/
within 1/2 SEC RAD/SEC 2 SEC 2
Ma,:imum Control No more than 600 No more than 600 wheel
Deflection wheel deflection deflection
or 5" of stick de-
flection
Force Maximum Force to 20 LB 3.3 LB -* 15 LB. j
Achieve Control
Power
Linearity Roll Acceleration Should not in- No objectionable non-
per Unit Stick De- crease linearities
flection I
Cross (A P/A 0 )MAX 0.3 Not excessive
Coupling A8 Not noticeable
an/g Less than -0.1
Apparent Number of Control No more than 2 -"
Roll Reversals to
Damping Stabilize.
LATERAL-DIRECTIONAL STABILITY AND DAMPING
Lateral Roll Time Constant Less Than 2 SEC Less Than 1 SEC
Damping
Dihedral Lateral Stick Dis- se >0 > 0
Effect placement
600<.5 62 MAX 6200MAX<0. 75
6 LMAX
Spiral Time to double Not less than 20 Not less than 20 SEC
Stability Amplitude SEC
Dutch Frequency-wnd - tdwad Less than - tdwand Less than zero
Roll
Damping Ratio-tv
~
-. 09
d greater than d greater than .08
1
10

24
13.0

TIMEC
GMIL-H-8501A
2.0
2.IM AGARD 408
PR=3.5 NASA TND5594
-'- CAL V/STOL SPEC

71.
42 .-- _ ___

I .
0
0 20 40 60 80 100 120 140

~2.0
0 0
0 20 40 60 80 100 120 140

I0'
II25 1.0

FIGUR 140OTO OWRCMAIO

I0
I
20
. . I

40 60
i p

80 100
a
120 140

I GRCSS WEIGHTi 1000 LB

] FIGURE 14. CONTROL POWER COMPARISON

1 2

Il u. . .i T - I - -i - ii • , - ., . . .
 0

11aU) '
0
$47
x

N0 L
00

#C4 0
0 E-' *

LnL
H

LML

TI
01
1.4 -'4

28
J Because of the variety of concepts employed for STOL aircraft, it is probable that
each concept should be treated separately in calculation of the ground effects.

AA typical measurement of the effect of configuration differences is shown in

Figure 16. These data were obtained from wind tunnel tests of models of the
Boeing 727 and 747 airplanes using a fixed ground board. It should be noted that
the 747 test shows greater ground effect with tail on than the 727 test does, even

i1 though the CL was smaller. In addition, the difference between tail-off and tail-
on configurations is difficult to explain.

Recirculation of air has caused problems with some STOL airplanes, resulting
in loss of directional control due to asymmetric drag changes occurring when the
airplane was banked in ground effect. This roll-yaw coupling is illustrated in
Figure 17 for the XC-142 configuration.

IIf the airplane is banked in ground effect the drag of the lower wing is reduced
more than that of the upper thus producing adverse yaw. The tunnel tests indi-
cate that at a wing incidence of 45, a flap setting of 60" and a speed of about 25
knots, the adverse yawing moment will equal the yawing control moment avail-
able to the pilot at a bank angle of about 7.5 degrees.

iOther ground effects that can occur include damage to and/or loss of power from
the engines due to reingestion of hot exhaust gas, foreign object damage due to
ground erosion by high velocity slipstreams and localized heating of the airframe
by recirculation of exhaust gases.

Testing for ground effect in a wind tunnel can introduce the requirement for a
moving ground board in order to avoid the strong interaction that would other-
wise occur between the disturbed airflow about the airplane and the boundary
layer on the tunnel floor. Figure 18, from Reference 3, presents the results
of a study made to evaluate the criteria for determining the conditions under
which the moving ground board is required. It is seen that if CL is greater
than about 4 (h/c) a moving ground board is required. This figure does not in-
elude concepts employing concentrated jets. Reference 4 indicates that moving
belts are not required for those concepts.

2.5 STOL MOR&L TESTING

225.1 introduction
It has long been recognized that testing models of high lift configurations re-
quires special modeling and testing techniques different from those employed
for conventional aircraft models. The use of powered lift systems that obtain
high lift from strong interactions between the lifting and propulsive systems
dictates the use of powered models. Furthermore, the large flow deflectioas
associated with high lift require that models tested in wind tunnels be small

1 27
I
 10
'00 U) -
ru r4 :3 U
..- 0
4 14.m

4'4 '
Ed 0

C q
OJ a) ,.e- - IN0

9: r < % r o0O 0

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ro 0

a'. 02

00
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'0
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z N

0 5.4'
*n 0

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23
 al 8

IfI

HH

29/
 WW

4.a -H 4

o4 ~M 4.o3,
0
it, OC ) 4J -in

N n0 e v 0

~~1 ~ r-4 O
9

a)Z Z zr
0) ~ E- ~
ar -ir et -

0 < I

H E-'

Wx E-'Z
z H

0 14
NE-1
A
44rz

;4 04
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030
I
relative to the tunnel working section or that adequate predictions of wall cor-
rections be available. These and other considerations have led to the construc-
tion of large wind tunnels an-' ther suitable test facilities, the development of
devices to simulate propulsion systems, new sophisticated instrumentation and
spurred a continuation of the theoretical investigation of wind tunnel wall effect
[.q corrections.

2.5.2 Simulation of Propulsion Systems

i Several different STOL high lift systems have been proposed using a wide variety
of propulsive devices including turboshaft driven propellers, turbojets, turbo-
fans and various systems employing gas generators to supply jet flaps, fans,
[etc. The requirement to simulate the operation of these systems with model-size
motors can lead to significant complication of model testing. Motor size/model
size matching is complicated by the requirement to provide high thrust coef-
ficients without testing at extremely low tunnel speeds. Motor heating can cause
interaction with balances, requiring careful attention to temperature compensa-

Iof tion. Accurate means must be provided to determine propulsive forces. Routing
air supplies around balances requires that unusually detailed attention be paid
to accurate, thorough calibration for data corrections. The large amount of
powered testing performed on tilt wing and deflected slipstream configurations
has led to the development of successful test techniques for propeller-driven
concepts. The important considerations of such testing are described in more
detail in Section 5.5.2.

2.5.3 Test Conditions

IIThe high lift levels and large momentum deflections associated with STOL
configurations at low speed can result in large wind tunnel wall effects, not
predictable by the conventional wall correction methods, unless special atten-
tion is paid to model/tunnel matching. In addition, very small models are in-
appropriately scaled due to practical limits on model motor size, and the large
models are required in order to obtain high thrust coefficients at acceptably high
Reynolds numbers. This has led to the development of large V/STOL wind
tunnels which have adji -table tunnel wall configurations - open, slotted, or
closed - and with moving or fixed ground boards. Table V lists the major
V/STOL wind tunnels currently in existence in North America.

Even in the larger tunnels a potential exists for significant flow distortion when
the test conditions create large deflections of tunnel flow. Figures 19 through
22, taken from Reference 5, show the progressive buildup of flow distortion and
interference fields with increasing downwash. A method of monitoring the tunnel
flow conditions to determine the onset of large flow distortions has been recom-
mended in Reference 5. The technique consists of observing the pressure dis-
tribution along the centerline of the tunnel floor. As shown in Figure 23, the
pressure distribution is rel .tively smooth when the flow distortion consists only

31
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40
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0

I~ WA

I O~z

0
HOrT

Hz u

I 0) UO
ow~

1 4, 0

1 35
 E-1~

z
0

0 0
E-HE-
KE4
E-4 Z
z 0~

u 04

-44

H4

36
 E-
Iz

K! 0

E-1

~r4
* I x
A~ HE-4

0
>

L44

I 37
 * 0
* LL.

(L 0L
0
0 4 ~01

r4))

0)

I.0 z

Z H4

z Q

ww
Iw >

N 0 N

C-C)

38
Iof minor disturbances of the tunnel boundary layer (such as Figure 20) but in
Figure 24 becomes non-uniform at the point of incipient stagnation (Such as Fig-
ure 21). Reference 6 gives the results of a study to determine the largest model
ii to tunnel size ratio permitted and indicates that the flow downwash angle due to
the lift/propulsion system has a major effect on the acceptable model size.

Hl Because of the complex flow interactions between the freestream and the propul-
sion system, great care must be taken to measure all important parameters and
avoid unwarranted assumptions about parameters that might be expected to re-
main constant or behave in a predictable manner. Further, it is necessary to
calibrate balances, propulsion units, etc. on an individual basis.

1 2.6 IMPORTANT TECHNOLOGY AREAS AND INFORMATION REQUIREMENTS

To serve as guidelines for the literature search described in Section 3 of this re-
port, the important factors discussed in Sections 2. 1 through 2. 5 have been classi-
fied into three major important technology areas, and the information required in
each of these technology areas is listed.

2. 6. 1 Basic Aero/Propulsion Technologies

2. 6. 1. 1 Aerodynamic Characteristics of High Lift Wings

0 What are the lift characteristics for clean wing, such as C CCLo'
CLmax, etc ?

0 What are the increments due to the addition of trailing edge and leading
edge devices?

2.6. 1.2 Aerodynamic Characteristics of Other Airframe Components

* What is the influence of stores, nacelles, fuselage, etc. on maximum lift?

SWht is the influence of nacelles, fuselage, tail and other components on

, aircraft stability ?

2.6. 1.3 Aerodynamic Characteristics of Boundary Layer Control Systems

Li . What is the change in lift and maximum lift due to application of boundary
layer control and what is the influence of BLC parameters (Cp, Vj/V.) on
Iithe change in lift?

* Whit is the influence of RLC on aircraft moments and drag ?

H
-1 3
2.(.1.4 Operational Envclopes of Propulsion Systems

0 What are the operational envelopes of various propulsive devices as influenced

by inlet flow distortions, sand and dust environment and recirculation of hot
gases ?

2.6. 1. 5 Behavior of Propulsive Devices at Low Speed

" What is the variation of thrust of the propulsive system with power setting,
flight speed and angle of attack?

* What are the moments and in-plane forces due to the propulsive system?

* What are the flow fields created by propulsive devices in terms of velocities
and downwash angles ?

2.6. 1.6 Aerodynamic/Propulsion Interactions

" What are the lift, drag and moment characteristics of the integrated
aerodynamic/propulsive system?

* What is the influence of the high lift wing on the propulsive system ?

" What are the flow fields created by the combined aero/propulsion system?

2. 6.2 Interacting Technologies

2.,6.2.1 Stability

* What are the regions of adequacy of linear stability analysis or small

perturbation methods ?

* What terms uf the stability equation which are normally neglected for
conventional ;ircrift must be included for accurate analysis of STOL
aircraft stability and control ?

2.6.2.2 G round E rrects

* What is the proximity to ground, flow field, lift, drag,

influence of configuration, airplane momnts, stability, ?
weight, flow deflection, propulsive system
forward speed, power

40
 I Ii 0

I z LL -
LL0
0

* ~LL LLQ

1~1 II I:0~OO~E-4H
*z
z
DT7 0
aii -
HO0

71
*10
ai
11 Ei
H)00

IE-

w
wt*.

N aD

jI 41
2. 2. :1 Surface Effects

0 What is the disc loading, airplane pressures, flow

influence of weight, configuration, on velocities, visibility,
forward speed, groind ground erosion, damage
composition to structure, engine
damage

2.6.2.4 Atmospheric Gusts

0 What is the sensitivity of various STOL concepts to operation in gusting

environment ?

2.6.3 Systems Requirements

2. 6. 3. 1 Power Required

* What is the power required to operate various types of BLC systems ?

* What are the power requirements for aircraft control ?

What are the power penalties due to slipstream deflection or thrust

vectoring?

2. 6. 3. 2 Control Requirements

* What is the control effectiveness of various types of control devices

proposed for low speed STOL operation?
- What are the aerodynamic problems associated with the use of these

control devices ?

2. 6. 3. 3 Margins and Criteria

* What are the appropriate criteria for STOL aircraft?

* What is tlie aircraft performance sensitivity to the criteria used?

2.6.3.4 Test Requirements and Limitations

* What special model and test techniques can be used to simulate the full
scale conditions?

* How do you correct the test data to compensate for the difference between
test and full scale conditions ?
4I2

" IJ
I
I What combinations of test conditions could potentially limit the ability to
compensate for the difference between test and full scale conditions ?

0 What techniques can be used to monitor the test to assure the validity of the
~results ?

SI The above exposition of the necessary information was prepared in the form of a
series of questions to serve as a convenient check list for the information search.

The main emphasis in this study has been placed on aerodynamic problems of
STOL aircraft operating in the low speed flight regime. The technology associated
with cruise and other high speed flight modes has accordingly been de-emphasized.
The study has concentrated on the technology of the aerodynamics of lifting sys-
tems, propulsion interactions, stability and control, criteria and margins and
model test techniques.

1 43
I
 3. LITERATURE SEARCH

3.1 SCOPE

The literature search was made in order to identify the available test data and
prediction methods relevant to the aerodynamics of selected STOL aircraft con-
cepts.

Seven STOL lift/propulsion concepts were researched:

1. Externally Blown Flap

2. Deflected Slipstream

3. Jet Flaps

4. High Lift Devices

5. Fan-in-Wing

6. Tilt Wing

7. Jet Lift

The areas of aerodynamic technology for which references were sought were:

1. Forces and Moments

2. Flow Fields

3. Ground Effect

4. Stability and Control

5. Handling Qualities and Criteria

6. Testing

In addition to the seven STOL concepts and the six technology areas, a further
classification of "general" was included to cover references that were relevant to
some or all of the concepts or technology areas.

3.2 SOURCES

The main body of references was obtained by use of automated literature searches
carried out by the Defense Documentation Center and by The Boeing Company

44
 I
I Library. Other sources of references were the NASA STAR index, the Royal
Aeronautical Society Library Acquisition Lists and a number of AGARD bibli-
ographies.

Of the references that were originally thought to be relevant, by virtue of their

title (or in some cases abstract), many were found to be outside the scope of this
literature search and were, therefore, discarded. The resulting bibliography
contains references to about 900 documents with abstracts from most of them and
comprises Volume II of this report.

3.3 CLASSIFICATION OF BIBLIOGRAPHY

The bibliography is classified in three levels. In the first level, the references
are classified into I, Prediction Methods and II, Test Data. Within these groups
-' the bibliography is organized, as a second level, into divisions according to the
STOL concepts as listed above. The third level of classification, within the
'STOL concept' divisions, consists of subdivisions identified by the technology
areas named above. The numbering system employed in classifying the ref-
erences has ft r components.

E I The first, Roman Numeral I or II, indicates whether the data contains prediction
? A techniques or test data.

The second and third components identify the STOL concept and technology area
.j respectively, the numbers employed corresponding with the above lists.

The fourth component is the position of the reference within the technology area
subdivision.

.1
*1

-U
 4. EXTERNALLY BLOWN FLAPS

4.1 AN ENGINEERING METHOD FOR ESTIMATING THE TAIL-OFF LONGI-

TUDINAL AERODYNAMIC CHARACTERISTICS OF AN AIRCRAFT WITH
EXTERNALLY BLOWN FLAPS

4. 1. 1 Introduction

The externally blown or external-flow jet augmented flap (EBF) is a means of

increasing lift in order to achieve short takeoff and landing (STOL) performance.
The propulsive jet is directed at the trailing edge flaps in order to augment the
wing lift at low speeds. Part of the propulsive jet flows through the flap slot and
the entire jet is turned by the flap. Lift is increased due to the vertical momen-
tum of the jet and induced flow around the airfoil. A sketch derived from water
tunnel flow visualization work reported in Reference 7 is used to illustrate the
concept, Figure 25.

The externally blown flap integrates the propulsion and mechanical high-lift sys-
tems to produce a powered-lift airpl-ne. The EBF is related to the jet-augmented
flap in which the propulsive jet is expelled as a thin sheet from the trailing edge
of the airfoil.

The externally blown flap concept was introduced by the NACA in 1956. However,
little interest was shown in incorporating it into flight hardware due to the high
jet velocities and temperatures of the pure turbojet engine. The advent of the high
bypass ratio turbofan with its low jet velocities and temperatures advanced the
concept from wind tunnel 1easibility studies to consideration for incorporation into
practical flight vehicles.

To evaluate the worth of a particular EBF configuration, a rapid method for

estimating its aerodynamic characteristics must be available. Since no such
technique was found during the literature search, the major portion of Phase 2 of
the study, Analysis of Techniques and Data, was devoted to the development of a
suitable method.

In 4. 1 of this section, an engineering method for predicting the tall-off longi-

tudinal aerodynamic characteristics of an EBF configuration is formulated and its
correlation with existing experimental data is established. In 4.2, the gaps and
deficiencies in the available wind tunnel data are defined in detail and programs
to correct these gaps and to extend the prediction capability introduced in this
report are recommended.

4.1.1.1 Summary. Developing prediction capability requires that the problem be

amenable to theoretical treatment, or that a large body of test data be available
to develop empirical techniques, or that parametric variations of the significant
parameters have been tested. The wind tunnel test data that is available in the

46
!

/>4

0
b-~HZ

z a

I
4.44 HN

E-4
14
Na 44

I 47
open literature is limited in quantity and scope and it is impossible to develop a
strictly empirical approach.

No theoretical treatment of the externally blown flap per se exists. However, the
externally blown flap is a type of jet flap for which some theoretical development
has been done. The approach developed in this report is to use jet flap theory to
derive incremental effects due to power. It is assumed that the unpowered char-
acteristics are known from wind tunnel testing or can be estimated by standard
methods.

The jet flap consists of a sheet of high energy fluid ejected from the trailing edge
of the airfoil. Lift, in addition to that given by standard thin airfoil theory and
the lifting component of momentum, is induced due to the jet deflecting the
streamlines of the flow downward in the vicinity of the trailing edge. An inviscid
linearized solution has been obtained for a thin jet of momentum coefficient Cj
emerging tangentially at the trailing edge, References 8 and 9.

This treatment of two-dimensional jet flaps has been extended, References 10

and 11, to the case of a flat, finite aspect ratio wing with a full-span zero thick-
ness jet sheet emerging at a small deflection from the trailing edge. The equa-
tions have been solved for the special case of large aspect ratio and elliptic
loading. Theoretical treatment of part-span jet flaps has not been undertaken.

In order to use jet flap theory to predict the characteristics of the EBF, a ra-
tionale for determining an effective jet angle has been developed. Since the
spanwise extent of the wing influenced by the propulsive jet cannot be established
from existing whid tunnel data, the EBF data have been analyzed to show that the
lift and drag can be estimated with fair accuracy without knowing this extent.
This is not true for pitching moments on swept wings.

Jet flap theory is manipulated to a form in which the powered lift curve can be
expressed as the product of the unpowered lift curve slope and a function of the
blowing momentum coefficient. Jet flap theory is a small angle, linear theory,
i.e., sin a a . The direct power effect on the lift curve slope is equal to Cj.
For the large EBF jet angles normally used, the final expression for lift curve
slope has been modified to contain a direct power effect dependent upon Cj and
6 jo
CLa CLa K (Cj) - Cj + Cj Cos 6j (1)
poweredupwre

Similarly, the incremental effects of flap lift due to power have been derived from
jet flap theory

ACL (Cj) ACL6 (Cj) F (A, Cj) 6J/57.30 (2)

48
 I
I A correction to the direct lift power effect has not been made because the test data
correlates better without it.

:, Maximum lift increment due to power has been correlated as a function of the
vertical component of momentum.

The force polars are derived by showing that the theoretical jet flap induced drag
polar is equivalent to unpowered elliptical loading induced drag, CJ/ 7r A, with the
thrust deflected to the optimum angle for level flight.

The pitching moments can be estimated by extracting a center of pressure of the

7 power-induced flap lift from jet flap theory.

4.1.1.2 Organization of the Method. The effective jet angle, extent of wing in-
* i 2luenced by the jet, and static turning efficiency are discussed in Section 4.1.2.
Expressions for lift curve slope, flap lift increment due to power, and incremental
maximum lift are derived in Section 4.1.3. Force polar estimation is explained
in Section 4.1.4. Pitching moment is discussed in Section 4.1.5.

The effects of asymmetric thrust are considered in Section 4.1.6. It is shown

that the methods of the prior sections can be used to adequately predict the longi-
- tudinal aerodynamic haracteristics of the configuratiop with engine out.

An example illustrating the application of the method is given in Section 4.1.7. A

correlation of theoretical and experimental results are contained in Section 4. 1.8.
Since the method is derived to some extent from this same data, the final proof of
the adequacy of this method must come from applying it to data that will be ob-
tained subsequent to publication of this report.
4.1.1.3 Nomenclature

A Aspect ratio

C Chord

jCI Developed chord with flaps extended

CD Drag coefficient

CDPmin Minimum profile drag

I Cf Flap Chord

Ci Developed flap chord

Cj Engine exhaust momentum coefficient, rh VE/qS

1 49
I
CJ Jet flap blowing momentum coefficient

C1 Rolling-moment coefficient T

C 1, Lift coefficient

C i'
c Circulation lift coefficient

C IT Total lift coefficient

CLa Three-dimensional lift curve slope

CLa CLa Two-dimensional lift curve slope

ACL (Cj) Increment in lift due to power at zero angle of attack

ACL 6 (CJ) Increment in 2-D flap effectiveness due to jet flap effect

AC 1 max (Cj) Increment in maximum lift due to power

Cm Pitching moment coefficient

do Function of flap chorc and Cj, Reference 9

(Identical with Do from Reference 9)
Do Function of Cj, Reference 11

F (A, Cj) Jet flap finite aspect idaio factor

FA Axial force

FN Normal force

FR Resultant force

K (A, Cj) Ratio of powered to unpowered 3-) lift curve slope for a
jet flapped wing

Engine mans flow

Slt F Reference wing area, rt2

S Gross wing area with flaps extended, t2

So Flapped wing area, ft2

T Gross thrust

50
 VE Engine exit velocity

Angle of attack

.5
- Jet flap deflection angle

5. Effective jet deflection angle, tan- 1 FN/FA

I-p
6u Flap upper surface slope

61 Flap lower surface slope

Static turning efficiency, FR/T

* 1 Effective roll arm of engine lift/wing semispan

%eng Spanwise location of engine centerline/wing semispan

X 2 sin- lff/c
XCp/c Chordwise location of center of pressure

4.1.2 Required Information

Jet flap theory, in its present state of development, requires that the blowing be
.1 full span on the wing and that the angle and blowing momentum coefficient of the
jet as it leaves the trailing edge of the wing be known in order to estimate the
aerodynamic characteristics of a jet flapped wing. In order to adapt jet flap
,T theory to the prediction of wings with externally blown flaps, the effect of part-
span blowing must be assessed and an effective jet deflection angle and momentum
coefficient must be defined.

t 4.1.2.1 Extent f Wing Influenced by the Jet. The extent of wing influenced by
the propulsive jet of the externally blown flap can be determined only by having
detailed wing and flap pressure data from wind tunnel testing. To date, no such
pressure data is readily available. A limited amount of Boeing pressure data
taken on the flaps only, indicate that the loading on the flap is highly localized in
the vicinity of the nacelle centerline. However, following a mechanical trailing
edge flap, most of the lift is expected to be induced on the wing. The distribution
of the induced lift on the wing can" nt bc determined from available test data.
Therefore, a method was developed which is independent of the cxtent of the wing
influenced by the jet.

The NACA test data Indicate that the lift and drag of an EBF configuration are a
function primarily of the blowing momentum coefficient and only weakly dependent
on engine location. Pitching moment is more sensitive to engine location,

151
especially on swept wings. Figure 26 taken from Reference 12 shows a compari-
son at zero angle of attack of lift, drag and pitching moment with the engine efflux
distributed along the span and with it localized well inboard. On the basis of this ;
and other data, the assumption was made that, for the level of accuracy being
sought, a method could bc developed which was independent of the extent of wing
influenced by the jet.

4.1.2.2 Jet Deflection Angle. It is necessary to be able Lo determine an effec-

tive jet deflection angle. Since the propulsive jet of the externally blown flap is
diffuse rather than infinitesimally thin as jet flap theory assumes, there is no
well-defined jet angle. A jet angle deduced from static force testing will be taken
to be the jet deflection angle. It is then necessary to relate this effective jet
angle to the trailing-edge geometry.

if a wind tunnel model is tested statically with engines operating and trailing edge a.
flaps down, the axial and normal forces on the model due to the jet can. be meas-
ured. Assuming that there is no induced flow around the model and therefore no
induced aerodynamic forces, an effective jet deflection angle can be defined by
tan- 1 FN/FA, Figure 27. NASA has done a substantial amount of static testing to
determine the effective jet anglc. lhese data have been examined to correlate
the measured effective jet angle with the flap system geometry. The effective .
angle can be estimated by
6j=1/2 ( u 6 1) (3) [

where 6u is the upper surface angle and 61 is the lower surface angle, Figure
28. This correlates well with the NASA static test data as shown in Section 4.1.8.

All of the data used are those in ,:ihich care has been taken to obtain efficient
turning of the jet. Therefore, jet angles may be lower than this if care is not
used. The assumption is made that the effective jet angle does not change with
forward velocity or thrust level.

4.1.2.3 Static Turning Efficiency. As with any system that redirects the propul-
sive jet, losses are incurred. Once again from static testing, a turning efficiency
can be inferred by knowing the static thrust input to the system and measuring the V
resultant force, 7 = FR/T, Figure 27.

Static turning efficiency has been correlated, as determined from NASA static L;
tests, as a function of the effective jet angle. The data of Reference 13 indicate
that the static efficiency is indepenlent of t ast, Figure 29. Figures 30 and 31
shcw the correlation for double- and triple-slotted flaps. In general, triple
slotted flaps appeaw to be more efficient thrust vectoring systems although some
double-slotted flap systems perform equally well. These data represent the cur-
rent state-oi-the-art for the flap systems considered. Turning efficiencies much

52
 r iJACA TN 3858

S IX-JZT ARRANGEMENT

CD

TWO-JET ARRANGEMENT

-2

LIL

4m4

0 0

]6

00

0 EI7~70-2,-
4
0 6 C
5
 I ]

Le

ii

'j = tan - FN/FA

EFFECTIVE JET DEFLECTION ANGLE

= F/T
R/T

STATIC TURNING EFFICIENCY

I:

Li

FIGURE 27. NOMENCLATURE, EFFECTIVE JET DEFLECTION

ANGLE AND STATIC TURNING EFFICIENCY

54
Ii

I! ii,

IANGLE FIGURE 28. NOMENCIATURE, EFFECTIVE JET DEFLECTION

i55
II
 14ACA TN 4079
_____ _____
__6f, DEG

F. 0 55
Fr
~ C> 65
Fr 75

.6

Fr__ _ _ _ _ _

Fi

.2

0
0 10 20 30 40 50 60 70
Fi ,lb

FIGURE 29. EFFECT OF THRUST LEVEL ON STATIC TURNING

EFFICIENCY

56
I
I

H:H
-, o0 NASM TN D-A 04
1.0 0 NAS;
I0 NASI TN D-5 64ED )AA
UN-UBL-S

!! ' ' A BOE:ING TEST DATA

0.8

T, ,. 0.6 ,-

z 0.4

0.0 -

0 200 40* 60* 800 1000

ESTITED J0T DEFLECTION ANGLE , 6 4
FIGURE 30. FLAP TURNING FFICIENCY CORREATION

DOUBLE-SLOTTED FLAPS
57
 I pi

t.

iI

0 LAS TN D-5164
1. NAS NUNPUBLISHED DATA
A BOE G 7EST DATI

0.8

0.6

o.4 !

0.2 -

0 200 400 600 800 1000

ESTIMATED JET DEFLECTION ANGLE - j
FIGURE 31. FLAP TURNING EFFICIENCY CORRELATION
TTTPLE-SLOTTED
FLAPS

58

LI
 worse than those shown can result. This report did not attempt to determine how
to ensure good turning efficiencies.

In determining theoretical jet flap lift, the important momentum parameter is the
momentum coefficient at the wing trailing edge. Correlating test data, externally
-blown flap lift correlates best when the engine exhaust jet momentum
is used di-
rectly and losses are considered only in determning drag.

14.1.3 Lift

Jet flap theory is still in its early stages of development compared to unblown
wing theory. The theories all assume small perturbations of the undisturbed
flow. Small jet angles are assumed. Inviscid theories have been developed for
the jet flap in two and three dimensions (References 8, 9, 10, and 11). The two-
dimensional theory, Reference 9, is derived for bent flat plates. The three-
dimensional theories, References 10 and 11, have been restricted to flat, elliptic-
ally loaded high aspect ratio wings with full span constant flap deflection and con-
stant sectional momentum coefficiern. along the span. For a jet-flapped wing, the
high aspect ratio assumptions require that the wing plus curved portion of the jet
i be of high aspect ratio.

Despite these limiting assumptions, jet flap theory compares well with experi-
mental data that lie outside the range of the assumptions made.

Since the externally blown flap is a type of jet-augmented flap, jet flap theory has
been used as a base on which to build a iemi-empirical lift prediction method.

utically 4.1.3.1 Lift Curve Slope. From Reference 11, the lift curve slope of an ellip-
loaded jet flapped wing of high aspect ratio is

CL
a 2D (4)

LaL 0-2 IF-v

1+1 2 2C~ J
where Do is a function of the jet momentum coefficient, and is reproduced in

Table VI. Rewriting equation 4 and adding and subtracting 21r in the denominator.
fj CL ( rA + 2 CJ)
cL = 2D '(5)
2
L fA+ C La 2 D -8D 0o + 21- *

II 59
U
 Factoring out A/ (A + 2)

CLa A 1;
2D A+ A
7 _ 2C
J (6)
La = + (CL 8D 0 - 2)/ (A + 2) 6
2D

From Reference 9, the lift curve slope of a flat plate airfoil with a jet flap is

C= 2 r~l+l~lC1/2
CL 2 (I+ .151 Cj + .219 CJ) (7)
a
2D

Substituting this into equation 6 and rearranging V

1/2 2Cj
(1+.1 5 1C + .2 19 C) (1+ 7rA)

CL = 2 12 (8)
a (1+ 2 (.151 C1+ .218 C-4Do)/(A+2)

The leading term, 2 7 A/(A + 2), is the lift curve slope of an unpowered elliptic-
ally loaded wing so that the powered lift curve slope of a jet flapped wing with
small jet deflection can be expressed as ,

CLa = CLa K(A, CJ) M

POWER ON POWER OFF

K (A, Cj) can be taken to be a function of blowing momentum coefficient only, with
an error of less than 1% for A of 6-10 and Cj up to 10. K (A, Cj) is given in
Figure 32.

Since the theory is linear and assumes small angles, there is a direct thrust
effect on the lift curve slope included in equation 9 of a I Cj ( a + 6)]/a a = Cj.
To account for large jet deflection angles equation 9 will be modified by replacing
(a + 6 ) by the c:)rrect sin (a + 6)80o that the direct thrust effect is given by
a[Cjsin(a+ 6)]/aa _Cjcos (a+ A)=Cjcos 6at a =0 °

CL CL POWER OFF c- CJ + CJ o 6 (RAD' (0)

CL = CL aooi

CL CL K(A, C) -j + jo (DE

La 7.3 57.3 (1)

POWER ON aPOWER OFF

60

I.
 0

Ln0

ii
I I- - - - _E-4

T! E40
-v-.-----------------

-. - -K ii - oa

[1 -.------------- ~Ow
[1 ----- ---- ---- _P_

------------- Nr -4-P -4

rIW d j rvi a ssu: a mIg~ LN a ooj

LI ------ ----- '-1-. --
 ..

TABLE VI
RELATIONSHIP OF Cj AND Do FOR
LIFT CURVE SLOPE CALCULATIONS V

Cj Do
0i
0.01 -. 0008
0.05 -.0040
0.10 -. 0080
0.20 -. 0158 "
0.40 -.0318
0.50 -.0398
1.00 -. 0798
1.50 -. 1198
2.00 -.1600
3.00 -. 2402
4.00 -. 3204
5.00 -. 4008
10.00 -. 8034

4.1.3.2 Flap Lilt Increment. In Reference 9, the increment in lift on a two-

dimensional bent flat plate airfoil with jet augmented flap is derived as

C
L
2(X+ sin X+ 2vdo)6 (12) L.
where X, 2sin-1 v'fKan! d = do (Cf/c, C J).

NOW

CL 2 ( X + silnX) (13)
is Just the lift increment of an unblown bent flat plate airfoil. Therefore, the
effect of blowing on the derivative of lift with respect to flap deflection is given
by

ACL(CJ) = 4.do (14)

This is given in Figure 33 as a function of flap chord ratio and jet momentum
coefficient. From Reference 11, the relationship between two-dimensional and

62
'AA

IP
three-dimensional lift is given by

_F(A, C (15)
CL2 (CLa2D 871) 0-

+ r 2 C

This has been calculated for A = 6-10 and Cj = 0-5 and is plotted in Figure 34.

The incremental flap lift due to power is given by

ACL(Cj) = ACL (Ci) F (A, Cj') 5. S (16) B
L L 57.3S

where S' is the flapped wing area and C' is reference to S' (Figure 35).

4.1.3.3 Maximum Lift. The estimation of maximum lift coefficient is not

amenable to theoretical calculation. Therefore, an empirical correlation ap-
proach has been undertaken. It is reasonable to assume that the increment in
maximum lift due to power will be proportional to i.e indcaced camber. In turn,
this induced camber is expected to be a function of the vertical component of the
jet momentum, r Cj sin 6j. The data from a number of tests has been plotted
in Figure 36. The data generalizes very well with no discernable trends due to
aspect ratio, flap chord, or leading edge device. p
This curve should be used with discretion. A configuration having substantial
leading edge separation for tie unpowered case may yield increments in CLmax -T

due to power much higher than these due to the boundary layer control effect of
the blNing. Also, as the vertical component of jet momentum increases, a
given leading edge geometry ;/ould be expected to depart from this curve as the
leading edge is no longer able to support the pressure distribution required with-
out separation.

4.1.4 Force Polars

The theoretical induced drag of a finite Pspect ratio wing with full span jet
flaps and liptical loading is derived in Reference 10

CDi C2T/ (A + 2 Cj) (17)

where CLT is the total lift.

64
 I6
H..
. ..

I Ai co Iw
,. _

u ,\
0 0

T
I

0 W I
. 0
%00

...
I_ I

65.

-
H
 DEELPEciAPANIINGCHRD

S WIIG ARA,
FLAPPD

FLAP AND WINGCHRS-

FIUE3.NMNTRDEVELOPE:

CHORDS AND FLAPPED WING AREA

661
I
!
1

I 7.0

B II 5.0-
-

E 4.0

113.0- X:
0 0 BOEJNG EST DATk

2.0
2.6 - - N
r NAS
NAS, TN
TN D-5364
D-943
0 UNP BLI 'HED NASA DArA.
HH LEA ING EDGE BLWN1t BL

I0.0
0.0 0.4 0.8 1.2 1.6 2.0 2.4 2.8
VERTICAL COMPONENT OF MOMENTUM, ?CT sin
FIGURE 36. CORRELATION OF MAXIMUM LIFT DUE TO
POWER

I67
This has been shown, Reference 14, to be equivalent to i

CDi [Cc/7rA (1 + 2Cj/7rA) (18) [

where CLc is the circulation lift coefficient.

The force polar is

2
CD- CDpmin =-Cj + [CLc/irA] (1 + 2Cj/irA) (19)

It will be shown that this is equivalent to the polar given by the unpowered theo-
retical induced drag for an elliptical loaded wing with the thrust deflected to the
optimum angle for level flight. V
&9

Consider the unpowered induced drag polar given by

CDi C2c/rA (20)

Now the optimum thrust deflection for level flight is given by the vector normal
to the polar. Therefore,

= A CLc /ACDc (21)

tan 0 =-2 CLC/rA

Now
2 + 2 2 (2
AC + AC c=()

Solving equations 21 and 22 for ACDc

SCCc (23)

Therefore, the total drag polar is given by

2
C D "CDm c cj
C-
D -pin wA [1+ (2Cl/A) 2]1/2 (24)

68I
1
1Since the theory considers large A, the denominator of equation 24 can be expanded
and only the first-order term kept

I C CDpi CL/1 A ji[1 - 2C ../(.DA)

32
- +2C/rA C Q.E.D. (25)

There are no theories and little test data for a wing with part-span jet flaps.
Most of the externally blown flap test data have part-span flaps and what must be
considered unknown spanwise blowing extent.

The force polars for the available externally blown flap tests with spread engines
are given closely by the sum of the minimum profile drag, theoretical unpowered
induced drag, the ram drag, and the thrust deflected to the optimum angle for
level flight, Figure 37. For the externally blown flap, the thrust must be multi-
plied by the turning efficiency to account for the losses in the system.

Some test configurations have the engines moved well inboard to reduce asym-
metric thrust effects on the lateral characteristics. This would be expected to
yield a spanwise. load distribution which was highly concentrated inboard. These
data do not yield polars consistent with the method just outlined. If, however,
aspect ratio is based on flap span, then good agreement is again obtained.

4.1.5 Pitching Moments

None of the three-dimensional jet flap theories consider pitching moments. There-
J fore, two-dimensional theory will be used for a starting point.

Since the theories are linear, the total pitching moment may be expressed as the
sum of the pitching moment due to the unpowered potential flow lift and the pitch-
ing moment due to power-induced lift. The pitching moment about the leading
edge can, therefore, be written

Cm = CLT XcP/cIT = ACLcJ Xcp/clcJ + ACLfu Xcp/Clfu (26)

I Solving for the center of the power-induced lift

I T-- cj ""CLc
J

S69
!I
 I.
Cv
CL

mu DLi

E::
I
H

0
H
U

DRAG COEFFICIENT, C D

FIGURE 37. FORCE POLAR BUILD-UP

70
ISubstituting known functions
acm T acmfu
aCLT as
2 c j_ a Cc j
_C (28)

a8 as

The terms on the right are known from two-dimensional theory, Reference 9.
This has been solved for a range of flap chord ratios and momentum coefficients
and the results plotted in Figure 38.

For conventional flaps, the effect of finite aspect ratio is to move the flap lift
center aft. Finite aspect ratio effects for conventional flaps have been derived
in Reference 15 and are presented in Figure 39. It has been assumed that this
same effect will be experienced on the jet flap. The spanwise center of load due
to flaps is taken at the spanwise location of the mean aerodynamic chord of the
flapped portion of the wing.

There is a constant pitching moment coefficient increment at constant angle-of-

attack given by
ACM = ACL(Cj) " cJ(29)

There is also an increment in 8 Cmi*a a given by

I a\cm)= ( - (a i " ) (30)

I where Xc../r/a is taken to be for a flap chord to wing chord ratio of one
(C'/C' = 1.0).
f
Pitching moment appears to be the longitudinal characteristic most sensitive to
engine or blowing location and extent, Figure 26.

4.1.6 Asymmetric Thrust

The lois of an engine on an externally blown flap configuration will result in a

reduction of high energy air supplied to that wing to augment the lift. The asym-
metric thrust condition may result in a lo..s in lift and a rolling moment due to an
asymmetrical lift distribution.

1 71
i
 "

0.-

0.4

I I
F
0.0

0.2- - - - - -

! --1.0
0.1. 7- - -

0.0 1,0 2.0 3.0 4.0 5.0

SLOWING MOMENTUM COE" ICIENTI, C3
FIGURE 38. CENTER OF PRESSURE OF POWERED LIFT!i

7l2
 I I

IIn
IIY
04
II)\ V,=
kI E-4 En
Bu
II __ t04
01

0I

m m~~~ zN.

P4

CI
Once again, the extent of wing influenced by the propulsive jet must be known in
order to compute the aerodynamic characteristics with engine out. NASA test data
indicate that the lift and drag of a configuration with asymmetric thrust at a given
t tal jet momentum coefficient level is the same as the symmetric thrust case at
the same total Cj, Figure 40 (Reference 16). Tall-on pitching moments are
affected due to the changed downwash at the tail. Therefore, the methods of the
previous sections may be used without modification.

The rolling moment due to engine-out cannot be established theoretically without

knowing the extent of the wing influenced by the jet. Therefore, an empirical
correlation considering engine location has been undertaken. The rolling moment
can be related to the lift loss by

AC 1 AC (31)
L)
where n 1 is the effective nondimensional roll arm.

There is substantial scatter in the experimental data, part of which is due to the
fact that we must divide increments obtained from test data for which a small
absolute error in determining the lift and rolling moments results in large error
in determining the effective roll arm.

The qpanwise location of the inoperative engine is the primary consideration af-
fecting effective roll arm. Jnfluences which are also important are wing sweep L
and the spanwise location of the operating engine. Insufficient data is available
to determine secondary effects and only the spanwise location of the inoperative
engine is considered here.

While there is substantial scatter, the equation

1 eng (32)

3 tide a reasn able first approximation. F .ire 41.

74
 z 0

II 9-4

000

ICM

I~ b4

0 4 0
1 0

_ 75 _ _ _ _ _
 i I I

i
0.6 -

0.4 - - - - - -- - - - - -I

0 .2O

00

ENGINE CENTERLINE SPANWISE LOCATION

FIGURE 41. ROLL ARM OF FAILED ENGINE LIFT

76
I
4.1.7 Application of the Theory

I
]
AREF = 8.4 AFlaps Down 7.71

Flap Span 1i/b= .091 =o/b.732

SC/C = .279

C'/C = 1.145
2
SRE F = 7.,6 ft
2
I SFlaps Down = 8.0 ft

S'/S REF = .697

Cj = .59, 1.24
f = 30/60 u =410 6 =590

1
=090 Deg
C Flaps Down
Unpowered

I Unpowered wind tunnel data available

STEP 1. Calculate 6

S6j = 1/2 (6 u + sx) = 500

I STEP 2. From Figure 30 Find n

6 = 500, double slotted flaps ' ? = .665

1 77

IA
 me LO

0 q t4

N00

eq9-

C.,
om
0
0
LOV- i
bl cli 0
00 0

CI L.J
W

C43

to eq I
03 - %0
ft H ta
t P4

78 ~4~[
 0
0
0

0 zI
ire 0 4

IH 0 0 U

L
-44
0 0

0 0CAE

* II 0o

I ~No 0 0
0

I a 00

I 0
%

U 79
 0 LO
150

I 1? 0o
. 1

000

eq t
'Alto

QII
C4 to ~~00t #

9 eq k CO

~Y
IIko.*
P4 II *

80q
I
4.1.8 Test-Theory Correlation
The validity of a theory must be inferred by comparing theoretical estimates
with test results. In this section, test-theory correlation are presented. While
1scatter exists, it is not sufficiently large to invalidate the methodology developed.

Estimated jet deflection angle is correlated with the jet deflection angle inferred
from static power-on testing, Figure 43. Data from eight NASA tests shows that
the method correlates within about 10%.

Lift curve slope also correlates within *10% except for one NASA test which is
over-predicted, Figure 44. However, subsequent test data from that model agree
well with the technique used.

1The largest discrepancy in any estimated increment is in the lift increment due
to power, Figure 45. This may be expected since this increment depends on other
factors that must be estimated; the extent of wing influenced by the flap and the
jet deflection angle. The lift increment should be considered reliable to only
+20%.

ii The estimated center of pressure of the flap lift due to power agrees well with
test data, Figure 46. It would be expected that the flap center of pressure should
be significantly different for the same configuration with clustered inboard or
spread engines. However, Figure 46 indicates that the difference is no greater
than the scatter due to flap angle differences. Actual pitching moment, of course,
depends on estimating not only the flap center of pressure but also the flap lift
increment.

The longitudinal aerodynamic characteristics of two dissimilar configurationb have

been estimated using the methods developed, Figures 42 and 47. Figure 42
is the configuration treated in the example of Section 4.1.7. Agreement is fair
with the largest discrepancy in pitching moment level which is due to overpredict-
ing the flap lift.

14.1.9 Conclusions

3be The determination of the lift and drag of a wing with externally blown flaps can
made using jet flap theory with suitable empirical values for jet angle and
turning efficiency.

There is a need for more data in order to refine the prediction techniques and do-
termine the influence of some secondary variables. Pressure data is needed in
order to determine the influence of the jet on the wing so that part-span loading
3 factors can be developed.

1 8
 T

o ASA TN D-7( 04
a A MEVO 3-8-59L--1
0 AC -5.64__

v AS TIND-9j3

.0 NSm
~o80 --- -

60~

E4 40'-------------------------

0-

00 20* 400 60* so* 1000

ESTIMATED EFFECTIVE JET ANGLE, 4j
FIGURE 4- -FFECTIVE JET ANGLE CORRELATION

82
I-
I

0 ASA TN -7 04
~f.&.N
1
d ASA TN -5",64
0 OEING 7STIAT

V OEG 7EST AT
S.16x

.10.

14

ow

L
i-

.06 .10 .12 .14 .16 .18

ESTIMATED LIFT CURVE SLOPE, C
I FIGURE 44. LIFT CURVE SLOPE CORRELATION

t!8
 7.0
0 4ASA TN -5 64
--- [
~AS TN -658
3OEIG ST ATA
6.0 -V -C -N -01
00 SA TS -943

3.0 [

2.0/
0 ..... [2
1.-- ----

o 1.0 2.0 3.0 4.0 S.o 6.0 7.0

ESTIMkTED
FIGURE LIFT INCREMENT
45. CORRELATION DUEINCRE-4ENT
OF LIFT To POWER,DUE
&CTo
L (Cj)

POWER AqT ZERO ANGLE OF ATTACK

3----
 I

0.6 TN r-5- 64

0.51

0 TN4 -704 US REE IN AM ENC INES

0.6 01 TN-6058 SP Eb GNIN|

0.75 O G S A IJ
0.4 'I
IQ

I 0.6
i
-

nml in-
......
-

1 0.3 - - D-
------------ - -

0 1.0 2.0 3.0 4.0 3.0

BLOWING HOMENTUX COEFFICIET BASED ON FLAPPED WIG AREA, Cj
FIGURE 46. POWER INDUCED LIFT CENTER OF PRESSURE
CORRELATION

I
 ICP 13
0001

c,

r34

444 oo11

44
0

000

ad 0
0
44 0
0 o
0 o

* PC.~ 0

0L
U0

0 r-
OD QD i
 4.2 DEFINITION OF DEFICIENCIES AND GAPS IN KNOWLEDGE

4.2.1 Introduction

In the second part of section 4 the deficiencies in the available test and analytical
techniques are defined. Programs to fill these voids and to aevelop improved
analytical techniques for externally blown flaps are recommended.

I] The externally blown flap concept has been tested by the NASA and others. In a
recent working paper, the NASA have summarized the objectives of the wind
tunnel testing to date and have noted the work they feel needs to be done.

"Although considerable wird-tunnel research has been conducted

on the external-flow jet-flap concept, the objective of most of the
work in the past has been maixdy to explore the general area of
performance and stability and control - with particular reference
to problem areas and to finding practical solutions to the problems -
such that the overall feasibility of the concept in terms of practi-
cal reliable application could be accurately assessed. This re-
search has provided the necessary information to show that the
*1 concept is effective for providing high lift on turbofan STOL air-
craft but has provided very little information relative to the
optimization of the jet-flap parameters involved. Because of
the increased interest at the present time in the jet-flap concept,
there is now a need for more detailed information for the
rational design of such systems. The overall objective of this
investigation is to provide the basic design information on the
effects of geometric variables such as wing planforri, engine
location, thrust deflectors, flap span, flap size and type,
leading-edge high-lift devices, and horizontal and vertical tail
location".

The NASA have established that the externally blown flap is aerodynamically
feasible and should be con. idered a candidate system for a STOL airplane. The risk
and cost of such an airplane can be reduced by further rational technology development.
In order to design an airplane using the externally blown flaps, it is necessary
to be able to estimate its aerodynamic characteristics with a reasori-ble degree
of accuracy. This requires theoretical, empirical, or semi-empirical esttmrtion
techniques substantiated by test data. The gaps and deficiencies in the available
test data and analytical methods will be defined.

4.2.2 TestDta
While the NASA have tested a range of representative configurations, a systema-
3tic seri" of tests intended to provide a basis for design of ar airplane with ex-
ternally blown flaps has not been done.

I * 87

!.
4.2.2.1 Modeling. An externally blown flap wind tunnel model differs from a V
conventional high-lift model only in the requirement for an engine simulation unit.
A variety of thrust units have been used on the existing tests: blowing nacelles
with domed inlets, ejector powered nacelles, and rotating machinery. If the I.
engines are located such that the inlet flow has little effect on the wing flow
fields, the criteria for a thrust simulation unit must be that the exit conditions be
properly simulated; scaled exit geometry and correct exit momentum coefficient.
This gives the proper relationship between the jet and the trailing edge flap system.
In addition, an exacting calibration of the engine simulation unit must be made.

The available testing, aimed at different test objectives, have not in general
used a calibration of sufficiently high standards to obtain accurate design data.

4.2.2.2 Test Conditions. Most of the test data %vailable in the open literature
have been obtained at very !c-v dynamic pressures. This has been dictated, in
part, by the use of existing thrust simulation units to obtain thrust coefficients
representative of high thrust-to-weight aircraft.

The majority of the available data were tested with model-tunnel combinations for
which wall correction theory is adequate and flow breakdown should not be a
problem. In future testing in other tunnels, model size may be limited by the
size of the wall correction that should be allowed. Tunnel flow conditions must
also be monitored to assure that the data is not affected by flow breakdown. i

4.2.2.3 Ground Effect. There is as of this writing, no published ground effect

data for the externally blown flap. This is a serious deficiency, since ground
effect will affect takeoff and landing performance as well as stability and control
characteristics. The effect of lift level and height above the ground must be
assessed. The significance of the ground effect on lift can be seen by looking at
jet flap data. Figure 48 indicates that as the jet flap approaches the ground, there
Is a substantial loss in lift, Reference 17. The figure also shows that care must
be tken to provido a realistic ground simulation since there is a substantial dif-
ference between fixed board and moving belt data.

Ground effect testing as it is normally done yields static data. Ground effect as
It is encountered in flight is a dynamic phenomenon. In order to provide realistic
inputs to airplane simulation, it would be advisable to be able to evaluate the
aerodynamic lag or h derivatives. This will be discussed further in Section4.3.1.

4.2.2.4 Flow Visualization. Understanding of force test results can be facili-

tated, insights necessary for theoretical development gained, and configuration
improvements defined from the use of flow visualization. This flow visualization
can range from simple tuft studies, to smoke, or the use of a separate model in
a water tunnel. The value of flow visualization is indicated in Reference 18
where smoke was used in understanding the destabilizing influence of a horizontal
tail.

88
Ir
A9
A0

.0

>0
0
CNo
01o
4a
4ez
w

.,c
co
Hrz

0 0

I
u-

A~~
9
Hn0
~ ~
___________
Very little flow visualization has been accomplished. The flow pattern of the
engine exhaust needs to be made visible to help determine the extent of the wing
influenced by the propulsion jet. This would also aid in optimizing engine-flap
arrangement. Boeing and NASA have used water injection into ejectors very
successfully to make visible exhaust jet location.

Water tunnel flow visualization requires a separate model. It is a powerful tech-

nique, however, which ONERA has used to understand power and vortex effects.
[1
Figure 25 shows a sketch made from a photograph of water tunnel visualization of
an externally blown flap in free air.

Flow visualization techniques complement a force test program and allow more
intelligent interpretation of the force data.

4.2.2.5 Pressure Data. Unlike the pure jet flap, the extent of wing influenced by
the propulsive jet of the externally blown flap is not known a priori. Wing and flap
static pressures would be needed to determine the extent and distribution of the
induced loading due to power. No pressure data is available in the open literature. I
The NASA has a limited amount of pressure data that will be available in a future [
technical note.

4.2.2.6 Parametric Force Testing. The available externally blown flap data
contains little parametric test data. In order to develop empirical or semi-
Ii
empirical estimation methods and to stimulate theoretical development, consistent,
high-quality, parametric test data is required. A summary of the force data
available with the deficiencies noted is given in Table VII.

4.2.3 Analytical Methods

The deficiencies in the methods developed in 4.1 will be defined. Also, the de-
ficiencies in jet flap theory on which the methods of 4.1 were developed will be LI
considered.

The approach and some of the approximations made in developing the engineering L_
method of this report were dictated by the limited range of data available and the
state of development of jet flap theory. LI
The method developed estimates incremental effects due to power. This assumes
that a good method is available for estimating the unpowered characteristics of an
unpowered wing. Another problem is that a part of the blowing goes into boundary
layer control before excess momentum is available to induce lift through the jet
flap effect. Thus, a wing that is badly separated unpowered may show much
larger increments due to power but lower overall levels than a wing that has mini-
mum separation. Using data generated from various sources for different pur-
poses, factors such as these cannot be determined.

90'
 144
'4
0
0.
bOcd
0 ~4 0 ~w~U. 0
o . . 0i Li
00 .A 'a

0'M0 0)1 0 00

> -

oi .,4 = I
rzo C)' oo 0 0 0d

BC .:(n

I Id

P4 0

0 0

-. 44

I - _91
 0) L4

-4

4C4
444 C)

E-4 4)o~ 1 Go ) t -l
4

0) 4 0 . 2 L4
C

jil 0L
V 01
>92 1i
 A 0

.- 0

411 4, lo ) lk~ (DS.

be o

a)d

'4

z ,4 ,

I a)a
aa

I 0

I 93
The method employs an effective jet deflection angle and static tuning efficiency.
How to ensure good angles and efficiencies was not determined. Obviously, such
L
parameters as engine orientation relative to the flap system, exhaust jet shape,
the relationship between flap chord and engine diameter, and engine bypass ratio[
may have an effect on both the turning losses and the angle to which the flap system
effectively turns the jet.

Insufficient data was available to develop a method to account for the effects of
flap span and blowing location on the aerodynamic characteristics. Fortunately,
the data indicated that the incremental effects on lift were only weakly dependent
on engine location. A force polar method for the two extremes of spread engines
B
and concentrated inboard loading only was developed. The pitching moments on
swept wings also depended on engine locations since the loading distribution is
important. Methcds more accurately treating flap span and engine location must
[I
depend on acquiring data giving detailed load distribution on the wing.

Methods for estimating lateral-directional characteristics were not considered.

These are extremely important and further work is indicated. Also, ground effect
must be considered.

4.3 RECOMMENDED PROGRAMS

An indication will be given of the program required to develop an estimation and

evaluation base for configurations using externally blown flaps. U
4.3.1 Wind Tunnel Test Programs

To avoid a proliferation of unrelated and uncorrelatable test data, a consistent,

parametric wind tunnel test program is required to eliminate the deficiencies
noted in Table VII. This would increase confidence in detailed design of an exter-
nally blown flap configuration. Any wind tunnel program should include, in addition
to basic free-air force data, flow visualization, flow field investigations, wing and
flap pressure distributions, and ground effect testing.

As discussed previously, ground effect as encountered by an aircraft is a dynamic

process. Aerodynamic lag may alleviate the adverse effects on lift at high lift
coefficients. It would be desirable to measure aerodynamic lag or (h/c) deriva-
Uves. This can be done in several ways. NASA has used the tow tank at Langley
with the model on a carriage moved over a ground board simulating several
approach angles. The Princeton Dynamic Track could also be used with the I
carriage programmed to simulate an approach and flare. A conventional wind
tunnel with a sting mounted model and moving ground belt and a data system [2
equipped to take r".namic data could also be used by placing the model at a given
angle of attack and traversing the model from free air toward the ground plane
while measuring a time history of the forces on the model. I
9U4
[1
I
1 4.3.2 Analytical Development

As additional test data becomes available the engineering method can be extended.
Since it relies on jet flap theory some additional jet flap theoretical work is in-
dicated. The effect of part span flaps, part span blowing, and high flap deflections
need to be theoretically evaluated. Theories to adequately estimate the flow fields
generated by Jet flapped wings must be evaluated.

IIn addition, tools for evaluating complete configurations using jet-flapped and
externally-blown-flapped wings need to be developed. The possibility of extending
existing general inviscid three-dimensional computer programs should be investi-
gated. These programs are potentially powerful tools to understand powered lift
configurations. Since the basic programs have already been developed, the inten-
tion would be to try to improve the simulation of engine operation and its inter-
jactions with the rest of the aircraft. The process will be difficult and is not
assured of success. The risk involved is high but the potential payoff is higher.

14.4 CONCLUSIONS

The feasibility of the externally blown flap concept has been determined from wind
tunnel testing. The testing has not, however, been intended to generate design
information so that there are a number of deficiencies in the data. These defi-
ciencies can be eliminated by a parametric wind tunnel test program. Concurrent
analytical development spurred by the accumulation of test data would improve the
present very limited design and evaluation confidence level.

I
I
I
I
I
I
I
 5. THE DEFLECTED SLIPSTREAM CONCEPT

5.1 INTRODUCTION

The feasibility and practicality of the deflected slipstream concept has been
clearly demonstrated by niddel test and by flight hardware both in flight test and
simulated operations. f
The development of the deflected slipstream type of configuration has been a
fairly gradual process based on the wing-propeller system of early aircraft.
The factors that have made such a concept suitable for STOL operation are the
development of the turboshaft engine and the high thrust levels available from
modern propellers. Additional refinements such as cross-shafting, boundary
layer control, and flaps with high turning effectiveness add to the possibilities
of the full exploitation of this concept. [
Large amounts of model test data have been obtained since the early nineteen
fifties. In recent years, interest in the tilt-wing concept has provided consider-
able amounts of data applicable to the deflected slipstream coefiguration. Most
of this data was obtained during tests designed for project evaluation or demon-
1
stration of the feasibility of the concept. Because oi the lack of a systematic
variation of parameters, it is difficult to apply much of this data to the develop-
ment of empirical methods (see, for example, Reference 19).

Flight test data have been obtained on a number of configurations. Of particular

interest are the Breguet 940 and 941. The latter first flew in 1961 and has since
been tested by the FAA and NASA and has been flown in simulated operations by
the commercial airlines. Other configurations that have been flight-tested in-
clude the Shin Meiwa STOL Seaplane, the Convair Charger, the Stroukoff YC-
134A and the Ryan 92 (VZ3-RY). The last named configuration, though, was a
V/STOL rather than a STOL aircraft and the emphasis on the VTOL performance
compromised the STOL considerations.

Tables VIII and IX show a brief summary of model test data and flight test data
respectively. The quoted reference numbers refer to the Bibliography, Volume II.

A number of analytical and empirical methods have been developed for the pre-
diction of the aerodynamic characteristics of deflected slipstream configurations.
None have been found, however, that can satisfactorily predict all the necessary
information required to assess a given configuration.

96

L
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5.2 REVIEW OF PREDICTION METHODS

5.2. 1 General Considerations

The available techniques for predicting the aerodynamic characteristics of de-

flected slipstream configurations fall between two extreme approaches.

One extreme is to represent the configuration by a purely mathematical model of

the flow processes involved and solve the resulting equations either analytically
or numerically. In principle, this leads to an evaluation of pressure and velocity
distributions everywhere and from that aspect is most desirable. However, a
comprehensive mathematical model of the real flow is difficult to solve analyti-
cally and requires large amounts of digital computer time to obtain numerical L
solutions. In addition, it is not clear that a purely potential flow model will
adequately represent the real fluid effects.

The other extreme is the empirical approach based on the evaluation of test data
and guided by consideration of the physical phenomena involved and by the simpler
theories. However, lack of systematic test data hampers the development of such
methods.

Most techniques combine the two approaches.

Traditionally, the approach to the evaluation of the characteristics of any aero-

dynamic configuration has been to first evaluate the characteristics of all the
major components in isolation. Then the effect of adding each additional compo-
nent to the major one (usually the wing) is computed until the whole configuration
has been "assembled. " In the case of propeller-wing aerodynamic characteris-
tics, it is necessary to add another term of major importance - representing the
mutual interaction between wing and propeller slipstream - including the induced
circulation lift on the wing due to the prop and the inflow distortion to the prop due
to the airplane structure.

In view of the above, the following contains a brief discussion of the characteris- U
tics of the isolated components of a deflected slipstream configuration, mainly
propellers and wing-flap systems and a discussion of the combined propeller-
wing-flap system.

The major components discussed are the propeller (including spinner, hub, blade
shanks and cuffs)and the wing-flap system.

Other components that affect the aerodynamics are the nacelles (including intake
and exhaust apertures and jet), the fuselage, the tail, etc. These items are
discussed below.

164
I
5.2.2 Propeller Methods

It is not intended to discuss the state of the art of propeller aerodynamic technol-
ogy but rather to indicate what slipstream predictions are available to the STOL
aerodynamicist who needs to evaluate the aerodynamic characteristics of a pro-
peller-wing-flap system.

The prediction of propeller aerodynamic characteristics has been under consid-

eration for a very long period with a considerable degree of success. In the, last
two decades the interest in VTOL and STOL applications has brought to light
shortcomings in the accepted prediction methods. Traditionally, the propeller
has been used in axial flow conditions and the effects of small deviations from
axial flow were easily estimated to a sufficient degree t_ accuracy. In STOL
applications, propellers experience very high disc loadings and can be subject to
large deviations from the axial flow condition. Under these conditions the pre-
diction of forces, moments and flow fields is less accurate than is necessary to
give reliable performance predictions for STOL configurations. Because of the
high disc loadings associated with deflected slipstream configurations, the flow in
the neighborhood of the propeller is subject to very high pressure and velocity
-l gradients.

The large deviations from axial flow that can occur due to, for example, gusts of
the same order of magnitude as the flight speed, give rise to in-plane forces and
moments that are no longer negligible. Even at fairly small angles of sideslip or
incidence, when the in-plane force may be small compared with the axial thrust
of the propeller, the induced velocity due to the in-plane force can cause a signi-
ficant change in the direction of local total velocity vectors in the slipstream.

In view of the importance of the flow field behavior of the propeller, the following
propeller related effects should be taken into consideration when analyzing a de-
flected slipstream configuration:

U(i) Thrust, side force and hub moment on the propeller in an originally
uniform, non-axial flow

(ii) Changes in thrust, side force and hub moment caused by changes in the
uniformity of flow due to

I(a) wing-flap system upwash field

i(b) nacelle-interference

(c) interference due to other airframe components

1 105

I
 (iii) Velocity and pressure distribution of the isolated propeller's
slipstream. This includes
i

(a) slipstream rotation, axial and radial velocities

(b) periodicity and unsteadiness of velocities and pressures

(iv) Interference effects on the slipstream caused by Items (ii) (a), (b)
and (c) above.

Various mathematical models of the propeller and slipstream have been proposed.

5. 2.2. 1 Actuator Disc. The actuator disc is a mathematical device replacing

the propeller that gives an impulse to the stream tube flowing through the propeller
disc. The simplest form assumes that in that streamtube there is no rotational
or radial velocity component anywhere in the field of flow and that the axial com-
ponent is uniform at any given axial distance from the propeller. With this
model the slipstream axial velocity can be related to the specified thrust of a
propeller of given size.

This model of the slipstream has been used in a number of prediction methods
and the resulting definition of slipstream dynamic pressure has been widely
adopted for the purpose of non-dimensionalizing test data. A ccnvenience stem-
ming from the use of slipstream dynamic pressure is the fact that coefficients
based on it do not approach infinity as flight speed approaches zero.

5.2.2.2 Inclined Actuator Disc. This approach as used i,1 Reference 20 Is the
same as the above except that the actuator disc affects only that component of the
freestream that is normal to the disc. Depending upon the choice of assumptions
about the nature of the flow, two slightly different values of slipstream dynamic
pressure and induced velocity are obteined.

5.2.2.3 Modified Actuator Disc. Glauert in "Aerodynamic Theory" (Ed.

Durand) outlines two modifications to the 'uniform' actuator disc, allowing radial
variations of axial velocity and introducing a rotational velocity term that is ob-
tained by relating the propeller torque to change of angular momentum of the
slipstream.

Reference 21 proposed two basic modifications:

A solid body rotation of the slipstream was assumed, the angular velocity being
a function of propeller speed and attitude and thrust and torque levels. The
angular velocity was expressed in terms of die propeller torque and thr ist coeffi-
cients and advance ratio.

106
I

Variations in the axial velocity of the sliestream are represented by construct-

ing the slipstream of a number of concentric annular streamtubes having differ-
ent values of the velocity potential and therefore axial velocity. The values of
some of the axial velocities must be specified and only a study of real slipstreams
will indicate their relative magnitudes.

Figure 45 of Reference 21 gives an indication of why such a process is necessary.

The figure shows the variation along a radius of the disc of the dynamic pressure
In the slipstream of a real propeller in the static thrust condition, at two differ-
ent blade settings.

5.2.2.4 Theodorsen's Wake Model. This was an attempt at simulating the real
flow in the wake while still retaining a potential flow model. The wake was rep-
resented by a helical sheet of vorticity shed from each propeller blade. The
sheets were assumed to the rigid and rotating with the blades. The flow field
-was obtained by solution of the potential flow equations.

Such a model -ave fairly good results for lightly loaded propellers. For highly
loaded propellers the distortion of the nearby portion of the wake is an important
factor in determining the flow conditions in the neighborhood of the propeller.

5.2.2.5 Vortex Wake Representation. Various models have been proposed that

f ]
use a discrete vortex representation of both propeller blades and the wake. Such
methods are amenable to numerical solution by digital computer.

i Two basic approaches are used, both of which represent the propeller blade with
a lifting line trailing a finite number of discrete vortex lines.

I Lup The first approach, called the free wake technique, is to allow the wake to take
the shape imposed by its own induced velocity field.

IThe second, called the fixed wake analysis, forces the wake to conform to a
shape predetermined by empirical methods.

The second approach is known to provide accurate predictions of propeller per-

formance in forward flight conditions but in the very low speed rang3 accuracy
decreases.

HNo attempts at the correlation of test data with a free vortex model have been
discovered during this study.

A refinement sometimes used in the above models is the avoidance of singulari-

ties by using vortices with viscous cores rather than potential line vortices.

The latter type of models offer the best promise of providing accurate predictions
of forces and flow fields since they have the scope of relating slipstream

j107
I
 °.

properties to blade geometry and of being expanded to include factors heretofore

omitted. Such factors as spinner and nacelle interference should probably be
included. One factor common to all of the above models is that the results given
are a time average of the flow conditions and not instantaneous values of the flow
conditions.

5.2.2.6 Empirical Methods. Numerous approximate methods are available for

the prediction of propeller forces and moments in uniform flow fields. The litera-
ture search has not revealed any empirical method for predicting the slipstream "
properties.

5.2.3 Wing-Flap System Methods

8.

In general, analytical methods for the prediction of aerodynamic characteris-

tics of wing-flap systems in uniform flow, are inadequate for all but "rough cut"
work. The usual technique for evaluating the aerodynamics of a wing-flap sys- &

tern is to evaluate the characteristics of the unflapped wing and then evaluate the
changes in the wing characteristics due to the addition )fthe particular flap con- --
figuration under consideration.

Frequently the first step and almost invariably the second are empirical ap-
proaches such as are outlined in USAF DATCOM or RAES Data Sheets. These
lead to evaluation of forces and moments and simple methods can be used to eval-
uate the downwash field at locations not too close to the wing-flap system. These
methods are not very accurate and are used only in the absence of test data on the
configuration under investigation or on a similar configuration. The most reli-
able procedure is to use test data from a configuration similar to or closely re-
lated to the configuration being studied. Empirical trends and theoretical pre-
dictions are used to account for small differences in configuration. However,
such a technique introduces the necessity of also taking consideration of differ-
ences between test and real conditions, wall effects, scale effects in addition
to differences of geometry.

The following wing-flap system methods can be used.

5.2.3.1 Horseshoe Vortex. This is the most rudimentary model and as such
is probably the least accurate. The horseshoe vortex model is capable of pre-
dictling downwash fields only near a streamwise line through the center of the span
of high aspect ratio wings.

5.2.3.2 Lifting Line. In this technique the wing is again represented by a line
vortex. But instead of the two trailing vortices of the horseshoe vortex model
there is a flat vortex sheet extending downstream from the lifting line. This .
model is better than the horsehsoe vortex but is still limited to high aspect ratio
wings. The flat wake assumption is a restriction to fairly low values of lift co-
efficient for at high lift coefficients the wing wake is highly curved so that the

108

!I
*! downwash field is not accurately modelled. Thus this approach is effectively
limited to wings of high aspect ratio and with no high lift devices.

15.2.3. 2 Lifting Surface. This approach represents the attempt to simulate the
downwash field in the vicinity of the wing. The technique uses a 'mesh' of horse-
shoe vortices located at the wing. The strength of the vortices is determined by
'tailoring' the local downwash velocity at each of a number of strategic points on
the wing surface to the slope of the wing surface there. If a sufficiently large
number of 'node points' is used presumably a good simulation of the wing-flap
could be obtained. Unfortunately, the amount of calculation required increases
with the number of node points specified so that usually a compromise between
I Iaccuracy and computation time must be made.

5. 2. 3.4 Vortex Lattice This approach is similar to the lifting surface techni-
que but is more ambitious and not limited to bodies of zero thickness. Here, the
vortex lattice is used to represent a complete surface (upper and lower surface
of a wing of finite thickness, for example) and local velocities are constrained to
follow the surface at node points.

Thie method has the prospect of being able to simulate almost any potential flow
but is subject to the same compromise between accuracy and computation time as
1 the lifting surface method.

For example, the vortex lattice simulation could be used for a slotted flap con-
figuration whereas the lifting surface theory would not be able to predict the
significant changes due to small variations in the slot configuration.

5.2.3.5 Empirical Methods. A variety of empirical methods is available -

mostly of a 'building-block' nature. The usual procedure is to calculate each of
[ the following items in turn, starting with the unflapped configuration:

J
* Airfoil lection characteristics:

lift curve slope

1angle of zero lift

maximum lift coefficient

aerodynamic center

1 profile drag, pitching moment coefficient at zero lift

1
1 109

I
 Wing planform characteristics:

airfoil characteristics subjected to the effects of finite

aspect ratio, taper, sweepback, twist; induced drag

Wing-flap characteristics:

wing planform characteristics subjected to the effects of flap

deflection, span and chord, extension, gap, etc.

5.2.4 Effects of Nacelles and Fuselage

Analytical methods for the effect of nacelles and bodies on wings exist and are
(in principle) reasonably simple to apply (slender body theory). However, empir-
ical charts based on wind tunnel data are readily available and simple to use,
employing only a few parameters such as wing incidence, body length and width.

Nacelles for turboshaft engines need special consideration because they lie in a
slipstream, and the location of intake and exhaust apertures has an influence on
the pitching moment of the system.

5.2.5 Combined Propeller-Wing-Flap System

The prectction of the characteristics of this system depends on the evaluation of

propeller charactxristics, wing-flap system characteristics, mutual interference,
and the influence of the propeller-wing-flap system on the rest of the airframe and
the flow field of the combination.

The approach can be made in two ways. First, the isolated characteristics could
be evaluated and then changes due to mutual interference between the two systems
could be calculated. Secondly, a mathematical model of the combination could be
postulated and used to evaluate its total characteristics.
5.2.5.1 Lift. The high lift of this combination can conveniently be viewed as
the sum of three major components, as illustrated:

Lift Lift induced by slipstream

Coefficient
.direct propeller force
component

wing lift in absence

of slipstream

Thrust Coefficient

110
 The following observations can be made about the three components:

Lift of the wing out of influence of the propeller has already been discussed in
. 5.2.3.

The direct propeller force component is not necessarily the same propeller force
as would be obtained by the propeller (or propellers) in isolation at the specified
attitude and freestream condition. A major factor influencing the propeller is the
upwash field ahead of the wing due to the high value of the circulation at the high
-. lift condition. This means that in general there will be a non-axial and non-
uniform flow incident at the propeller. This in turn introduces a propeller in-
plane force. The net direct propeller force then is composed of both thrust and
in-plane components. Some authors include in the 'direct thrust' term the lift
due to the turning of the propeller slipstream by the wing-flap system. This is
the same as assuming that the thrust vector is rotated by the action of the wing-
flap system, and the amount of rotation (called "static turning effectiveness") is
usually an empirical value obtained from tests at zero forward speed. The
assumption that the turning effectiveness at forward speed is the same as at static
conditions is suspect. The drag of the wing in the slipstream produces a loss of
momentum of the slipstream so that the effect of turning the slipstream is less
• than might be predicted by an amount depending on the deflection. Because of
the upwash field of the wing it is seen that the direct propeller force component
of lift depends on the wing-flap system to some extent even if the lift gained by
turning the slipstream is not included as part of the direct force term.

The lift induced by the slipstream is due to a number of different effects, the
main ones being

(a) increased dynamic pressure in the slipstream.

(b) change of local angle of attack of that portion of the wing within the
slipstream. This is caused by f~e increased mean velocity vector,
the spanwise variation in downwash resulting from rotation of the
propeller slipstream and the spanwise variation of axial induced
velocity resulting from non-uniform blade loading, and

(o) the deflection of the propeller slipstream by the wing-flap system.

3Other factors that can affect the lift induced by the slipstream include

(d unsteady and/or periodic fluctuations of velocity components in the

I slipstream. Test data indicates that this effect is negligible except
perhaps at high incidence. it is a well known phenomenon that cyclic
pitching or heaving motions of an airfoil result In changes of maxi-
I mum CL, and that the magnitude of the change deperds upon the fre-
quency of the motion. This may explain the difference between the

II
 lift curves obtained without propellers and those obtained with propellers
producing zero net thrust.

(e) spanwise spreading (and contraction) of the slipstream effect due to

spanwise pressure variations near the edges of the immersed portions
of the wing,

(f) increase of local Reynolds Number and Mach Number within the
slipstream,

(g) effects of regions of velocity shear, and

(h) entrainment of the freestream by the slipstream.

5.2. 5.2 Drag. The drag can also be thought of as being composed of three
major components.

Drag of the airframe in the absence of the slipstream can be calculated by stand-
ard methods as discussed in 5. 2. 3.

The direct propeller component of drag is subject to the same remarks as the
propeller component of lift.

The slipstream induced drag is subject to the same considerations as the corre-
sponding lift term. In addition, it should be noted that the spanwise variation of
lift directly affects the drag; due to the large spanwise variations in lift, relatively
high values of induced drag are likely to occur.

In addition to the above, there are significant contributions to the net drag from
the ram drag and jet efflux of the nacelles that are normally placed behind the
propellers in the slipstream.
5.2. 5.3 Pitching Moment. Pitching moment of the propeller-wing-flap system
can be analyzed into three components. They are:

The isolated wing-flap pitching moment.

The direct propeller term. This includes moments arising from thrust,
in-plane force and hub moment. Also, the ram drag and jet efflux of the
engines will cause moments of magnitude dependent on the location of
nacelles with respect to the reference point.

The induced moment due to the effect of the slipstreams on the wing. It
is to be expected (and is shown by test data to be so) that a large nose down
pitching moment is the penalty paid for turning the slipstreams through
large downward angles by means of the flaps. The necessary download

112
I
required at the tail to trim these moments may considerably reduce the gain
in lift obtained by turning the slipstream.

5.2. 5.4 Downwash. Because of the complex flow resulting from the interaction
of the propeller slipstreams with the wing-flap system and the high deflection of
the airflow by the flaps the downwash field cannot be accurately calculated by the
simple horseshoe vortex or flat vortex sheet models that are suitable for low
-- lift levels.

5. 2. 5. 5 Lateral-Directional Characteristics. The aspects of lateral-directional

aerodynamics that are of primary interest are control response, sideslip and yaw
characteristics and engine (or propeller) failure. It is probable that the forces
resulting from lateral control deflections and those due to engine failure could be
evaluated by consideration of the local changes of lift, drag, and moments on the
wing. The sideslip and yaw dependent forces will be due in part in the in-plane
forces and hub moments of the propellers; other components will be as for the
isolated airframe unless the sideslip is so great that propeller slipstreams will
impinge upon the fuselage or vertical tail.

5. 2. 5. 6 Stability and Control. The necessity of trimming the large nose down
pitching moments due to the use of flaps has already been noted. Calculation of
the required restoring moment provided by the horizontal tail depends upon a
detailed knowledge of the local dynamic pressure and the downwash in the vicinity
of the tail. The changes in trim due to flap deflection and due to power variation
can be large because of the fact that the tail lies close to and sometimes within
- -the propeller slipstreams and therefore is subject to large and rapid variations of
downwash and dynamic pressure.

The tail provides the nose up restoring moment by means of a download which
reduces the total lift of the aircraft significantly at high flap deflections. Predic-
tion of this effect is impossible without a thorough knowledge of the details of the
flow field in the vicinity of the tail and the variations in the flow field due to geo-
metric and power changes.

No special methods for predicting dynamic stability have been discovered relating to
deflected slipstream configurations and inthe STOL operation doubt has been cast on
classical dynamic stability analysis. Because of the low flight speeds involved
in STOL operations the assumptions of small disturbances and linear equations
are possibly not valid and may even be misleading. Also, since the takeoff and
landing operations are acceleration maneuvers, the assumption of an initial steady
state may also lead to erroneous stability predictions. Time lag terms that re-
sult from the finite time required for slipstream to the tail are of importance
since the moment provided by the tail may be very sensitive to changes of slip-
-* stream properties and location.

113
5.3 A DISCUSSION OF SELECTED METHODS FOR THE DEFLECTED

SLIPSTREAM CONCEPT

5. 3. 1 General

Section 1. 2 of the Bibliography (Volume II of this report) contains 38 references

that include prediction methods or that include discussions of the problems asso-
ciated with the development of methods.

Of these references that include prediction methods, a representative sample has

been reviewed below. The methods considered include both the purely analytical
and empirical types and the table below indicates the nature of each of the re-
viewed methods.

The me-'ods listed in Table X are reviewed in detail in the order shown.

TABLE X
PREDICTION METHODS FOR THE DEFLECTED SLIPSTREAM CONCEPT

Reference Description

"A Preliminary Theoretical Inves- Combines actuator disc represen-

tigation of the Effects of Propeller tation of prop with lifting line or
Slipstream on Wing Lift", Douglas slender wing theories. Also dis-
Rep. SM 14991, E. W. Graham cusses lifting surface theory.
et al (Reference 22)

Series of reports on wings extend- Combine uniform jet with

ing through slipstreams by Vehicle Weissinger lifting surface theory
Research Corporation V. R. C. Re- to consider lift of wings with
ports 1, 8, 9, 9a, 10. and multiple jets, inclined jets, at
"Aerodynamics of non-uniform flows high angles of attack and in
as related to an airfoil extending separated flow conditions. Only
through a circular jet". J. Aero. theory is included, no comparison
Sci. Vol. 25 No. 1 S., Rethorst with data except in Reference 39.
(References 23 through 27 and 39)

"Semiempirical procedure for Uses actuator disc theory to cal-

estimating lift and drag charac- culate slipstream properties.
teristics of propeller-wing. flap Evaluates lift and drag by estima-
configurations for vertical-and- ting change of momentum of prop, -
short-take-off-and-landing air- slipstream and 'wing stream-tube'.
planes". NACA Memo 1-16-59L Not limited to unflapped wings.
R.E. Kuhn (Reference 28)

114
 I
TABLE X - Concluded

Reference Description

"A stability analysis of tilt-wing Inclined actuator disc theory is

aircraft (analytical)". Princeton used to calculate local flow condi-
U. Rep. 177, C. H. Cromwell and tions at wing. Mass flow correc-
H. E. Payne (Reference 29). tion used to account for the fact
that only part of the wing may be
inside the slipstream.

"Effects of propeller slipstream Uses inclined actuator disc

on V/STOL aircraft performance theory and slender wing theory.
and stability". TRECOM TR Includes consideration of flapped
64-47, Dynasciences Corp. wings.
(Reference 20)

"An investigation of propeller Extends Reference 20 to predict on-

slipstream effects on V/STOL set of stall, including slipstream
aircraft performance and stability", rotation and uneven axial flow.
USAAVLABS TR 65-81 L.
Butler et al. (Reference 21)

"Lifting surface theory for V/STOL Combines an inclined actuator

aircraft in transition and cruise", disc with a wing represented by
USAAVLABS TR 68-67 E. Levinsky a lifting surface.
et al. (Reference 30)

5. 3. 2 Discussion of Speoific Methods

5. 3. 2. 1 A Preliminary Theoretical investigation of the Effects of Propeller

Slipstream on Wing Lift (Reference 22). This report discusses three different
methods for calculating the increase of lift on a wing due to interaction between
it and a propeller slipstream. In each case the propeller slipstream is repre-

I
[ ..
-
sented by a uniform, straight tube of fluid of specified velocity. The three
methods differ in their represent,-ion of the wing.

-_ First, the lifting line model as originally proposed by Koning (Reference 31) is
used and has been solved by solution of Laplace's equation in the Treffz plane and
the adoption of a Fourrier sine series for the circulation on the parts of the wing
in-and outside the slipstream. The boundary conditions at the edge of the slip-
-stream are satisfied L the adoption of an "image" horseshoe vortex system of
the appropriate strength located at the inverse points corresponding with the
horseshoe vortex system representing the wing. It is stated that the aspect ratio
of the portion of the wing in the slipstream, D/C, (ratio of slipstream diameter

1. 115

1
to wing chord) is an important parameter and that the lifting line approach is of
doubtful value when D/C becomes small (less than 6, say).

The second approach was to employ the slender wing theory of R. T. Jones (Ref-
erence 32). By assuming a separated solution for Laplace's equation in terms of
a perturbation potential in polar coordinates at the trailing edge of the wing a
closed form solution is obtained for the lift increment at any spanwise station.

It is acknowledged by the authors that the slender wing approach is likely to be ""
valid only for small values of D/C (less than about 1. 0, say) and so a third ap- ,.
proach is taken in order to bridge the gap between large and small values of D/C.

The third approach was the application of Weissinger's lifting surface theory
(Reference 33). The only case for which a solution was obtained for this model
was that of an infinite span wing with a sinusoidally varying angle of attack span-
ning a slipstream of finite width and infinite height.

A comparison is made between the three methods for the case of a wing in a
slipstream of infinite height and is shown in Figure 49. Also a comparison with
some test data from Reference 34 is made using the lifting line and slender wing
theories in Figure 50.

Figure 49 indicates, as might be expected, that the Weissinger lifting surface

approach approximates the slender wing theory at low values of D/C and the
lifting line theory at high values of D/C..

Figure 50 illustrates the relative merits of the lifting line and slender wing ap-
proacies in predicting the spanwise variation of lift on a wing spanning a slip- ,.
stream with constant velocity distribution and no rotation. It is seen that the
slender wing approach is more accurate than the lifting line, at least for the por-
tion of the wing that lies in the slipstream. Some doubt is cast on the accuracy
of this test data because of the apparent influence of the jet on the wing at large
distances from the center of the jet.

The report also includes a comparison of predictions using the slender wing
method with test data from Reference 35.

116
T
IThe comparison shows poor agreement between predicted and test values of lift
increment due to slipstream.
Ck S ender wing 1.4 Tet Dt a

27a T ory Ck0 Ving alone

0.8 - 1. - kith j=t

Lifting Line
0.6 Theory 1 .0

0.4 0.8

0.2 Lifting 0.6

Surface Theory lifting line
S i -- slerder ing
L 2 3 4
J Distance fro.' Jet
Jet Width/Wing "hord Axis in Radii
FIGURE 49. COMPARISON OF LIFTING FIGURE 50. COMPARISON OF LIFTING
LINE, SLENDER BODY AND LINE AND SLENDER WING
LIFTING SURFACE METHODS PREDICTIONS OF REFERENCE
OF REFERENCE 22 FOR A 22 WITH TEST DATA FROM
WING WITH SINUSOIDALLY REFERENCE 34 FOR A WING
VARYING ANGLE OF ATTACK EXTENDING ACROSS A CIR-
IN AN INFINITE PLANE- CULAR JET
SIDED JET

I ti
The methods of this report contain the following general assumptions:

The froestream and slipstream are composed of an ideal, incompressible, non-

viscous fluid;

I The propeller is an actuator disc producing a uniform pressure discontinuity

across its surface and introducing no rotation to the slipstream;

I The axis of the slipstream inte ,sects the wing;

Interference effects of wingtips, fuselage, nacelles, etc. are Ignored;

The boundary conditions of equal pressure and zero normal velocity at the edge
I of the slipstream must be satisfied.

I 117

!I
In addition, the following assumptions apply to the lifting line model:

The wing is represented by a straight vortex normal to the axis of the slipstream;

The trailing vortex system is represented by a series of 'elementary' horseshoe

vortices (or vortex doublets).
The two following assumptions apply to the slender wing and lifting surface
models respectively:

Flow at the trailing edge of the wing is tangential to the wing;

The wing is replaced by straight lifting line located at the quarter chord of the
wing and the flow is forced to match the local airfoil slope at the three-quarter
chord of the, wing.

Slipstream rotation does not greatly affect the overall lift except when the wing
is partly stalled as is seen in References 36 and 37. However, slipstream rota-
tion does affect the local spanwise lift variation (Reference 38). This effect
should be included so that the onset of stall can be predicted.

5. 3. 2. 2 Vehicle Research Corporation. A series of reports published by the

Vehicle Research Corporation (References 23 through 27) presents the develop-
ment of a lifting surface theory for wings extending through propeller slipstreams.
Special aspects of the aerodynamics of the deflected slipstream concept that are
considered include:

Wings extending through multiple jets, (Reference 24);

Wings at high angle of attack extending through multiple jets,

(Reference 25);

Wings at high angle of attack extending through inclined jets,

(Reference 26);

Wings extending through multiple jets in separated flow conditions,

(Reference 27);

Highly cambered wings "as used in deflected slipstream V/STOL

arrangements" (Reference 25).

The work described in References 23 through 27 was the result of a program de-
signed to generalize and extend the basic lifting surface theory of Reference 39.

The basic lifting surface C eory of Reference 39 is an application of Weissinger's

lifting surface theory. The representation of the propeller slipstream in the

118
I
analysis, as in the analysis of Reference 22, is a circular jet of uniform flow
properties and no rotation. The approach to satisfying the boundary conditions
at the edge of the slipstream differs, however. The wing is assumed to consist
of "even" and "odd" parts of horseshoe vortices with their bound vortices at the
wing quarter chord. The even parts of the system are pairs of parallel vortices
of strength r/2 extending to infinity both upstream and downstream of the bound
vortex. The odd parts consist of a conventional horseshoe vortex of strength r,'2
extending to infinity downstream and another of strength -F/2 extending to infinity
1upstream. The effect of the slipstream on tie wing is represented by a pertur-
bation velocity potential, 0 , that is expressed as either of a pair of infinite series
in Bessel functions depending on whether points are being considered inside or
outside of the jet. The value of 0 is obtained by imposing the boundary conditions
of constant pressure and tangential flow at the jet boundary and the downwash
condition at the three-quarter chord line of the wing.

Predictions of spanwise lift distribution obtained by the use of the method of Ref-
erence 39 show good agreement with the test data of Figure 50 for that part of
the wing lying in the slipstream.
The special aspects of the deflected slipstream concept noted above are all

treated by modifications of the method of Reference 39.

II High angles of attack are treated in Reference 26 by satisfying the Weissinger

downwash condition at the three-quarter chord points of the wing rather than at
the corresponding poirts in the horizontal plane through the bound vortex.

I! Multiple jet effects are accounted for in Reference 24 by the application of iniage
systems for each jet as before with the addition of further image systems related
to the original image systems. Logically this lends to an infinite set of image
I vortices but to simplify the problem the author of Reference 25 used only the
basic vortices, their images as required for each jet ant the images of those
gimages as required for each jet.

Wings extendin, ,hrough inclined jets (Reference 26) are treated by applying the
Weissinger downwash condition to an "effective" local airfoil slope at the three-
quarter chord point. The effective slope is simply the difference between the air-
foil local slope and the jet slope inside the jet and the difference between the local
slope and a modified jet slope outside the jet. The jet slope is calculated by an
approximate inclined actuator disc theory.

Separated flow conditions are allowed for in Reference 27 by placing the bound
vortex of the separated portion of the wing at the one-third chord location and im-
posing the Weissinger downwash requirement at the mid chord porttion or by
employir actual two dimensional airfoil characteristics in separated flow
conditionb.

1 119

I
In some of the above work a term is included to describe rotation in the propeller
slipstream.

All of the work presented in References 23 through 27 and 39 requires a large

amount of computation to evaluate the lift distribution.

References 23 through 27 do not contain any numerical calculations or compari-

sons between prediction and test data.

5.3.2.3 Semiempirical Procedurc ior EstimatiNg Lift and Drag Characteristics

of Propeller-Wing- Flap Configurations for Vertical-and-Short-Take-off-and-
Landing Airplanes. The approach taken in Reference 28 is different from those ""
considered above in that no attempt is made to model the details of the airflow.
Only the gross effects of wing-slipstream interaction are considered.

The propeller slipstreams comprise one body of air and their change of momentum
is the result of the increment of velocity due to passing through the propellers and
turning downwards by the action of the wing and flap.

The effectiveness of the flaps in redirecting the propeller slipstream is assumed

to be the same as that achieved in static tests. Reference 28 includes correlations
of the flap effectiveness for a number of different configurations, together with.•
data indicating the loss of momentum due to the turning action of the flaps.

The second body of air considered is that flowing through a circular stream tulke
of diameter equal to the wing span; a correction term is included to allow for the
fact that the propeller lipstreams are inside this larger one.

Simple actuator disc theory is used to relate slipstream velocity to propeller

thrust.

The lift and drag of the propeller-wing-flap combination are evaluated by computa-
tin of the gain of downward and forwnrd momentum of the combined stream tubes.
The expressions developed for the lift and drag coefficients contain three distinct
terms:

(a) the "power-off" coefficient

(b) a term representing the rotation and modification of the propeller

thrust vector and

(c) a so called "augmentation" term

A correction factor of 1.6 is applied to the augmentation term as a result of com-

paring predictions with several sets of test data.

120

i
 Figure 51 shows a comparison of predictions obtained using this method with
test data from Reference 40.

C LO8 =30' -
2SF
Props =-C))8

6 6

4 4

2 == 2 15 °

0 0
0 . 10 10
Th zust Coaff. T/qS Thrsz CoeLf. T/.5
FIGURE 51. COMPARISON OF PREDICTIONS OBTAINED USING
THE METHOD OF REFERENCE 28 WITH DATA FROM
REFERENCE 40

This method has the merit that by the judicious use of correction factors it could
be adapted to fit a related family of configuration.

The method uses the "power-ofP characteristics of the configuration as a basis

for calculation of the slipstream effects. This indicates that such considerations
as the effect of aspect ratio are automatically accounted for a least in part.

The possibility exists of extending the method to include, for example, the effects
of propeller slipstream rotation, without the need for lengthy analysis or
computatior.

Further comparisons are made in Section 5.4 of test data with predictions made
using this technique.

5. 3.2.4 A Stability Analysis of Tilt Wing Aircraft (Analytical). Although the

prediction method for lift and drag outlined in Reference 29 is intended for tilt
wings it is also applicable to deflected slipstream configurations. Also, it rep
resents a third type of approach to the problem of assessing wing-slipstream In-
teractions which is different from the purely analytical approach of References
22, 23 through 27, 39, and from the momentum approach of Reference 28.

First, an approximate inclined actuator disc theory is used to relate the velocity
and inclination of the propeller slipstreams. The angle between the wing chord-

II
line and the slipstream velocity vector is chosen as an effective angle of attack for
the wing and the assumption is made that the wing possesses the same lift and

121

!
drag coefficients as it would in a free stream of speed equal to the slipstream
velocity and placed at the effective angle of attv'k. It is recognized that only part
of the wing may be immersed in the slipstreams and so a mass flow correction
factor is calculated, which is the ratio of the "actual" mass flow to the "assumed"
mass flow. The "assumed" mass flow is the product of the velocity through the
propeller disc and the momentum area of the wing. The "actual" mass flow is .i
the sum of the mass flow through the propeller disc and the mass flow through the
momentum area of the wing (which iz adjusted to allow for the presence of the
slipstreams within it.)

Figure 52 shows predictions of the lift and drag coefficients (referred to slip-
stream dynamic pressure) as a function of "effective" angle of attack for the
model test of Reference 29. Correlation appears to be poor especially at high
values of effective angle of attack.
0 "
L .6 0

0.5 C6-

0.4

0.3 03 3
1.2

0.11 23

0 10 20 30 40
Effective Incidence, al

FIGURE 52. COMPARISON OF PREDICTION BY METHOD OF

REFERENCE 29 WITH TEST DATA FROM THE SAME
SOURCE
The method, in effect, assumes that the whole of the wing isat the effective angle
of attack experienced by that part of the wing that is in the slipstream.

The inclined actuator disc theory employed is good only for small angles of attack.

The report continues, having evaluated lift and drag, to study the stability and
control of tilt wing ai- 'raft. The equations of motion are formulated and the con-
ventional assumptions of uncoupled longitudinal and lateral motion and small per-
turbations are made. The stability deiivatives are defined and evaluated for a
specified configuration. Derivatives that could not be evaluated were assumed
negligible. The response of the configuration to various perturbations is shown
in tht report.

122
 5.3.2.5 Effects of Propeller Slipstream on V/STOL Aircraft Performance and
Stability (Reference 20). The lift and drag of a defduted slipstream (or tili wing)
airplane are calculated and expressed in simple frrm.

The increment of lift on the part of the wing immersed in a slipstream is calcula-
ted using the slender wing theory of Reference 32. Other contributions to lift and
drag calculated are, propeller thrust and normal force and the "free stream"
lift and drag of the wing.

Inclined actuator disc theory is employed to calculate the inclin!A.on and velocity
of the slipstream at the wing.

Figure 53 shows a comparison of predictions with test data from Reference 40.
Note that the coefficients are referred to slipstream dynamic pressure, CL,
L/q+T/A) etc.

1.6
C
LS,
Cxs 1.4 -

CLS
1.2

1.0 /
0 o CXs
0.2

AngX: of attack.. uegrees

~FIGURE 53. COMPARISON OF PREDICTIONS USING METHOD OF
-- REFERENCE 20 WITH TEST DATA FROM REFERENCE
36

0123

0
 i
Comparisons are included of test data from Reference 36 and calculations pre-
dicting lift and drag for the model test of Reference 36. Generally, correlation
appears to be good at low angles of attack and provided flaps are not deflected.

The effect of flap deflection has been assumed to be equivalent to a change of angle
of attack of the same magnitude as the flap deflection.

5.3.2.6 An Investigation of Propeller Slipstream Effects on V/STOL Aircraft

Performance and Stability (Reference 21). The prediction method of Reference
21 is based o, that of Reference 20. The lift increment due to the slipstream in-
teracting with the wing is calculated as before by the slender wing theory of Ref-
erence 32. Reference 21 differs in the evaluation of the local angle of attack in
the slipstream by allowing for a rotational term.

Using Schrenk's spanwise loading approximation the prediction of the onset of stall
is demonstrated. Figure 54 illustrates predictions of stall angle of attack in com-
parison with test data from Reference 41. Onset of stall was considered to occur
at the angle for which the lift curve first became non-linear.
50

Angle o0 -

Attack at
Onset of
Stalldeg 30

20

10

0
0 2 .4 .6 .8 1.0
Thrust Coefficient CTS

FIGURE 54. PREDICTION OF ONSET OF STALL

Two further modifications are made to the basic method of Reference 20. First :
the effect of flap deflection on the wing effective angle of attack is modified by a
factor depending on the flap deflection and based on test data comparisons. This
resulted in improved predictions for large flap deflections but deteriorated the
predictions for flap deflections of less than 200.

Secondly, a scheme for taking into account radial variations in the axial velocity
in the slipstream was indicated. A sample case was calculated using this scheme

124
1

and it was shown that if the peak velocity in a propeller slipstream is close to the
axis the gain in lift due to wing-slipstream interaction will be higher than if the
peak velocity is near the outside of the slipstream or if the slipstream is uniform.

The rotational term included was in the form of a solid body rotation whereas a
vortex type of rotation may be more applicable.

5.3.2.7 Lifting Surface Theory for V/STOL Aircraft in Transition and Cruise
(Reference 30). The prediction method of Reference 30 is developed in two parts.

First, an inclined actuator disc theory is developed in incompressible flow, pro-

ducing a circular slipstream that is simulated by a set of ring vortices coaxial
with the slipstream and by vortices, sources and sinks distributed along the length
of the slipstream. The usual boundary conditions are imposed to evaluate the
velocity field of the slipstream. Representation of the propeller by a set of bound
vortices with a trailing vortex extending helcally downstream from the tip and
another extending axially from the blade root allows of the introduction of rota-
tional velocity terms inside the slipstream. The authors illustrate that the actua-
tor disc theory can predict the downwash at the centerline of the slipstream fairly
well when compared with test data showing average downwash in the vicinity of a
tail.

The second part of the development of this prediction method is the simulation of
the wing-slipstream combination. The wing is represented by a system of dis-
crete small horseshioe vortices with the bound vortex elements on the quarter
chord line of the wing.

A system of horseshoe vortex elements is distributed around the slipstream in

order to satisfy the "equal pressure" boundary condition. The tangential velocity
boundary condition is satisfied by the assumption of a 'reduced potential' inside
the slipstream. The Weissinger downwash condition at the three quarter chord
int is imposed. Satisfaction of tht. boundary conditions leads to set of linear
'I sultaneous equations whose unknowns are the strengths of the wing and slip-
strc am vortices.

Several sample calculations were carried out and comparison with test data from
Reference 42 shows fairly good correlation with measurements of spanwise lift
variation at small angles and attack. Downwash measurements, however, are
considerably larger than the method predicts. An improvement in the prediction
was achieved by assuming the slipstream had been deflected by only a half of the
value calculated from the inclined actuator disc theory.

All trailing vortices (from the wing and from the propeller) arc inclined at the
3 same angle to the freestream direction.

125

1.
The propeller slipstream starts at the wing quarter chord location.

The actuator disc theory may be used to predict flow fields only at large distances
from the actuator disc.

5.3.3 Summary Comments

The prediction methods reviewed above are a sample of the available methods.
Those available concentrate on the problem of the interaction between the pro-
peller slipstream and the wing.

The mathematical methods consisting of distributions of vorticity 'tailored' to fit

certain boundary conditions appear to have employed models that are oversimpli-
fied particularly with regard to flapped-wings. The fact that all such methods
employ an actuator disc to represent the propeller and ignore the effect of the wing
on the propeller performance sheds doubt on their validity.

The early empirical or semi-empirical methods (Reference 28) were somewhat

oversimplified but later methods have indicated that the semi-empirical approach
has several advantages:

They can easily be forced to fit the data from tests of a family of configura-
tions in order to predict the change in characteristics caused by alteration
of important parameters;
They are easily extended to include the effects of phenomena that were -.

originally omitted;

They require a minimum of computation time;

The more complicated analytical methods have not been shown superior in
their ability to predict lift.

It is to be noticed that of the methods available only lift, drag and the induced velo-
cities due to the wing-propeller slipstream interaction are evaluated, and these
only for the symmetrical flight condition.

Reference 21 indicates a method for the pr-,diction of wing pitching moment by de-
monstrating that the center of pressure of the wing, when expressed as a fraction
of extended chord, does not shift when propeller thrust coefficient is changed ex-
cept near the static thrust condition (CTs > 0. 8). Reference 21 also includes a
survey of data related to propeller normal force and hub moment including the
changes in these items due to the presence of a wing. This aspect is mentioned
here because in most of the methods reviewed above these items, although of im-
portance, have been ignored.

126
 In all of the above, the characteristics of real propellers have not been taken into
consideration. Various methods exist for the prediction of propeller loads in axial
and inclined flow conditions. Again, there are two basic approaches, analytical
and empirical. In the past, analytical methods have been of reasonable accuracy
for axial flow conditions at low thrust coefficients but in the near static case and
in inclined flow conditions these methods are inadequate. .ent analytic at-
tempts, Reference 43 for example, taking advantage of the large capacity digital
computational facilities now available have been encouraging. The approach taken
has been to represent each propeller blade as a bound vortex with a system of
trailing vortices. The trailing vortices are either constrained Reference 43 to
fit in a wake of given shape determined empirically (as in Reference 43) or allowed
to follow the motion imposed upon them by the induced velocity field of the whole
system. Such a technique is necessary in the high thrust conditions achieved in

tusuallySTOL operations because of the large distortions of the trailing vortices from the
assumed regular helical form.

Theoretical predictions of the stability and control of deflected slipstream con-

figurations are, in principle, easily obtained by traditional methods if the prob-
lem of defining the forces and moments on the airplane in all relevant flight condi-
., tions has been solved. This area of study has been hampered by the lack of
knowledge about the trajectory of propeller slipstreams and the induced flow fields
"
*'about the aircraft, particularly in the vicinity of the tail.

The validity of classical stability methods applied to V/STOL aircraft has been
questioned (Reference 44) for variety of reasons. Aspects upon which doubt has
been cast include: the assumption of small perturbations, the validity of lineari-
zation of the equations of motion and the assumption of an equilibrium steady state.
Evid.nce of non-linearities -in the longitudinal characteristics of tilt wing airplanes
has been reported in Reference 19.

No methods have been found that predict ground effects for deflected slipstream
configurations. The best approach to date has been to employ test data (for
example, Reference 45) or use the general method of Heyson (Reference 46).

5.4 CORRELATION OF TEST DATA WITH PREDICTION

The method of Reference 28 has been used to predict ligt and drag coefficients for
two deflected slipstream configurations. In the first case the method is compared
with flight test data for the Breguet 941 (Reference 47). Secondly, the method is
compared with wind tunnel data from NASA TND4448 (Reference 48).

1
~127

I
5.4. 1 Prediction Method of Reference 28

The equations developed for lift and drag coefficients by the method of
Reference 28 are

CL L Tsin (a + I+T (S/A (33)

and

CD =Dow +(IT1 Va1 ) - cos 0)] (34)

where:

CLow, C Dow are "power-off" lift and drag coefficients at angle of

W attack aw

F is thrust recovery factor (Figure 3 of Reference 28)

T

T is thrust coefficient, T/qS

c
q is freestream dynamic pressure

S is wing area

au. is wing angle of attack relative to the freestream direction

0 is flap turning effectiveness (Figure 2 of Reference 28).

A is totalpropeller disc area

K is an empirical facto used to make the above equations

fit test data used in the formulation of the method. A
value of K-eI. 6 was recommended by the author f
Reference 28 and will be used for the predictions
calculated below.

5.4.2 Flight Test Data and Calculations

The data used for this comparison was obtained from Reference 47, Figure 29(a).
Reference 47 is a report of flight test carried out by NASA on the Breguet 941
deflected slipstream airplane. Figure 29(a) contains the flight test lift-drag polar

128
 for the aircraft in the take-off configuration with settings of 450 on the inboard
flaps and 300 on the outboard flaps.

I i Since the method of Reference 28 predicts the lift and drag of the propeller-wing-
flap system the basic flight test data was adjusted to allow for the tail contribu-
tions required to trim the airplane. The tail characteristics and the downwash at
the tail were estimated using the methods of Appendix B of Reference 49.

The basic data used for the predictions were the lift and drag corresponding to
Tc = 0 in Figure 29(a) of Reference 47.

-u The values of CLow, C Dow were obtained from the trimmed values of CL, C D
at T' = 0 using the following expressions:
ST
'9
CLow
C LCL LT S
(35)

S
CDow CD DT (36)

=a (aF - e -a )+a (7
CLT 1T T - T 2T e (37)

2
-. C DT e
=C LT sin i+ C LT nR TT(38) (8

where

"" ST is tail area

obtained from
a1T is tail lift curve slope Appendix B
( of Reference 49
a2 T is elevator effectiveness

e is downwash at the tail

aF is fuselage angle of attack

iT is horizontal tail setting

a oT is zero lift angle of attack of tail

1 6 is elevator deflection

CLT is tail lift coefficient required for trim

1 129

in
 CDT is tail drag coefficient required for trim

ART is tail aspect TJ

ratio

T
T is Oswald efficiency of tail

A sample calculation for this set of test data is shown below for the case of
5 atT =1.0.

The following data are required for the calculation:

Wing Area, S 889 ft2

Wing Incidence, w 30 relative to fuselage reference

line

Wing MAC, F 12.15 ft

Wing span, b 76. 1 ft

Wing chord at of inboard prop, c 13. 2 ft

Wing chord at £ of outboard prop, c 10. 1 ft

Extended flap chord at 300, cf 0.42 local chord

Extended flap chord at 450, cf 0.43 local chord

Spanwise extent of inboard flap 0. 56 semispan

Spanwise extent of outboard flap 0.44 semispan

Propeller diameter, D 14. 76'ft

Propeller total disc area, A 684 ft 2

Tail area, ST 320 ft 2

Tail aspect ratio, ART TL3.3

1T 5.60
Tail setting,
0

130

L
 The following values will be required for the calculation of "power-on" lift and
drag coefficients:

(cf/D) i = 0.43 x 13. 2/14. 76 = 0. 385 inboard

(cf/D)o = 0.42 x 10. 1/14.76 = 0.287 outboard

Figure 2 of Reference 28 gives corresponding values of flap turning effectiveness

of:

(8/S)i = 0.65

(0 )=0.55

and for inboard and outboard flap deflections of 450 and 300 respectively the flap
turning anglt6 are ubtained

oi = 29.2*

Oo = 16.5

A mean of these two values, weighted in proportion to the spanwise extent of the
respective flaps gives:

0 = 0.56 x 29.20 -0.44 x 16.5 = 23.60

Corresponding with this value ot 0. Figure 3 of Reference 28 gives

T
T =0. 98 for the thrust recovery factor.

We have selected the condition a = 5.

The corresponding fuselage angle, a P is therefore, 2. This angle must be

related to a u, the angle of attack indicator reading as this is the parameter
against which the test data are plotted. Figure 30 of Reference 47 is a calibra-
tion of a u vs a F at various values of Ti.

J AtT, =0 aF=2"gives a u= 3.4

Corresponding values of trimmed lift and drag coefficients are then obtained from
JFigure 29 (a) of Reference 47.

CL = 1.72
unpowered

C D = 0.15
I 131
Values obtained for the trim lift and drag of the tail use the following data:

al 0.058/deg.

a2T 0.075/deg.
aoT = 40 (this is an assumption based on the information given in

Reference 47)

eT = 0.7 (value suggested in DATCOM)

Downwash, e, has been evaluated using the method of Appendix B of Reference

49 and is illustrated in Figure 55 as a function of wing lift coefficient.

Downwash angle
at tail
--deg. 16 /

12 --

8 / .

0
1 2 3 4 5
Wing Lift Coefficient
FIGURE 55. ESTIMATED DOWNWASH AT THE TAIL OF BREGUET
941 IN TAKE-OFF CONFIGURATION
At CL = 1. 72 we obtain e = 9. 2 and from Figure 20 of Reference 47 the value of
a e is estimetted to be -0. 5. Evaluating the expressions for tail lift and drag we
get CLT = -0.36 and CDT = 0.04.

Thus the values of C Low and C DDw at aw = 5" are

C .1.72- (-0.36) 320 1.85

Low 8!R9

C = .15- (-0.04) 320 0164

Now, all the information required to calculate CL, CD at the "power-on" con-
dition has been obtained and we substitute in the appropriate euations, as
follows:

132
 CL 1.85+0.98x1.Oxsin (5.00+23.60) x (39)

and
1.6 x I-cos (5.00 x 23. 6"A (40)
C 0.164+0.98xl.Ox
D l+1.Ox (889/684) -Cos (5.00+23.60)

C =2.81andC D =- 0 . 57 at a W =50, T'c = 1.0 (41)

L

The values calculated here must be compared with the test data suitably adjusted
for the trim lift and drag, in the same manner as the Tc, = 0 data was corrected.
At Tc, = I a F = 20 corresponds with au = 4. 80 which corresponds with
CL = 2. 76 and CD = -0. 52 in Figure 29 (a) of Reference 47. Figure 20 of Ref-
erence 47 gives 6e " -5. 60 and from Figure 5 the downw..sh e = 130.

Hence C =-0.965 and CDT = -0.89 and this results in CL = 3.11 and
CD = -0.49l compared with predictions of CL = 2. 81 and CD = -0. 57. Compari-
son of predictions with the test data is shown in Figures 56, 57, and 58.

5.4.3 Wind Tunnel Test Data

The test data used for this correlation was obtained during wind tunnel tests of a
large scale model of a four propeller deflected slipstream configuration that had
no horizontal tail. Thus no trim corrections were required. The data was ob-
* tained from Reference 48, Figure 11(e).
Predictions were made for "power-on" lift and drag at thrust coefficients of 1. 0
and 3. 0 based on the test data obtained at T; = 0 which is approximately equiv-
alent to the "power-off" condition.

-. Relevant data for this correlation are:

." Wing Area 329 ft 2

Propeller diameter 9.3 ft

I Flap type Triple slotted

Flap deflection 50"0

I Extended flap chord 0. 38 local chord

J Correlation between predictions and test data is illustrated in Figure 59.

1 133

I
 Lift
Coef f

CC

5.0 - - - --

-I

'.0

1.0 to -rd

00 2 4 6 8 10 12 14 16 18 20
Wing Angle of A~ttack, qw deg~ees

FIGURE 56. COMPARISON BETWEEN LIFT COEFFICIENT AND

TEST VALUES FROM REFERENCE 47

134
t Drag
Coef f

0.4

0.2

2.

.6~~ -- - -- -- -- - 1--- -- -- -

1 -.2

I.4
-1.4

0 2 4 6 8 10 12 14 16 18 20
Wing Angle of Attack, czw
I FIGURE 57.
degrees
COMPARISON BETWEEN PREDICTED DRAG

I COEFFICIENT AND TEST VALUES FROM

REFERENCE 47

135
 U

17

I-
ii
1 ift
;oe f
CT-
L
5.0

C
T' =1.0
/ /-C
/ / TI =).6
' !/
II

A R 2.T N VF

FIUR
58. COPRSNBTENPRDCE
1.2

Test-
Predictio 1.0
OC OA- -. -0. 0 -. - -. 8

iL

-1.2 -0.8 -0.4 0 0.4 0.8

Drag Coefficient CD i
FIGURE 58. COMPARISON BETWEEN PREDICTED FORCE POLAR
AND TEST VALUES FROM REFERENCE 47 j
136 [

..I-
..
I
i4

E-

I II 0 C,
I

' I

|I 4 - -
a.,
m-

0 0

!~
I--AW N,
i L
I0

Ei4 H

137
5.4.4 Comments on the Predictions

The comparison of the flight test data from Reference 47 with predictions
(Figures 56, 57' and 58) shows poor agreement at the higher values of thrust
coefficient. Errors of 0. 5 in lift coefficient and 0. 2 in drag coefficient occur at
zero angle of attack when the thrust coefficient is 1.6. These numbers represent
at 10% overestimate of take-off speed, an overestimate of drag of about 40% and
an error in climb angle of about 60 for the airplane in question.

Although this represents poor correlation the fault may not lie entirely with the
basic prediction method. The method of calculating the downwash at the tail is
certainly in error to some extent as it is based on the assumption that the wing
wake is not "rolled-up" and neglects all but the gross lift effect in calculating
local downwash.

The comparison of predictions with wind tunnel test data (Figure 59) shows
slightly better agreement than with the flight test data. In this case no trim cor-
rections were required.

It is evident that the prediction method of Reference 28 gives only fair pre-
dictions of lift and drag (at least for the test data considered here) and further
improvements are necessary before it could be used with confidence. S.

5.5 DEFINITION OF DEFICIENCIES AND GAPS IN KNOWLEDGE

5.5.1 General

Successful development of the deflected slipstream configuration requires a

thorough understanding of these aerodynamic effects. I'

The important mutual interference effects between wing and propeller including
stall onset, lateral control power, and the effects of power changes on forces
and moments.

The wake charac.aristics includirg position and thickness of the wake core and
distribution of strength and direction of flow within the wake.

These are also the most difficult characteristics to predict accurately and re-
quire detailed attention to develop an adequate test program.1

5.5.2 Test Data

L.
A large body of force test data has been obtained on a number of basic config-
urations representing the deflected slipstream concept. These data are valuable
in the design and analysis of similar configurations. However, many of the test
programs had omissions that limit the usefulness of applying the data to the

138
 development of basic aerodynamic technology. In particular, the lack of instru-
mentation for the measurement of propeller loads has limited the usefulness of
some of the data. Much of the data has been obtained at specific design-point
conditions and, though useful for certain design work, lacks the systematic vari-
ation of parameters which is important to the development of methodology.

5.5.2. 1 Modeling. Considerable advances have taken place in the development

of small model motors. However, careful attention must be paid to the matching
of motors and propellers to the airplane model.

Power/thermal limitations of electric motors may restrict the thrust coefficient

attainable or require testing at extremely low tunnel speeds to obtain the required
thrust coefficient. Heating of the electric motors requires that special attention
be paid to cooling the motors and the ducting of coolant lines (usually water) must
be carefully integrated within the model. The motor heating can also cause in-
accuracies in the measured propeller forces through temperature drift in the
Ibalances, which must then be compensated.

Hydraulic motors have been developed which do not have the inherent power
limitations of the electric motors but the routing of hydraulic lines complicates
the model design and fabrication and the heating and expansion of the lines can
cause serious interference with the propeller balances.

IPneumatic motors tend to be smaller per unit horsepower developed than electric
motors and therefore lend themselves more effectively to use on small models.
The routing of the air supply around balances and to the motor, however, requires
that unusually detailed attention be paid to model design and to accurate, thor-
ough calibration for data corrections.

I Size and shape of nacelles should be compromised to the minimum extent possible
in order to avoid premature flow separation, particularly with power-off and to
minimize interference with flaps and control surfaces. This includes both phys-
ical interference which restricts control surface size and/or deflection and the
aerodynamic interference, when using air motors, created by the impingement of
tail pipe exhaust flow on the wing and trailing edge surfaces or its influence on
the trailing wake. At high torque conditions, the air motors can produce sizable
tailpipe thrust - much larger, relative to the prop thrust, than would be en-
countered on the full scale airplane. Since it may be important to simulate not
only the correct total thrust coefficient but also the correct thrust-split between
prop and tailpipe, careful attention should be paid to the selection of motor rpm
I during the test.

Motor rpm is also influenced strongly by the model dynamics - with the rpm often
specified by the requirement to avoid resonance bands - and by the relationship
between tunnel speed, required thrust, rpm, and prop blade angle.

1 139

dI
Relatively little has been done, by analysis or experiment, to show what the
effect might be on a wing-flap-slipstream system of simulating more accurately
the propeller span loading and slipstream swirl content. These variables should
be evaluated.

Excited by the motors and propellers, the model's dynamic motion can influence
data collecti)n accuracy, even with high sampling rates and digital recording
equipment. It may be necessary to experiment with tuning systems which corn-
pensate for these effects.

As previously stated, it is of the utmost importance to measure propeller forces

and moments - both to accurately calibrate the thrust coefficient and to extrapol-
ate the data to full scale. The changes in both prop thrust and normal force as
the propeller is pitched through an a -sweep can cause a significant error in the
airplane induced drag if prop instrumentation is not available to account for these
effects.

5.5.2.2 Test Conditions. As noted, the development of small, powerful air

motors has reduced one problem which has plagued powered testing for years-
namely, the requirement to produce high thrust coefficients without testing at
very low tunnel speeds (and correspondingly-at low Reynolds number).

In addition, modern V/STOL tunnels now exist, providing a wide range of con-
trolled tunnel speeds, with large test sections and with the ability to adapt the
wall treatment to the test conditions - from open to slotted to closed.

5.5.2.3 Ground Effect. Some ground effect data has been obtained mainly,
however, on tilt wing configurations. These data may be incomplete in that they
appear to show some Inconsistencies and contradictory trends. The significance
of ground effect to the performance and stability of deflected slipstream air-
planes in landing is shown by flight test (Br 941 for example, Reference 47). The
necessity for realistic ground effect simulation has already been demonstrated
for various configurations and a criterion has been evolved (NASA SP-116) to
determine the conditions for which a moving belt facility is required.

Dynamic ground effects have been noticed, primarily in the flight testing and tun-
nel testing of the XC-142. Very little systematic testing has been done to explore
the dynamic ground effects of deflected slipstream configurations, but the ex-
ploratory work on tilt wings may be applicable.

5.5.2.4 Flow Visualization, A number of different techniques have been used to

visualize the flow from a prop or wing or complete model including smoke,
bubble, tufted screens, schlieren (using local heat addition), etc. None of these
have provent universally acceptable and no standard, practical method of flow
visual ization exists today. The development of a useable technique is of

140

....
II
I paramount importance - both to configuration developmental testing and to devel-
opment of more realistic mathematical models to aid in theoretical studies.

5.5.2.5 Pressure Data. Pressure distributions on wings have been measured on

a number of configurations but mainly in static flow conditions. Such measure-
ments have indicated the large variations of spanwise loading in the part of the
wing in the slipstream due to slipstream rotation. The amount of data is suffi-
cient to give a broad understanding but is inadequate to support an analytical
effort or to verify theoretical prediction techniques.

5.5.2.6 Parametric Force Testing. As previously discussed, a large body of

force test data has ben obtained. However, limitations in test instrumentation
or range of parameters has severely reduced the usefulness of much of the data
for the purpose of creating new configurations or assembling analytical method-
1ology. In addition, it is expected that there is a large amount of data, obtained
during contractor's configuration development, and unpublished NASA testing
which has never been properly analyzed for the development of new methodology.

5.5. 3 Analytical Methods

1 ?The review of the analytical methods reveals that the basic mathematical models
used to simulate the real flow characteristics are inadequate. Furthermore, it
is not clear that the boundary conditions imposed to solve the resulting equations
are valid. It is clear that a much more realistic model must be used including
real propeller effects and the influence of the wing flow induced at the propeller.

I I The empirical methods available are generally capable of predicting the trends in
aerodynamic characteristics due to change of the relevant parameters though the
accuracy is not sufficient without appropriate test data for back-up. In addition,
inconsistc cies have been found in the correlation of these methods with test
data - giving good agreement with one set of data and poor agreement with another.

4, The methods available cover only overall lift and drag and spanwise loading
variation. The calculation of lateral-directional behaviour has not been possible
to any level of accuracy because of the ignorance of the characteristics of the
wake and the influence of prop direction of rotation. For the same reason, it is
difficult to predict the power effects on longitudinal stability and specifically - to
predict analytically the horizontal tail height required to minimize these power
effects.

The direction of propeller rotation influences both the wing's spanwise loading
(and therefore the onset of stall) and the lateral-directional behavior with
asymmetric thr 'qt. Yet, It is not possible to predict these effects.

I1
1 141

i]
5.6 RECOMMENDED PROGRAMS

Outlined below are the necessary programs required to correct the voids in de-
flected slipstream aerodynamic technology.

5.6.1 Wind Tunnel Test Programs

The main areas in which test data are required are STOL propeller character-
istics and the combined propeller-wing-flap flow field.

The first area requires parametric testing of typical STOL propellers in axial F'
and inclined flow conditions over the range of speeds from static to maximum
cruise speed at the appropriate thrust levels. The tests should include measure-
ment of thrust normal force and hub moment and comprehensive studies of the
slipstream structure and trajectory. In addition, the effects on the propeller and
slipstream characteristics of non-uniform inflows (such as may be induced by
high lift wings) should be evaluated.

The second area of importance is the study of the flow field consisting of the com-
bined effects of the propeller slipstreams and the wake of the wing-flap system.
Detailed flow surveys are required for the basic wing-flap system and for the
combined propeller-wing-flap system with variab! - propeller position, overlap,
etc. Half model tests are generally agreed to be suitable for testing incremental
configuration differences in longitudinal characteristics of wing-body-nacelle
configurations, but unsuitable where fuselage aerodynamics or downwash may be
important. Complete model tests should be carried out in order to evaluate char-
acteristics of the flow field for sideslip and yawed flow conditions.

The lateral tests are required because it is necessary to assess the strong
influence of the slipstreams on the lateral characteristics that result from the
proximity of the slipstreams to the fuselage and vertical tail. This is especially
true of testing for "engine-out" stability and control.

Dynamic testing is required to assess the time lag effect as disturba -s at the
propeller are convected in the slipstream since such changes could have strong
effects on the stability and handling characteristics of the configuration.

All of the above testing should be carried out both in and out of ground effect.

5.6.2 Development of Prediction Methods

The lack of a suitably complete and accurate mathematical model of the propeller
slipstream is at least partly responsible for the inability of the resulting wing-
flap-slipstream methods to accurately predict the flow field characteristics and,
in particular, lift and drag of the wing-flap-slipstream systerm.

142
I

It is important to find a practical technique to predict the propeller forces and

Imoments in the presence of a non-uniform inflow, as can occur at high angles of
attack and large values of circulation in the proximity of a wing-flap system. Of
equal importance with the propeller forces and moments !s the corresponding
slipstream vector distribution within which the wing and flap must work.

For an analytical solution to the problem of predicting the flow fields two ap-
proaches are possible.

The mathematical model for the first approach should consist of a bound vortex
representation of the propeller with trailing vortices to simulate the wake. For
the flapped wing a vortex lattice representation is recommended rather than the
Weissinger lifting surface which is probably insufficient for all but the simplest
plain flapped configuration.

The second approach is similar to the first, except that some of the propeller
wake parameters (eg; wake shape, contraction and vortex spacing downstream)
91 may be specified. This approach must await sufficiently detailed test data not
now available on the wake characteristics and on the identification of parameters
which most determine the wake shape.

Empirical techniques for the calculation of overall lift and drag characteristics
"} have been developed from the simplest up to a fairly comprehensive level without
making large gains in the level of accuracy. Items that have typically been
omitted but which should be included in these methods include:

II. The effect of propeller normal force on the induced velocity in the propeller slip-
stream. Even when the normal force is sufficiently small to be negligible the
induced velocity resulting from it can make a significant change in the downwash
angle in the slipstream.

The effect of propeller rotation on the variation of downwash in the slipstream.

This has been considered On Reference 21) by assuming a solid body type of
3 rotation of the slipstream with some measure of success;

The effect of the location of the slipstream relative to the wing. This aspect is
worthy of study particularly if the case arises that the axis of the propeller slip-
stream to far away from the wing chord line (of the order of half a diameter,
say);

JThe panwise influence of propeller-induced flow over the wing. These data may
be obtainable empirically from test data or analytically by vortex-based analysis
of the flow about wing tip panels outboard of a slipstream, or about wings oper-
ating behind overlapped propellers with gaps between them.

143
I
 7
5.7 CONCLUSIONS

The feasibility of the deflected slipstream concept has been demonstrated by a

considerable amount of model test data and by full scale flying hardware both in
test flights and in simulated operations.

Some important test data required for the accurate prediction of performance and
stability characteristics are not available and accurate analytical methods of pre-
diction have not yet been developed. Present empirical methods are of sufficient
accuracy to permit conceptual design studies with confidence only when model
test data is available that is not too dissimilar from the configuration being
studied.

No technical problems are foreseen to prevent the acquisition of the required

test data.

Development of a realistic mathematical model for the analytic solution of the

propeller-wing-flap system and its flow field is, in principle, a fairly simple
matter. However, the solution of the equations resulting from a truly compre-
hensive mathematical model will require large amounts of computer time and
storage, and empirical methods which can produce solutions quickly and inex-
pensively are still needed in the preliminary design process.

144
[I

6. A COMPARISON OF LIFT/PROPULSION CONCEPTS APPLIED

TO A MEDIUM STOL TRANSPORT

6.1 INTRODUCTION

The externally-blown flap concept is a leading contender for a me!um-sized STOL

transport (MST) because of the relative simplicity by which it achieves powered
lift. In this section the technology of the EBF concept is compared to that of four
other lift propulsion concepts which could be applied to an MST configuration:

l Internally Blown Flaps

Augmentor Wing
Direct Jet Lift
Mechanical High Lift Devices

For each of these, the material is assembled here as a guide to the comparative
evaluation of proposed activity in the design, analysis or testing of a medium STOL
transport. Section 2 contains a description of the factors which are common to all
of these concepts in that applicatio- This section treats each lift-propulsion con-
cept separately in terms of data credibility and limitations and criteria for a com-
parative evaluation.

IThe externally blown flap is treated in detail in Section 4. Those data are not re-
peated here, but only referred to in comparison of the externally blown flap with
J other concepts.

6.2 DATA CREDIBILITY AND LIMITATIONS

6.2.1 Externally Blown Flaps

j The state of the technology of external blown flaps as well described in Section 4.
Briefly, the status is this: there are recent wind tunnel tests of good quality which
may be applied to specific configurations of the concept; parametric test data are
still needed, test facilities Pnd techniques are available which can proeuce high
quality data, this report contains a method of calculating lift, drag ane pitching mo-
ment to fill a void noticed in the literature; and the feasibility would now appear to
be only ciptingent upon a development of curmit trbofan transport design.

Typical NASA data are shown in Figure 60 to illustrate the large increase in lift
& forces available with externally blown flaps. These untrimmed lift noefficients are
achieved with thrust coefficients representative of all-engine operation during short
takeoff and landing. Pitching moments which result from this lift are large but are
consistent with the cbordwise centers of pressure of mechanical flap systems. The
ift coefficient as used here includes the large vertical thrust component as defined
in Figure 61. Some effects of configuration geometry upon these components are

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Ij147
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 I.

shown in Figure 62. It is seen that the more even spanwise distribution of thrust
results in improved lift increments. However, in Section 2, it is noted that failure
of engines placed at outboard stations results in large rolling moments. These
considerations affect the spanwise positioning of engines for an actual aircraft
design.

The engine-out condition on the externally blown flap results in the loss of the lift
associated with the dead engine, as well as a rolling moment associated with that -:
lift loss. The rolling moment problem will be most severe at or past stall since j
there will be an asymmetric stall with the Y? dead " wing stalling first as shown in
Figure 63. Lateral control power may place design limitations on the permissible F
spanwise location of engines. Engine-out roll can be minimized by moving the
engines well inboard or by using a number of smaller engines distributed along the
span. T

Section 2 indicates that ground effects are unknown but potentially very important.
The ability to flare in ground effect and reduce rate of sink to an acceptable level
at touchdown may limit the usable lift level.

Although, conceptually, lift coefficient with power is limited only by the amount of
thrust we choose to use, the level that can be used operationally is limited. For
conventional takL Pff and landing (CTOL) aircraft, usable lift has been limited by
speed (or resulting "g') margins, minimum control speed criteria, climb capa-
bility, longitudinal trim capability, etc., as discussed in Section 2. For the ex-
ternally blown flap configuration, appropriate margins and criteria must be deter-
mined for safe flight operation and it is necessary that their effect on performance
be determined. Since CL is dependent on engine operation, consideration of
engine-out lift loss and lift loss due to trimming engine-out rolling moment must
be made in determining the permitted operational lift level.

6.2.2 Internally Blown Flaps

Internally blown flaps have been extensively treated in the literature on boundary
layer control, both analytically and experimentally in both model and full scale.
The many service applications have demonstrated feasibility of the primary con- i
cept. For STOL aircraft, both leading edge and trailing edge blowing (see Figure [1
64) are likely to find application. Figure 65 shows the progress which has been
made in blown flap design and the increase in section lift coefficient due to various
amounts of flap blowing coefficient. The most efficient amount of blowing is that
which isjust sufficient to prevent flow separation and which is applied to the best
available mechanical high lift design. Increases in lift beyond the knee of the curve
are due to the jet reaction component and supercirculation effects. Experiments
have shown that to obtain maximum effectiveness from the larger blowing coeffi-
cients, the blowing should be applied not only to the trailing edge, but a portion of
the blowing should be used at the leading edge, even If mechanical leading edge de-
vices are used. Figure 66 shows that an increase of maximum lift coefficient of

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blown, double-slotted trailing edge flap.

The use of trailing edge and leading edge blowing together permit the achievement
of a desirable range of pitch attitudes and angles of attack for use during takeoff
and landing. The use of very high blowing coefficients and correspondingly high
circulation lift coefficients requires that ground effects be investigated. When the
large blowing coefficients are used, the problem of providing sufficient lateral
control becomes acute, and is often compounded by the desire to use a full span 1,
high-lift system. Blown lateral control devices, other than those demonstrated
already on such aircraft as the NC-130 and augmentor wing demonstrator will re-
quire wind tunnel and functional tests to prove effectiveness and practicality. 1.
6.2.3 Augmentor Wing

Augmentor wing technology has developed rapidly, drawing upon jet flap and BLC
background, although distinct from either, and drawing upon large-scale wind
tunnel testing. Geometry of the augmentor wing flap is depicted in Figure 67,
which shows that the jet which issues from the nozzle is directed not to attach to a
surface as in the usual BLC, but to mix with entrained air from the slots provided,
so as to obtain ejector action. At the present time there appear to be no analytical
methods especially applicable to augmentor wing configurations.
Typical longitudinal characteristics of an augmented jet flap wing are shown in
Figure 68. These are unpublished NASA test data and were obtained in the Ames
40 by 80 foot wind tunnel on a 44.15 foot wing span model that geometrically simu-
lated a CV-7A aircraft with an augmented jet flap extending over 55 percent of the
wing span. Blown ailerons extended from the flap to wing tip. The compressed
air for the augmentor was supplied by axial flow compressors with their turbines
driven by exhaust gases of a jet engine. A J85 turbojet was installed under each
wing with an exhaust diverter valve simulating a rotating type nozzle with capacity "
for vectoring thrust aft for takeoff and cruise and downward for approach and land-
ing. With the augmentor flap deflected but with augmentation flow off, the CLMA I.
is about 2.3. With jet augmentation the CLM 1o 5.7 untrimmed and with the
addition of 1500 lbs. of thrust vectored 850 downward the CLMA X becomes 6.8 or
an increase nCLofl.l. This isabout 35 percent higher than Just the 1500
pound, of thrust would produce as a vector thus indicating an increase in circula-
tion lift due to the vectored jets.

The two curves labeled "early flap design" are for a more complicated augmentor
flap design and show that the performance of the present simplified flap is supe-
rior. Blowing the knee of the early flap showed insuflicient improvement in lift to
warrant the added complexity.

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One of the strongest points favoring the augmentor wing is that it produces high lift
coefficients with relatively low increases in pitching moments compared to mc
other high lift concepts. For example, at constant a, as the amount of blowing is
varied to increase lift coefficient the ratio of ACL/AC M is 8 to 10 while the &CL/
ACM for a slotted flap is about 3. This means that for a given increase in wing
-, lift coefficient the amount of lift that must be sacrified to trim the airplane with an
augmentor wing is about one third of that for a slotted flap. Therefore the useable
lift is higher.

IUsing elevator effectiveness from the above test data, the pitch acceleration capa-
bilities of a typical 40,000 pound aircraft employing an augmentor wing were cal-
culated r a sea level standard day as shown in Figure 69. The aircraft is in the
landing coDiguration with the augmentor flap deflected 75 degrees. The line
drawim through the calculated points represents the control power from a conven-
tione.l tail that is capable of a ACLTAIL = .55 at full deflection. Indications are
that the requirements of Reference 1 would be satisfied above a velocity of about 69
knots, which id a speed constant with STOL operation from field lengths of about
II 2000'.

The augmentor wing lateral control data presented in Figure 70 are also from the
unpublished NASA tests and are for the landing configuration with the flap at 75
degrees and the starboard aileron drooped to 45 degrees. The port aileron deflec-
tion is varied from its normal droop of 45 degrees landing position to 0 and 65
degrees.

With the ailerons operating in this manner at all times data are presented showing
CQ, C and C w en inboard spoilers are applied and when the augmentor on one
wing panel is throttled to produce roll control. Two cases of blockage were tested-
25 and 5G percent of the outboard flap semispan. In both caswd 75 plvcent of the
3 augmentor exit area was blocked by . of wedges.

The data show that # maximum rolling moment coefficient of asbout .115 could be
3attained with the ailerons alone. The ailerons and inhiard spoilers combined pro-
duce Ce a .2.

3Throttling the augmentor 25 an 50 percent of its span, and applying aileron, pro-
duced maximum rolling moment coefficients o, about .21 and .26 respectively.

Using the lateral control data of the unpubllthed NASA test pre'iously reirred to,
the rolling angular acceleraton provided by each Individual roll control device was
calculated for a 40,000 pound airplane at sea levei vwlth the results shown on Figure
J1I. The aileron was operative when the augmentor wv throttled and also when the
spoilers were tested so that the rolling effectiveness of aileron plus spoiler and
aileron plus throttled augmentor can be obtained by adding the components. The
I spoilers were not tested while throttling was applied so their iaterference effects
on each other are not known. However, it will probably be suJiciently accurate to

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To obtain sufficient roll control, all three devices will probably have to be used.

STOL minimum speed margins of augmentor-wing aircraft must allow for reduc-
tions in maximum lift consistent with provision of lateral control. Lift losses
obtained from the wind tunnel tests discussed above showed these values as
increments:

z LMA LOSS

00/650 blown aileron .6 to .8

Spoilers outboard .1

Spoilers inboard .4

25% span augmentor throttled .25

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Ground effect data on lift are available for one augmentor wing configuration and
are shown in Figure 72.

The data were obtained using a fixed ground board. It should be noted from the
criteria of Figure 18, that for the h/5 = 1.3 and lift coefficients of 3.85 and 4.4
at which these tests were run, that a fixed r ound board is sufficient.

Noise is of concern with all powered lift devices, including the augmentor wing.
,-, In this case the source of noise is expected to be primarily the mixing region
between the air from the primary nozzle and that entrained through the ejector
slots. It is expected that the presence of surfaces on both sides of the mixing
region will make the augmentor wing amenable to acoustic treatment.

Tests were conducted in 1970 at NASA Lewis to evaluate the acoustic character-
istics of an augmentor wing co figu1ration and an externally blown flap configura-
tion. Large scale models were used, as depicted in Figure 73, to investigate the
near and far field noise and azimuthal pattern above and below the flaps. The
tests were planned to obtained additional data on turning effectiveness, panel flut-
ter and thermal effects.

16.2.4 Direct Jet Lift

The direct lift engine concept appears to be a simple selection to the problem of
achieving high lift, but is also subject to the same two major difficulties that af-
fect most of the other STOL concepts: low-speed control and ground effects.
practical solutions have already been demonstrated for both of these areas.

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Satisfactory roll control for VTOL (and that should be sufficient for STOL) has
been demonstrated by using the main lift engines for lateral control as on the Dor-
nier Do-31 or by using auxiliary reaction controls as on the Hawker-Siddeley P-
1127. The category of "ground effectd' as applied to lift engine configurations, i;
includes not only the interaction of the lift engine flow upon the aerodynamic lift,
drag and moments, but also the recirculation of hot gases or debris from the
ground into the engine and the effect on ground erosion.

Direct lift engines can heavily influence the flow field and the aerodynamic charac-
teristics of the vehicle depending upon location of the engine relative to fuselage,
wing, aerodynamic high-lift devices and tail surfaces.

For example, there are three major ways in which lift engines cause changes in
the pitching moments of an aircraft even if they are concentrated at the aircraft
I
center of gravity as shown in Figure 74. The intake momentum drag and the ex-
haust thrust component tend to pitch the airplane up. This pitching moment varies
with the engine length, thrust produced and aircraft forward speed.
1:
Changes inlift are produced by jet induction on the surfaces fore and aft of the I,
engine inlet and outlet which vary with engine location, forward speed and engine
size. The shape and size of the surfaces surrounding the inlet and outlet are also
important.

Vorticity generated by the exhaust may complicate stability and trim of the vehi-
cle by causing irregularities in the flow field of the tail. These effects must be
considered not only in the aggregate, as when considering a complete configura-
tion, but also in the individual component contributions as when designing and
calibrating a wind tunnel model.

Physically, lift engines offer a wide choice of places where they can be mounted.
They can be placed in the fuselage or mounted in pods on the wing or fuselage.
Because of this versatility in mounting, many things must be considered before
selecting a location. If they are placed far from the aircraft's center of gravity
they may seriously affect moments of inertia yet this location is desirable if they E
are used also for longitudinal, lateral and directional control. The pitch and roll
moments resulting from engine failure must be considered as well as reingestion
and aerodynamic interference in ground effect.

The problem of mechanical installation of engines, and fuel and control lines and ]
of providing thermal protection for structure and other components as well as E
their maintenance, may be a deciding factor in chosing one location rather than
another.

Though the lift engines are only used during takeoff and landing, cabin noise lev-
els should be considered especially if commercial use is anticipated.

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0 0 0

I eed-

I 165
Work is being done to reduce the noise of lift engines by acoustical treatment.
Theoretical predictions of possible noise reduction are encouraging but full scale
tests are required for verification.

Figure 75 shows the noise levels for a bypass ratio 0 turbo-jet engine, a bypass
ratio 5 cruise engine and a bypass ratio 12 lift fan. The noise level of these three
F.
sources shows that the treated cruise engine is still slightly noisier than the aux-
iliary lift systerns. The dominance of fan generated noise tends to make noise
F
levels constant with bypass ratio for a given state of the art. Possibly the future
will offer improvements in basic fan noise for cruise engines that will allow use
of bypass ratios above 5 in cruise.
F
The direct lift concept has been proved feasible down to zero speed by flight test
of the Do-31.

The addition of lift engines cannot be considered as just " strapping on extra cans
of lift" because their high velocity exhausts can cause serious losses in the circu-
lation lift of the high lift wings with which they are associated. This interference
is so dependent upon configuration that wind tunnel tests must be conducted for a
particular design. The model testing is complicated by the high lift engine ex-
haust velocities and it mev be that inlet velocities should be simulated.

Ground effects can be very critical and they are so dependent upon the particular
configuration that tunnel tests with a ground board and perhaps a moving belt are
required. I
Although theoretical analysis indicates a good noise reduction potential exists for
lift engines, a full scale test demonstration should be conducted.

6.2.5 Mechanical High Lift Devices

Mechanical high lift devices - that is, the non-powered-lift category which includes
the leading and trailing edge flaps, vanes, slats and similar articles - are well
founded in theory and experiment.

Figure 76 portrays the improvement in flap design that has occurred during the
years from 1947 to the present. The improvement in maximum lift coefficient of
the 737 from 2.6 to 3.53 indicates what a well planned and intense design effort '
can accomplish. Based on this experience it is estimated that with reduced sweep
and a triple slotted flap with about 30 percent Fowler action a CLMAX of 4.0 is I
attainable. With a full-span high lift device, the provision of lateral control may
require some ingenuity, such as the development of mechanisms which produce
flaperon action with a triple-slotted flap.

Figure 77 shows data obtained on a straight wing semiapan model with a full span,
large Fowler action, triple slotted flap. A leading edge slat of 15% chord was

166

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deflected 20 degrees. A maximum untrimmed lift coefficient of about 4.2 was
obtained partly due to the 42% chord increase from Fowler action. As is typical
with flaps of this type, the large nose down pitching moments result in a large
reducton in useable lift when trimmed.

For comparison, lift data were obtained with the same model operated as a T"
double-slotted flap. Medel geometry is as shown on Figure 78. The same wing
panel was used, but a 35%-Fowler-action double-slotted flap configuration was
used instead of a triple-slotted flap. The high antrimmed maximum lift coeffi-
cient shown on Figure 79 will be considerably reduced when the high pitching
moments are trimmed.

Lift coefficients of this magnitude are achievable with complete aircraft configu-
rations. Figure 80 shows data from Reference 48 which reports test results from
a large scale, four propeller STOL airplane with a triple slotted flap and a leading
edge slat.

To obtain the high lift coefficients required for STOL o-eration from a mechanical
high lift device will probably require large full span flaps with much Fowler
action. This leads to a serious lateral control problem. The lateral control
would have to be obtained by using a combination of lateral controls. Sections of
the flap could be used - aperons. Spoilers would undoubtedly be required. if
a portion of the wing wt. reserved for ailerons they would be blown for maximum
effectiveness and, for very low-speed operation, it may be necessary to use
reaction controls.

The vast amount of data that has been accumulated through the years is for flaps
that develop lift coefficients of about 2.8, but very little is available for the higher
lift coefficients required for STOL. Test data are needed on control power -
especially lateral control and yaw coupling. Ground effects on the aerodynamic
coefficients at high values of lift are a subject for further testing.

There is very little ground effect data on mechanical high lift wings at the high
values of lift coefficient required for STOL aircraft. Tests have shown that the
influence of ground effect increases at higher lift coefficionts and also as the
trailing edge of the flap gets closer to the ground at high flap deflections.

Wind tunnel tests of mechanical high lift models capable of producing lift coeffi-
cients of about four are required. These tests should be made using a moving
rather than a fixed ground board where possible.

6.3 CRITERIA FOR A COMPARATIVE EVAL ,ATION

In evaluating the relative merits of STOL aircraft designs based upon the several
lift/propulsion concepts, important differences in the concepts will be apparent
from an appraisal of cost effectiveness, reliability, maintainability and a host of

170
 'I

iI7
:!

I ,m

6321

ALTEewSTE 5%.. • VAME 4

AI
AtC. ~lI.SECToO0N

FIGURE 78. GEOMETRY OF DOUBLE AND TRIPLE SLOTTED

IFLAP MODEL

17I

I
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CCA
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W V~ v\/V
~V ~ ~ r-4~r!

C3

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IICl
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AL
-, 10
0 1.0 -10 0 to 20 3o 0 -1

REFERENCE:, NASA TND 4448

NOTES: 1. R,, 3x10 6

2. AR= P.06 b=56ft. c 7.3ft
3. 632 k!6 airfoil
-. 4. C' =0

I GEOMETRY OF LEADING EDGE SLAT AND TRIPLE SLOTTED FLAP

FIGURE 80. MECHANICAL HIGH LIFT DEVICES ZERO THRUST

LONGXTUDINAL CHARACTERISTICS OF 4 PRO-
PELLER DEFLECTED SLIPSTREAM STOL AIRCRAFT
WITH LEADING EDGE SI.AT AND TRIPLE SLOTTED
FLAP17
other factors as well as the technological factors which affect the STOL aerody-
namic performance. The latter are treated here in six headings:

Performance
Handling qualities
Ground effects
Ride quality *

Noise
Failure characteristics

These factors, some of which are discussed in more detail in Section 2, are re-
capped here.

The performance flight envelope of the aircraft must be clearly defined for the all-
engines-operative mode. Flight path angle and maneuver G capabilities as a func-
tion of speed and altitude should be compared for competing aircraft.

The same type of comparison should be made for the most-critical-engine out
condition

Another important evaluation item is the sensitivity of the aircraft to configuration

changes such as flap setting, blowing and thrust deflection, and to what extent
these configuration changes may be permitted during normal takeoff or landing
operations, abort operations, or transitions between the two. It is important to
know the consequences of flight beyond the normal flight envelope boundaries, in
order to establish margins of safety. These should be converted into equivalent
minimum speeds. Reverse thrust should be available with flow directed to result
In the least disturbance to the aircraft itself and to surrounding objeers, so as to
provide a useful reverse taxi capability.

In evaluating the flying qualities of STOL aircraft, in addition to the conventional ::

stability and control considerations, it is important to consider the rate of change
of characteristics with speed. At the low speeds of STOL operations, the changes
d0 do dO dTEQ-
of control power d - - and uTREQ (which will probably be on the back side
iV' dV' dV dV
of the thrust curve) may have an important bearing on minimum speed require-
ments. o.

The sensitivity of the aircraft to gusts during approach and landing is critical in
STOL aircraft because the gust can be such a large percentage of the low approach -
speed. Aircraft should be compared on the basis of the forces and moments re-
sulting from a foot per second of gust velocity. Vertical, lateral, longitudinal
and asymmetrical gusts should be considered.

The approach attitude should provide good visibility to the pilot.

174
 If a flare is required, the amount of rotation and the pitch acceleration available
to obtain it should be evaluated and considered in establishing field length factors.
The ability to develop lift without rotation is a strong asset in a STOL aircraft.

The presence of control cross coupling in any portion of the flight envelope should
be accounted for, particularly when the use of lateral control may affect the usable
lift.

The susceptibility of STOL aircraft to being tipped over while parked or during
ground handling should be considered. The relatively low wing loading high-
aspect-ratio straight wings with large flaps with the trailing edge near the ground
may make them subject to tip over by side angled gusts. Landing gears should be
designed to compensate for this and to provide a desirable turn radius during
taxiing. To provide good STOL characteristics the reverse thrust mechanism and
basic braking system must: reliable and effective and any possible asymmetries
should be easily controllable.

* The ability to taxi backwards by means of reverse thrust must be provided, espe-
cially for use on short, unprepared, fields that do not provide tractors for reverse
taxi. On fields of this type the design of the reverse thrust unit must preclude the
possibility of the reverser flow causing reingestion problems.

During all ground handling operation accomplished by application of the aircraft's

• own power, the resultant noise levels of the competing aircraft should be evalu-
ated.

Ground effect on the high lift systems of STOL concepts cannot be predicted theo-
retically but must be obtained by wind tunnel tests on the specific configuration
proposes. The evaluation of any STOL concept should demand of the proposer
sufficient test data to be convinced that the capabilities of the concept in ground
effect are known with a reasonable degree of confidence.

Force and moment effects and instabilities induced in ground effect must be evalu-
ated as well as changes in control power.

The recirculation and foreign object damage possibilities of some concepts such as
direct lift engines may be greater than those of another concept such as mechani-
cal high lift devices.

The ground signature or trailing vortices resulting from very high lift concepts
during takeoff and landing could prove potentially dangerous for another aircraft
and may require dimensional separation of STOL aircraft in ground operations or
in the traffic pattern.

The STOL aircraft will likely be used for rather short missions and therefore will
spend a relatively large portion of its flight time at low altitude in turbulent air.

J 175

I
A STOL configuration with low wing loading, high aspect ratio, low sweep and a
high lift curve slope embodies features which are detrimental to smooth ride qual-
ities. Also, the need for large flaps which reduce the size of the wing structural -.

box may affect the elastic gust response in both cruise flight and during STOL
operation. The aircraft should be evaluated on the basis of acceleration per ft. L
per second of gust velocity for not only the rigid body but also the elastic case.

Gust alleviation systems may be provided and their effectiveness and reliability
should be assessed especially if credit is taken in weights trend calculations for
reduced load factor history.

STOL aircraft have high power loadings and, achieving lift by directing high veloc- -
ity gases, can have high potential noise levels. These noise levels, both internal
(in the cockpit and cabin) and external, should be evaluated and compared for
different proposed designs. This should be done for takeoff, approach and cruise
conditions.

When noise criteria are specified for takeoff and landing in terms of distance and
azimuth, the power levels and flap settings and other configuration variables may
be limited and the flight path and STOL distance may be affected.
Evaluation of the failure characteristics of an aircraft involves determination of
the effect of the failure on the STOL performance of the aircraft and its ability to
complete the mission.

It is important to know the effect of an engine failure on trim and control power in
evaluating performance.

Other critical component failures could occur in interconnects including ducting,

valves, control links, etc. Some of these are inherent possibilities in some con-
cepts and not in others. Air duct failures would be suspect in augmentor wings
and internally blown flaps and not in mechanical and externally flaps systems.
SAS, electrical, hydraulic, brakes and thrust reversal systems are common to all
concepts.

It is not only necessary to evaluate the results of failures, especially those inher-
ent to particular lift concepts, but it is important to assess the relative probability
of their occurrence and the severity of the consequences after occurrence.

6.3.1 Internally Blown Flaps

High lift levels are attainable with internally blown flaps much of it due to super-
circulation. As with any high aerodynamic lift device and especially when large
chord highly deflected flaps are employed, ground effects are large. A high wing
configuration seems preferable in order to reduce ground effect and for other -.
reasons.

176
 Engine development work is required to efficiently furnish the flap blowing air
from high-bypass ratio engines which will produce the large thrust to weight ratios
-needed for STOL operation.

A design effort is needed to quantify the weight and configurational effects of the
large high temperature air ducts in the wing and to verify their fatigue life.

Blown flaps require small aircraft rotation angles for takeoff but do produce large
pitching moments. By cross ducting a portion of the engine blowing air from each
engine to the flap on the opposite wing rolling moments resulting from an engine
failure can be kept small.

The initial cost of internal blcwn flap STOL aircraft may be higher than for con-
ventional aircraft due to additional cost for engines with blowing capability and for
Sducting, however the aerodynamic benefit in performance and controllability may
-; be substantial.

I:.6.3.2 Augmentor Wing

The augmentor wing is capable of producing very high lift coefficients with much
of the lift due to supercirculation. Because of its high lift, ground effects are
large and a high wing design is indicated.

To efficiently supply the augmentor blowing air some engine development work
will be required. Design work will also be needed to insure that the large high
temperature gas ducts are efficient and have a long fatigue life.

Some takeoff thrust augmentation will be available.

The noise level may be potentially low due to the high equivalent bypass ratio and
. rapid mixing of the augmentor air. Also, by the nature of its design, it is amen-
able to acoustic treatment.

Engine out rolling moments will be small if a proper portion of the blowing air
from each engine is ducted to the augmentor on the opposite wing.

The initial cost of the du-'ng and flap system will be high as will its maintenance.

6.3.3 Direct Lift Engine Concept

The effect of the lift engines on the lift of the basic wing is unknown in ground
effect. Wind tunnel tests on a particular proposed configuration must be conducted
to determine tins effect.

Lift engines are in limited development and have been used successfully on the
Dornier DO-31, which is the best example of a lift-jet transport configuration.

177
 U
Engine reingestion problems must be avoided during landing and when reverse
thrust is applied.

The installation of lift engines with their fuel and control lines is complex as is the j
addition of the required inlet and exhaust louvers.

The use of lift engines is a potentially simple and effective method of providing
STOL performance from a strictly aerodynamic point of view. By obtaining lift
this way, smaller wings can be used which will result in higher wing loadings in
cruise and good ride comfort.

Lift engine developments promise favorable noise characteristics.

In locating lift engines on the aircraft care must be taken to keep pitching and
yawing moments resulting from an engine out as small as possible.

To obtain the best aerodynamic performance, STOL configurations with lift engines
must bear the additional cost of lift engine development and procuremet in the
initial cost and of the maintenance of these multi-engine combinations in the dilect
operating cost.

6.3.4 Mechanical High Lift Devices

There is a large quantity of wind tunnel and flight test data available on mechanical
high lift devices. As a result the technical risk on this concept is low. Also no
new engines need be developed specifically for this concept.

A STOL aircraft depending completely on mechanical high lift, rather than powered
lift devices would probably have a low wing loading and high gust reaction response
unless a gust alleviation system is used.

It is difficult to prevent interaction of the lift and propulsion systems, however,

and the moments resulting from engine failure may require consideration of special
control system or interconnected propulsion (as with a cross-shafted deflected
slipstream propeller driven vehicle).

6.4 CONCLUSIONS

Figure 81 compares the maximum or stall lift coefficient and the usable lift coeffi-
cient for five lift concepts. The usable lift reflects the approach CL rules that
provide margins necessary for maneuver, gusts and instrument or operation
errors. The externally blown flap with an all engine operative CL stall at lg of 6.2
represents a CL stall of 5.4 with one engine out. This results in a usable CL of
only 3.7.

178
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z .E-4

4~- ' H Hr)

A 'P4

Z))

i
r, o kU

1___9I
If the line marked usable on the direct lift engine bar represents the usable lift of
the aircraft with the lift engine thrust set at zero, the lines above it represent in-
crements in" effective CL stall" that is available by adding lift engines. New stall
and maneuver margins might be considered in this design because the lift engine
thrust is insensitive to gusts.

Figure 82 compares the noise levels of various STOL aircraft types designed for
the same mission and takeoff distance. The basis is the external blown flap which
had the highest noise level in takeoff, partly due to a focusing effect of the flap on
the noise. Each of the other high lift concepts has a noise level lower than that of
the external blown flap in both takeoff and approach.

In takeof, the turbojet or turbofan direct lift designs are 12 to 13 PNdB quieter
than the externally blown flap and 6 to 7 PNdB quieter than the conventional me-
chanical flap or internally blown flap designs. The advantage of the direct lift de-
signs is that less propulsion is lost overcoming the high drag associated with high
aerodynamic lift. On the design with high CL'S considerable potential climb thrust
is sacrificed to overcome the high drag and the rate of climb is reduced with a
resultant loss in altitude and increase in community noise.

The approach case indicates that unpowered lift designs are significantly better
than powered designs. The externally blown flap aircraft is noisy because it re-
quires about an 80 percent power setting to maintain the desired descent gradient
and aerodynamic lift. Improvement, in direct lift systems in approach do look
promising as more is learned about inlet fan noise suppression techniques.

Recent studies indicate that the differences shown between concepts are extremes
and can be reduced by engine cycle changes or increased bypass ratio, but the in-
crease in engine size, aircraft weight and fuel weights reflect a large penalty of -o

noise constraints upon configurations with STOL performanc,.

Table XI summarizes the availability of analytical prediction methods, criteria

and test data for the five lift concepts being considered.

It will be noted that there is a need for an analytical method for determining ground
effects for every lift concept. It should be added that the same need exists for
wind tunnel data on ground effect.

Another field where information is lacking for each STOL concept 1o basic criteria
for performance, stability and control, and design. The main reason for this
deficiency Is that there is such a small experience backgrtwmd available on opera-
ting STOL aircraft. Many years and many aircraft have contributed to the evolu-
tion of these criteria for CTOL aircraft, both c,*.,il and military, and it isdesire-
ble to evolve STOL criteria from operational experience with powered-lift
aircraft.

180
iI

(1.5 Naut. Miles from Brake Release)

0
(Ref.)
External
Blown
-* Internal Flap
-5 - Mech. Blown
Flap
.0 Flap Flap
-- a
z

Jet
"- Flap
--

U) Turbo
•HO * Jet Turbo
-15L- Fan

1APPROACHI
(1 Naut. Mile from Threshold)

"" C H (Ref.)
6External
Blown
Fla

-5
Jet Turbo
Flap Jet
Turbo

Flap Flap

I Fa

FIGURE 82. NOISE COMPARISON

It is concluded that, if the MST takeoff and landing distance is not less than 2,000
feet, each of the five high lift concepts is feasible. It may be desirable to combine
concepts. For example, it may be desirable to add lift engines to the mechanical
high lift wing to enable use of a smaller wing. Internal BLC may be needed to
provide lateral control for the externally-blown configuration.

The wind tunnel data that are available are valid.

Lateral control limits the STOL performance of the five high lift systems. Pow- o.
ered lift systems such as blown flaps and the augmentor wing are control limited
for the critical-engine out condition. To obtain STOL performance with mech-
anical high lift will require full or near full span flaps. Lateral control will
probably have to be obtained by using a portion of the flap as a flaperon (which is
difficult to mechanize), plus spoilers. The lateral control engine out problem of
the augmentor wing can be alleviated to some degree by ducting a portion of the
blowing air from each engine to the flap on the opposite wing so that in the event
of an engine failure neither wing flap will be completely unblown.

The most serious deficiency in aerodynamic knowledge of high lift devices is

ground effect on the aerody.,amic coefficients and particularly on lift. Theory and
available tes t.s show that the higher the lift coefficient the more sensitive is the
high lift device to ground effect. Very little ground effect test data are available
at the high lift coefficients required for STOL performance.

Ground effect is so dependent on configuration that detailed tests must be con-

ducted on the particular design selected.

A flight demonstration is recommended for several reasons. The development of

CTOL aircraft has been an evolutionary process through the years. This process
has not been available to STOL aircraft. Some of the aircraft that have been used
to obtain STOL experience were basically V/STOL types - not STOL and a back-
ground is needed for STOL operational and design criteria.

4.

-S

182 ii

__________________________________________________
I1 V

z z

~zz z z

a) co

Vo 'go

Ea)

18a)
 7. CONCLUSIONS

Of all the STOL concepts studied, the technology base is greatest for the deflected --
slipstream, tilt wing, and mechanical high lift concepts and smallest for the
externally-blown-flap and jet flap types. Nevertheless, there are sufficient tech-
nical data for all of the concepts to show that each could be developed into a
feasible STOL aircraft system.

The low speed , peration required to take off or land in short distances must be
obtained by increasing the lifting capability through propulsive-lift augmentation li
or low wing loadings. Either course creates additional problems. Powered lift
tends to increase sideline noise, aerc'lynamic ground effects, and system com- --
plexity. Low wing loading reduces comfort of ride quality. Techniques exist in
today's technology to solve all of these problems.

STOL aircraft require a set of performance and flying qualities criteria which
differ from those applied to conventional aircraft. Many different sets of criteria
have been proposed and studied. None has been adopted. Additional operational
experience with STOL flight demonstrators is a necessary prerequisite to such an
action.

Wind tunnel testing of STOL aircraft is more complex and requires more detailed
attention than that required for conventional aircraft. The following factors re-
quire special attention: matching of model size to tunnel size, ground effect testing
including requirements for a moving ground plane, model motor selection and
installation, instrumentation, and flow visualization.

Much of the available test data have limited applicability resulting from incomplete
calibration or the omission of measurements of parameters later found to be im-
portant. The development of analytical methods for prediction of aerodynamic
characteristics of STOL aircraft would benefit from systematic testing of these
configurations.

Methodology for deflected slipstream includes several analytical and empirical

force prediction methods. These have limited applicability. No such methodology
was found for the externally blown flap. A preliminary method has been developed
and is presented in Section 4 of this report.

Additional method development is required for all STOL concepts to improve the
understanding of the comple%- aerodynamic effects of these configurations.

184
I
RF FERENCES

1. Innis, R. C., Holzh'iser, C. A., and Quigley, H. C., AIRWORTHINESS CON-

.. SIDERATIONS FOR STOL AIRCRAFT, NASA TND5594, National Aeronautics
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2. TENTATIVE AIRWORTHINESS STANDARDS FOR VERTICRAFT/POWERED

LIFT TRANSPORT CATEGORY AIRCRAFT, FAR "XX", Federal Aviation
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portation, Washington, D.C., July 1970.

3. Turner, T.R., ENDLESS BELT TECHNIQUE FOR GROUND SIMULATION,

NASA SPl16, Conference on V/STOL and STOL Aircraft, Ames Research
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il
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186
!I
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