1. Master thesis in Aeronautical Engineering
Development of a Flight Dynamics Model
of a Flying Wing Configuration
Candidate Supervisor
Jacopo Tonti Prof. Guido De Matteis
External supervisor
Prof. Arthur Rizzi
(Kungliga Tekniska högskolan)
Academic Year 2013/2014

3. cbn 2014 by Jacopo Tonti
Some rights reserved. This thesis is available under a
Creative Commons CC BY-NC-3.0 IT License.
(creativecommons.org/licenses/by-nc/3.0/it/)

5. Alla mia musa,
a Vale

7. «As ailerons, these damn spoilers make great rudders!»
Bruce Miller, after flying the Marske Pioneer 1A

9. Abstract
The subject of UCAV design is an important topic nowadays and many countries have
their own programmes. An international group, under the initiative of the NATO RTO
AVT-201 Task group, titled “Extended Assessment of Reliable Stability & Control Pre-
diction Methods for NATO Air Vehicles”, is currently performing intensive analysis on
a generic UCAV configuration, named SACCON. In this thesis the stability and control
characteristics of the SACCON are investigated, with the purpose of carrying out a compre-
hensive assessment of the flying qualities of the design. The study included the generation
of the complete aerodynamic database of the aircraft, on the basis of the experimental data
measured during TN2514 and TN2540 campaigns at DNW-NWB low speed wind tunnel.
Moreover, system identification techniques were adopted for the extraction of dynamic
derivatives from the time histories of forced oscillation runs. The trim of the aircraft was
evaluated across the points of a reasonable test envelope, so as to define a set of plausible
operative conditions, representing the reference conditions for subsequent linearization of
the dynamic model. The study provided a thorough description of the stability and control
characteristics and flying qualities of the unaugmented SACCON, laying the groundwork
for future improvement and validation of the configuration in the next design stages.
Keywords: Aerodynamic Modelization, System Identification, Stability & Control, Linear
Dynamics, Flying Qualities, Flying Wing, UCAV, SACCON.

11. Table of Contents
Contents i
List of Figures v
List of Tables ix
Nomenclature and Symbols xi
Frames of reference . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xi
Notations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xii
1 Introduction 1
1.1 Background of the NATO RTO program . . . . . . . . . . . . . . . . . . . . 2
1.2 Problem description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
1.3 Objective and methodology . . . . . . . . . . . . . . . . . . . . . . . . . . . 5
1.4 Thesis outline . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7
2 Literature review 9
2.1 Historical perspective . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9
2.1.1 Modern stealth UCAVs . . . . . . . . . . . . . . . . . . . . . . . . . . 14
2.2 An overview on flight mechanics analysis . . . . . . . . . . . . . . . . . . . . 15
2.2.1 Static stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15
2.2.2 Dynamic stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
2.2.3 Flying and handling qualities . . . . . . . . . . . . . . . . . . . . . . 20
2.2.3.1 Cooper-Harper rating scale . . . . . . . . . . . . . . . . . . 20
2.2.3.2 MIL-HDBK-1797A . . . . . . . . . . . . . . . . . . . . . . . 22
2.2.3.3 CAP criterion . . . . . . . . . . . . . . . . . . . . . . . . . 27
2.3 Flying wing design issues . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
2.3.1 Longitudinal issues . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29
2.3.2 Lateral-directional issues . . . . . . . . . . . . . . . . . . . . . . . . . 30
3 Aerodynamic database 33
3.1 Foreword . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33
3.2 Wind tunnel campaigns . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35
i

12. TABLE OF CONTENTS
3.2.1 Wind tunnel model . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35
3.2.2 Experimental setup . . . . . . . . . . . . . . . . . . . . . . . . . . . . 37
3.2.3 Tests and results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38
3.3 Database generation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42
3.3.1 Database format . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45
3.3.2 Static data processing . . . . . . . . . . . . . . . . . . . . . . . . . . 47
3.3.3 Dynamic data processing . . . . . . . . . . . . . . . . . . . . . . . . 49
3.4 Aerodynamic analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50
3.4.1 Baseline . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51
3.4.2 Dynamic behavior . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53
3.4.3 Control authority . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56
4 Static analysis 63
4.1 Flight envelope definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64
4.1.1 Airspeed limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . 65
4.1.2 Altitude limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . 66
4.1.3 CG limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66
4.2 Longitudinal static stability . . . . . . . . . . . . . . . . . . . . . . . . . . . 68
4.3 Trim assessment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70
4.4 Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79
5 Dynamic analysis 81
5.1 Aerodynamic identification . . . . . . . . . . . . . . . . . . . . . . . . . . . . 81
5.2 Dynamic modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 87
5.2.1 Longitudinal dynamics . . . . . . . . . . . . . . . . . . . . . . . . . . 88
5.2.2 Lateral-directional dynamics . . . . . . . . . . . . . . . . . . . . . . . 95
5.3 Flying qualities assessment . . . . . . . . . . . . . . . . . . . . . . . . . . . 106
5.3.1 Longitudinal flying qualities . . . . . . . . . . . . . . . . . . . . . . . 108
5.3.1.1 Short period . . . . . . . . . . . . . . . . . . . . . . . . . . 108
5.3.1.2 Phugoid . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 110
5.3.2 Lateral-directional flying qualities . . . . . . . . . . . . . . . . . . . . 112
5.3.2.1 Roll subsidence . . . . . . . . . . . . . . . . . . . . . . . . . 112
5.3.2.2 Dutch roll . . . . . . . . . . . . . . . . . . . . . . . . . . . . 113
5.3.2.3 Spiral . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 114
5.3.3 Control dynamics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 115
5.3.3.1 Response to step elevator . . . . . . . . . . . . . . . . . . . 115
5.3.3.2 Response to step aileron . . . . . . . . . . . . . . . . . . . . 116
5.3.3.3 Response to step rudder . . . . . . . . . . . . . . . . . . . . 118
5.4 Chapter summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 120
ii

13. TABLE OF CONTENTS
6 Concluding remarks 123
6.1 Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 123
6.2 Further research . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 127
Appendices 131
A SACCON configuration 131
A.1 General description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 131
A.2 Mass and inertia properties . . . . . . . . . . . . . . . . . . . . . . . . . . . 132
A.3 Geometric properties . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 133
B Theoretical basis and definitions 135
B.1 Physical model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 135
B.1.1 Assumptions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 135
B.1.2 Coordinate systems and transformations . . . . . . . . . . . . . . . . 136
B.1.3 Mathematical relations . . . . . . . . . . . . . . . . . . . . . . . . . . 138
B.2 Conventions and customs . . . . . . . . . . . . . . . . . . . . . . . . . . . . 140
B.2.1 Control sign convention and definitions . . . . . . . . . . . . . . . . . 140
B.2.2 Aerodynamic parameters convention . . . . . . . . . . . . . . . . . . 142
B.2.3 Propulsion system customs . . . . . . . . . . . . . . . . . . . . . . . 143
B.2.4 Mass and geometry . . . . . . . . . . . . . . . . . . . . . . . . . . . . 143
C Linearized Model 145
D XML database structure 151
D.1 Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 151
D.1.1 Fundamental table structure . . . . . . . . . . . . . . . . . . . . . . . 151
D.2 Database structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 152
D.2.1 Aerodynamics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 153
D.2.2 Geometry and mass . . . . . . . . . . . . . . . . . . . . . . . . . . . 157
D.2.3 Propulsion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 158
D.2.4 Flight control system . . . . . . . . . . . . . . . . . . . . . . . . . . . 159
Bibliography 163
iii

15. List of Figures
2.1 Nature’s noteworthy flying wing designs. . . . . . . . . . . . . . . . . . . . . 9
2.2 The Penaud and Gauchot “Amphibian” - 1876 [46]. . . . . . . . . . . . . . . 10
2.3 Dunne’s D.8 flying wing biplane - 1912 [52]. . . . . . . . . . . . . . . . . . . 11
2.4 Chyeranovskii BICh-17 experimental fighter - 1934. . . . . . . . . . . . . . . 12
2.5 The Horten Vc - 1941. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
2.6 The Northrop-Grumman B-2 “Spirit” - 1989. . . . . . . . . . . . . . . . . . . 13
2.7 Modern stealth flying wing UCAV designs. . . . . . . . . . . . . . . . . . . . 14
2.8 Pitching moment curves (fixed elevator) [2]. . . . . . . . . . . . . . . . . . . 16
2.9 Conventional wing-tail arrangement [2]. . . . . . . . . . . . . . . . . . . . . 17
2.10 Dynamic response of a statically stable aircraft [52]. . . . . . . . . . . . . . 18
2.11 Cooper-Harper rating scale [52]. . . . . . . . . . . . . . . . . . . . . . . . . . 21
2.12 MCH-UVD diagnosis tool [33]. . . . . . . . . . . . . . . . . . . . . . . . . . 22
2.13 CAP requirements for Category B flight phase [35]. . . . . . . . . . . . . . . 28
2.14 Northrop N-1M. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31
2.15 The drag rudder deployed on the wing tip of the Northrop N-9M. . . . . . . 32
3.1 Planform and geometric parameters of the DLR-F17SACCON [16]. . . . . . 36
3.2 The DLR-F17/SACCON in the DLR-NWB with yaw link support [47]. . . . 37
3.3 Lateral coefficients of the DLR-F17 versus α at different β [15]. . . . . . . . 41
3.4 Influence of sting mounting on longitudinal coefficients (Body axes) [38]. . . 42
3.5 The frame of reference convention adopted in TN 2514 and TN 2540 [20]. . . 43
3.6 Baseline drag and lift coefficients versus α, varying β. . . . . . . . . . . . . 51
3.7 Baseline pitching moment coefficient versus α, varying β. . . . . . . . . . . . 52
3.8 Baseline lateral-directional coefficients (Body frame) versus β, varying α. . . 53
3.9 1-cycle average of lift driven by pitch oscillations [20]. . . . . . . . . . . . . 54
3.10 1-cycle average of pitching moment driven by pitch oscillations [20]. . . . . . 54
3.11 1-cycle average of lateral coefficients driven by 1 Hz roll oscillations [20]. . . 55
3.12 1-cycle average of lateral coefficients driven by yaw oscillations. . . . . . . . 55
3.13 Elevator contribution to lift. . . . . . . . . . . . . . . . . . . . . . . . . . . . 57
3.14 Elevator contribution to pitching moment. . . . . . . . . . . . . . . . . . . . 58
3.15 Total lift and pitching moment with elevator. . . . . . . . . . . . . . . . . . 58
3.16 Rolling and yawing moments induced by the ailerons. . . . . . . . . . . . . . 59
v

16. LIST OF FIGURES
3.17 Rolling and yawing moments induced by the drag rudders. . . . . . . . . . . 60
3.18 LCDP. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62
3.19 Cnβ DYN. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62
4.1 The analysis envelope of the SACCON . . . . . . . . . . . . . . . . . . . . . 67
4.2 Limit locations of the CG of the SACCON (in red the ARP). . . . . . . . . 67
4.3 Variation in static margin with CG position and angle of attack. . . . . . . 69
4.4 Variation of static margin with CG position and velocity. . . . . . . . . . . . 70
4.5 Flow chart diagram of the double variable iteration procedure. . . . . . . . . 72
4.6 Variation of angle of attack to trim with CG location. . . . . . . . . . . . . 74
4.7 Variation of elevator to trim with CG location. . . . . . . . . . . . . . . . . 75
4.8 Map of angle of attack to trim versus altitude and CG location. . . . . . . . 76
4.9 Map of elevator to trim versus altitude and CG location. . . . . . . . . . . . 76
4.10 Sketch of forces acting on an airplane in horizontal level flight. . . . . . . . . 77
4.11 Variation of required thrust with CG location. . . . . . . . . . . . . . . . . . 78
4.12 Variation of aerodynamic efficiency with CG location. . . . . . . . . . . . . 79
5.1 Comparison of the longitudinal combined dynamic derivatives calculated
with the two linear methods for 1 Hz oscillations. . . . . . . . . . . . . . . . 85
5.2 Roll motion combined dynamic derivatives. . . . . . . . . . . . . . . . . . . 86
5.3 Yaw motion combined dynamic derivatives. . . . . . . . . . . . . . . . . . . 86
5.4 Short period root locus varying airspeed and static margin at sea level. . . . 89
5.5 Phugoid root locus varying airspeed and static margin at sea level. . . . . . 90
5.6 Phugoid normalized shape at U8 “ 150 m/s and h “ 0 m for K » ´5%. . . 91
5.7 Third mode root locus varying airspeed at h “ 0 m (K » 0%). . . . . . . . . 92
5.8 Third mode normalized shape at U8 “ 70 m/s and h “ 0 m for K » 0%. . . 92
5.9 Short period root locus at h “ 12 000 m. . . . . . . . . . . . . . . . . . . . . 93
5.10 Phugoid root locus at h “ 12 000 m. . . . . . . . . . . . . . . . . . . . . . . . 94
5.11 Tumbling normalized shape at U8 “ 100 m/s and h “ 0 m for K » ´5%. . . 95
5.12 Dutch roll root locus varying airspeed and static margin at sea level. . . . . 96
5.13 Dutch roll normalized shape in different flight phases for K » ´5%. . . . . 97
5.14 Dutch roll shape approaching coalescence at U8 “ 200 m/s for K » ´5%. . 98
5.15 Variation of lateral-directional derivatives with airspeed at sea level. . . . . 99
5.16 Variation of yaw damping derivative N1
r with airspeed at sea level. . . . . . 100
5.17 Variation of dutch roll and spiral poles with airspeed at sea level (K » ´5%).100
5.18 Variation of spiral pole location with airspeed and SM at sea level. . . . . . 101
5.19 Variation of spiral pole location with airspeed and SM at h “ 12, 000 m. . . 102
5.20 Variation of dynamic stability parameters with airspeed at sea level. . . . . 104
5.21 Dutch roll root locus varying altitude and static margin at U8 “ 200 m/s. . 105
5.22 Analysis envelope with prescribed flight phases. . . . . . . . . . . . . . . . . 107
vi

17. LIST OF FIGURES
5.23 Short period degree assessment. . . . . . . . . . . . . . . . . . . . . . . . . . 108
5.24 Short period frequency variation with airspeed and CG position. . . . . . . 109
5.25 CAP assessment for short period characteristics. . . . . . . . . . . . . . . . 110
5.26 Phugoid degree assessment. . . . . . . . . . . . . . . . . . . . . . . . . . . . 111
5.27 Phugoid frequency variation with airspeed and CG position. . . . . . . . . . 111
5.28 Roll subsidence degree assessment. . . . . . . . . . . . . . . . . . . . . . . . 112
5.29 Dutch roll level assessment at sea level (Category A). . . . . . . . . . . . . 113
5.30 Spiral divergence degree assessment. . . . . . . . . . . . . . . . . . . . . . . 114
5.31 Longitudinal response to step elevator at sea level. . . . . . . . . . . . . . . 115
5.32 Longitudinal response to step elevator at 12 000 m. . . . . . . . . . . . . . . 116
5.33 Variation of roll rate response to step aileron with airspeed (K » ´5%). . . 117
5.34 Variation of roll angle response to step aileron with airspeed (K » ´5%). . 118
5.35 Variation of sideslip angle response to step aileron with airspeed (K » ´5%).118
5.36 Lateral-directional response to step rudder in different flight conditions. . . 119
A.1 Designed arrangement of the control surfaces of the SACCON [37]. . . . . . 131
A.2 Wing airfoils of the SACCON UCAV configuration [16]. . . . . . . . . . . . 133
A.3 Radius distribution and relative thickness of the SACCON configuration. [16].134
B.1 Attitude angles and positive direction of velocities. . . . . . . . . . . . . . . 136
B.2 Wind axes arrangement and positive signs of α and β. . . . . . . . . . . . . 137
B.3 Positive deflection of fundamental control surfaces. . . . . . . . . . . . . . . 140
B.4 Relative arrangement of Geometry (blue) and Body (black) frames. . . . . . 143
D.1 Comprehensive view of the structure of the input file. . . . . . . . . . . . . . 153
vii

19. List of Tables
2.1 Definition of handling quality levels in MIL-HDBK-1797A [35]. . . . . . . . 23
2.2 Definition of flight phase categories in MIL-HDBK-1797A [35]. . . . . . . . . 24
2.3 Short period requirements in MIL-HDBK-1797A [35]. . . . . . . . . . . . . . 24
2.4 Phugoid requirements in MIL-HDBK-1797A [35]. . . . . . . . . . . . . . . . 25
2.5 Roll subsidence requirements in MIL-HDBK-1797A [35]. . . . . . . . . . . . 25
2.6 Spiral requirements in MIL-HDBK-1797A [35]. . . . . . . . . . . . . . . . . 26
2.7 Dutch roll requirements in MIL-HDBK-1797A [35]. . . . . . . . . . . . . . . 26
3.1 Prospect of control deflection combinations tested during TN 2514 and TN 2540
at DNW-NWB (OB and SP are mutually exclusive). . . . . . . . . . . . . . 39
3.2 Prospect of forced oscillation tests of TN 2514 and TN 2540 at DNW-NWB. 40
3.3 Template of a generic spreadsheet with three states. . . . . . . . . . . . . . 46
3.4 Control authority of the SACCON at α “ 50 and β “ 00 (in 1/rad). . . . . . 60
4.1 Summary of SACCON stability and trim issues. . . . . . . . . . . . . . . . . 80
5.1 Spiral eigenvector variation due to dutch roll collapse at sea level (K » ´5%).103
5.2 Lateral departure eigenvector at sea level and U8 “ 250 m/s for K » ´5%. 104
5.3 MIL-HDBK-1797A aileron response requirements for Class II aircrafts. . . . 117
A.1 Mass and inertia properties of the SACCON. . . . . . . . . . . . . . . . . . 132
A.2 Geometric properties of the SACCON. . . . . . . . . . . . . . . . . . . . . . 133
ix

21. Nomenclature and Symbols
Frames of reference
NOTE
Body frame is chosen as default frame of reference. Thus, for the sake of clarity, the
superscript for all quantities expressed in that coordinate system will be omitted.
Frame Tag Definition [1, 5]
Fixed Vertical E
O: at sea level on the vertical of the aircraft initial position
x: pointing North, tangent to meridians
y: pointing East, tangent to parallels
z: pointing to the center of the Earth
Local Vertical V
O: aircraft center of gravity
x: pointing North, tangent to meridians
y: pointing East, tangent to parallels
z: pointing to the center of the Earth
Wind W
O: aircraft center of gravity
x: in the direction of velocity relative to air
z: in the aircraft plane of symmetry, pointing toward its belly
y: consequently, on the right
Stability S
O: aircraft center of gravity
x: in the direction of the projection of velocity relative to air on
the plane of symmetry
z: in the aircraft plane of symmetry, pointing toward its belly
y: consequently, on the right
Body B
O: aircraft center of gravity
x: along fuselage reference line, pointing toward the nose
z: in the aircraft plane of symmetry, pointing toward its belly
y: consequently, on the right
Geometry G
O: foremost point of the aircraft fuselage
x: parallel to fuselage reference line, pointing toward the tail
z: in the aircraft plane of symmetry, pointing towards its topside
y: consequently, on the right
xi

22. NOMENCLATURE AND SYMBOLS
Notations
NOTE
Vectors are denoted by an under line; matrices by bold symbols; tensors by a double arrow;
time derivatives by a dot 9pq; space derivatives by a prime pq1, dimensionless quantities by
a hat ˆpq. Names of variables and parameters used in the present document and within
the simulator are, in general, the combination of a symbol and a subscript. The frame of
reference to which a variable is referred is specified (if applicable) as superscript tag.
Some symbols and acronyms may be used as subscripts.
Symbol Tag Description Unit
A“ b2{S - Aspect ratio -
B B Engine gyroscopic moment N ¨ m
C C Aerodynamic coefficient -
D D Drag N
E “ CL{CD E Aerodynamic efficiency -
F F Force vector N
F - Aerodynamic force vector N
G - Linear momentum kg ¨ m{s
H - Angular momentum kg ¨ m2{s
I I Element of the inertia tensor kg ¨ m2
K “ kn ´ k - Longitudinal static margin (Body frame) -
L L Lift N
L - Rolling moment (never used as subscript) N ¨ m
M8 Mach Free stream Mach number -
M M Moment vector N ¨ m
M - Pitching moment (never used as subscript) N ¨ m
M - Aerodynamic moment vector N ¨ m
N - Yawing moment (never used as subscript) N ¨ m
P P Power W
Q Q Engine torque kg ¨ m
S S Surface m2
T T Thrust N
T BA T_BA Rotation matrix from frame A to frame B -
U8 TAS Free stream velocity (true airspeed) m/s
W W Weight N
X, Y , Z X, Y, Z Aerodynamic force components N
a A Acceleration m{s2
a8 a Free stream speed of sound m/s
b b Span m
xii

23. NOMENCLATURE AND SYMBOLS
c c Chord m
f - Frequency Hz
g g Gravitational acceleration m{s2
h alt Altitude m
k - Distance in ratio of reference chord -
m m Mass kg
n n Load factor -
p8 pres Free stream air pressure Pa
p, q, r p, q, r Roll, pitch, yaw rates rad/s
q8 “ 1
2ρ8U8
2
qd Free stream dynamic pressure Pa
t t Time s
v v Velocity vector m/s
u, v, w u, v, w Velocity components m/s
x x Position vector m
x, y, z x, y, z Position components m/s
Θ8 temp Free stream temperature K
Λ - Sweep angle deg
Ξ euler Attitude angles vector deg
Ω Omega Engine nominal rpm 1/s
α alpha Angle of attack deg
β beta Angle of sideslip deg
γ gamma Flight path angle deg
δ d Control surface deflection angle deg
ζv zeta_v Engine vertical incidence deg
ζh zeta_h Engine horizontal incidence deg
ϑ theta Pitch angle deg
κ “ πfx{U8 - Reduced frequency (x is a reference length) -
λ “ ctip{croot - Taper ratio -
ρ8 rho Free stream air density kg{m2
ϕ phi Roll angle deg
ψ psi Yaw angle deg
ω omega Angular velocity vector rad/s
Subscript Tag Description
0 0 Initial or equilibrium condition
a a Aileron
air air Airbrake
asym asym Asymmetric deflection
xiii

24. NOMENCLATURE AND SYMBOLS
c c Canard
eng eng Engine
e e Elevator
glob glob Global
in in Inboard
l l Rolling moment
ldg ldg Landing gear
lef lef Leading edge flap
m m Pitching moment
n n Yawing moment
out out Outboard
port (LX) port (LX) Port (left side)
r r Rudder
ref ref Reference
root root Root of the aerodynamic surface
spl spl Deceleron (split surface)
spo spo Spoiler
star (RX) star (RX) Starboard (right side)
sym sym Symmetric deflection
t t Throttle
tef tef Trailing edge flap
tip tip Tip of the aerodynamic surface
x, y, z x, y, z x, y, z axis
Acronym Tag Description
AC AC Aerodynamic center (1/4 root chord in general)
ARP ARP Aerodynamic reference point
CG CG Center of gravity
CP CP Center of pressure
DoF - Degrees of freedom
ERP ERP Engine reference point
FCS FCS Flight Control System
LE - Leading edge
mac - Mean aerodynamic chord
NP - Neutral point
SM - Static margin
SPO - Short period oscillation
S&C - Stability and Control
TAS TAS True airspeed
TE - Trailing edge
xiv

27. Chapter 1
Introduction
The flying wing design is a very attractive configuration due to the aerodynamic perfor-
mance advantages it offers over its conventional counterpart, primarily higher aerodynamic
and structural efficiency. However, the omission of horizontal and vertical stabilizers leads
to even severe stability and control authority issues. Further difficulties arise from the
problem of fitting the pilot, engines, flight equipment, and payload all within the depth
of the wing section. Moreover, due to the practical need for a thick wing, the flying wing
concept is relegated in the slow-to-medium speed range. These compromises are difficult
to reconcile and efforts to do so can reduce or even negate the expected advantages of the
flying wing design, such as reductions in weight and drag. These reasons contributed to tie
to tailless designs the reputation of being impracticable and notoriously difficult to control.
As a consequence, no flying wing aircraft has ever been designed and flown successfully for
the civil aerospace sector, where the technical advantages associated with such a config-
uration are easily outweighed by the requirements on safety and degradation in handling
qualities. Not even the military sector, with its relaxed aviation regulation requirements,
along with the advent of fly-by-wire control systems, ever considered the flying wing con-
figuration appealing and feasible at first glance, since its flaws seem to easily match or
overcome its strengths.
However the flying wing, despite the disadvantages inherent in its design, gained re-
newed interest, due to its potentially low radar reflection cross-section. In fact, stealth
technology relies, among other features, on the presence of flat surfaces that coherently
reflect radar waves only in certain directions, thus making the aircraft hard to detect unless
the radar receiver is at a specific position relative to the aircraft. Hence, the flying wing
proves to be a successful design whenever stealth requirements become a prominent con-
cern, especially since the introduction of modern fly-by-wire flight control system (FCS)
allowed to mitigate the stability and control deficiencies affecting the design.
Nevertheless, of the few attempts made to build a military flying wing aircraft (espe-
cially by the Horten Brothers and Northrop Corporation [45]), only one survived premature
cancellation and entered mass production: the Northrop B-2.
1

28. 1. Introduction 1.1: Background of the NATO RTO program
During the last decade, the ever more widespread development of increasingly advanced
unmanned combat air vehicles (UCAV) has definitely boosted the interest of the aerospace
community on flying wind design. The massive use of state-of-the-art flight control systems,
along with the lack of issues related to the presence of crew onboard, has allowed for less
conventional configuration to be adopted with much more confidence, by exploiting their
full potential, while alleviating their flaws.
1.1 Background of the NATO RTO program
The ability to accurately predict both static and dynamic stability characteristics of
air vehicles using computational fluid dynamics (CFD) methods could revolutionize the
vehicle design process for air vehicles. A validated CFD capability would significantly re-
duce the number of ground tests required to verify vehicle concepts and, in general, could
eliminate costly vehicle “repair” campaigns required to fix performance anomalies that were
not adequately predicted prior to full-scale vehicle development. As a result, significant
reductions in acquisition cost, schedule, and risk could be realized. Historically, many mil-
itary aircraft development programs have encountered critical stability and control (S&C)
deficiencies during early stages of flight test or worse still in service, despite thousands
of hours of wind tunnel testing. These surprises have occurred across the speed range
from takeoff and landing to cruise flight, and particularly at the fringes of the operating
envelope, where separated and vortical flows dominate [14].
The dramatic increase in computing power and the affordability of PC clusters now
make it feasible to undertake time-accurate high-fidelity CFD simulations. Current capa-
bilities of CFD to predict static and dynamic stability of aircrafts were assigned an overall
Technology Readiness Level (TRL) of 4. For such a reason, it has been widely believed
within the research community that the next significant improvement in the state-of-the-art
for predicting the S&C characteristics of a new vehicle might be through the application
of CFD tools. In particular, the application of high-end CFD codes with a Reynolds-
avereged Navier-Stokes (RANS) or a better level of technology to specific areas of S&C
interest before first flight can help focus the wind tunnel program and provide improved
understanding of the underlying flow physics. When S&C parameters are determined by
wind tunnel or CFD methods, the basic principle is to determine forces and moments on
an aircraft when it is undergoing oscillations in pitch, roll or yaw. While S&C problems
most commonly occur near the margins of the flight envelope, the computational solution
of S&C for even steady flight conditions can be demanding. Fortunately, a critical mass of
international researches has been assembled to address this challenge through the North
Atlantic Treaty Organization (NATO) Research and Technology Organization (RTO).
2

29. 1.1: Background of the NATO RTO program 1. Introduction
In 2007, a 3-year international collaboration was initiated through the NATO1 RTO
Applied Vehicle Technology AVT-161 titled “Assessment of Stability and Control Predic-
tion Methods for NATO Air and Sea Vehicles”. The aim of the task group was to investigate
the applicability of current CFD tools for predicting S&C characteristics of air and sea ve-
hicles. The specific objectives were to 1) assess the state-of-the-art in computational fluid
dynamics methods for the prediction of static and dynamic stability and control charac-
teristics of military vehicles in the air and sea domains; 2) identify shortcomings of current
methods and identify areas requiring further development, including aspects of compu-
tational uncertainty [17]. Improvements to the prediction capability (specifically in the
ability of turbulence models to predict flow separation around blunt surfaces) were sug-
gested by AVT-161 and are being carried forward by AVT-183 (2010-2014). In addition,
AVT-161 has led to a specialists meeting task group, namely AVT-189, where the results
and progress of AVT-161 will be evaluated by experts external to AVT-161. The success
of AVT-161, and the input and identification of potential new partners from AVT-189,
has motivated the establishment of the AVT-201 task group in 2012 (scheduled to end in
2015), with the desire to pursue and extend the efforts of AVT-161 in using CFD methods
to predict S&C characteristics of aircrafts [14].
The target configuration conceived and intensively studied by the RTO team is a generic
UCAV geometry called “Stability And Control CONfiguration”, or simply SACCON. The
design consists of a highly swept lambda wing with a planform primarily devised to meet
marked stealth requirements and later adjusted to improve its aerodynamic performances.
The detailed description of the aircraft is presented in Appendix A.
The main objective of the ongoing AVT-201 is to determine an overall strategy for cre-
ating S&C databases for vehicle simulation at full-scale conditions, including the deflection
of control surfaces, throughout the operational envelope of the vehicle. The investigations
carried out have the purpose to assess the usefulness of engineering methods not only as
an analysis tool during the early aircraft design, but also as a design tool to improve the
shape definition of the vehicle, in order to achieve better performance [21].
The topics to be covered in the pursue of the main objective are summarized below [14].
• Perform additional in-depth correlation studies, using the detailed flow field mea-
surements obtained by AVT-161 to enhance understanding of discrepancies between
predicted and experimental dynamic derivatives.
• Carry out further wind tunnel testing to extend the dynamic data set to include
multiple frequency and amplitude maneuvers, in order to obtain, where possible,
full-scale test data for a maneuvering vehicle, that can be used for validation of the
1
In addition to the participating NATO members, Sweden (FOI) and Australia (DSTO) were also
invited to join the task group.
3

30. 1. Introduction 1.2: Problem description
methods and capabilities that are being developed.
• Design, build and test modified S&C wind tunnel model with trailing-edge control
surfaces, to evaluate the ability to predict control effectiveness, stability character-
istics, and other flight dynamics characteristics of the configuration with controls
deflected.
• Investigate techniques for generating flight simulation models from CFD predictions,
that is build S&C databases from experimental and CFD data to determine level of
accuracy and sensitivity of flight simulation using CFD when compared with exper-
imental data model.
• International collaboration, specifically the concept of “virtual laboratory”, as pio-
neered by AVT-113, and used to great effect in AVT-161.
The key benefits expected from the research are the reduction of project risk and design
cycle time and, hence, of sub-scale testing, manpower and financial resources; as well as
an enhanced understanding and prediction capabilities of aircraft characteristics before
fabrication, leading to an early identification and explanation of unexpected behavior,
through the medium of reliable flight simulation [14].
1.2 Problem description
The challenges faced by the task group of the RTO program when designing and testing
the SACCON configuration are manifold, but they can be roughly summarized as follows.
1) Optimization of the planform for attainment of both stealth requirements and ade-
quate aerodynamic performance.
2) Execution of two parallel campaigns of wind tunnel experiments and computational
studies carried out with several CFD tools.
3) Understanding of the extensive non-linear behavior of the vortical flow field generated
by the highly swept wing at a wide range of angles of attack and sideslip, in order to
improve the accuracy of numerical modelization.
4) Understanding of the particular nuances of each distinct CFD grid and flow solver,
before broad application of the tools in this class of problems.
5) Generation of two complete aerodynamic databases, one with experimental (wind
tunnel) data and the other with CFD data, to be used to “fly” the aircraft in a
simulator.
6) Evaluation and analysis of S&C characteristics, based on both databases, in order to
assess the reliability of CFD predictions as a design tool since early design stages.
4

31. 1.3: Objective and methodology 1. Introduction
Points 1), 3) and 4) have been extensively treated as part of the research of AVT-161.
Wind tunnel tests and CFD computations, point 2), were carried out in several rounds
during both AVT-161, where several leading edge and surface treatments were tested, and
AVT-201, where control deflection and engine flow effects were measured.
This thesis work focused on the issues represented by points 5) and 6), that is the generation
of a complete aerodynamic database of SACCON and the subsequent evaluation of its S&C
characteristics, limited to the data provided by wind tunnel campaigns of AVT-201.
1.3 Objective and methodology
The main objective of the thesis work was to use the experimental data provided by
wind tunnel tests, carried out at DNW-NWB facility as part of the AVT-201 research
schedule, to perform a complete analysis of the stability and control characteristics and
flying qualities of the SACCON aircraft, a lambda wing flying wing design, whose detailed
description is provided in Appendix A. The results thus obtained will serve as reference for
future validation of S&C characteristics prediction based on CFD estimations, ultimately,
to assess the reliability of CFD methods as design, as well as analysis, tools.
In the context of providing exhaustive results about S&C characteristics of the SACCON
aircraft, the underlying principles were to develop a robust and general algorithm to per-
form the calculations and to define a convenient format for the aerodynamic database to
be processed. The latter to be adopted when generating the database from the CFD data,
so as to facilitate data correlation and validation.
The work methodology adopted to achieve the target set was analytical and can be
broken down into the following aspects.
• As a first step, it was defined a structure for the aircraft database suitable for fast S&C
analysis. This includes, (i) the choice of the file format and of (ii) the structure of the
tables (multidimensional or spreadsheet), (iii) the identification and implementation
of all parameters necessary for the S&C algorithm. Particular regard was used in the
setup of the aerodynamic portion of the database, in order to facilitate processing
without comprising data readability.
• Then all available wind tunnel results were aggregated into a single complete aerody-
namic database of the SACCON. This database was then consolidated with geometry
and mass, propulsion and flight control system data, which were integrated in a fash-
ion similar to that of the aerodynamic portion. Without any information about
mass distribution of the SACCON (neither wind tunnel model, nor full-scale), these
parameters were deducted by comparison to a comparable aircraft [29].
• The next phase involved the analysis of the aerodynamic data collected in the
database of the aircraft, with the purpose of (i) identifying the range of linear aerody-
5

32. 1. Introduction 1.3: Objective and methodology
namic behavior, expected to prevail at least during the cruise and loiter segments of
the mission profile of the UCAV [21]; (ii) examining the aerodynamic characteristics
beyond the condition of departure of the linear behavior, dominated by nonlinear flow
effects. Moreover, during this stage the data were manipulated in order to isolate the
contributions of control surfaces for a preliminary assessment of their effectiveness
on the aircraft forces and moments.
• Next logical step was to shape an analysis envelope for the SACCON, in order to
define a set of plausible operative conditions, each defined by a combination of air-
speed, altitude and CG location. The scope of the analysis envelope is to investigate
the sensitiveness of the S&C characteristics and flying qualities of the aircrafts to ve-
locity, altitude and centering. Again, no information regarding aircraft performances
nor mass distribution were available, so those of comparable aircrafts were adopted
and, if necessary, scaled. In particular, the same performance figures and powerplant
of the Dassault nEUROn were considered.
• Static analysis and trim evaluation were then carried out over the full flight envelope,
with the purpose of (i) exploring the influence of altitude and CG location on the
equilibrium characteristics of the SACCON, especially elevator deflection and aero-
dynamic efficiency, and (ii) provide a first estimation of required thrust in straight
leveled flight. Moreover, the results were used to adjust the limits of the flight enve-
lope, taking into account stall and control surface saturation.
• Finally a comprehensive linear dynamic analysis was conducted over the revised flight
envelope. The outcomes of the analysis were then used to (i) evaluate the dynamic
behavior of the SACCON, quantifying the sensitiveness of its dynamic response to ve-
locity, altitude and CG location, to (ii) verify the compliance of the design’s handling
qualities with MIL-HDBK-1797A aviation regulations throughout the flight envelope
and to (iii) highlight the control and handling problems of the configuration and
outline a feasible strategy to make it a practical design.
The work described in this dissertation was initiated at the Department of Aeronautical
and Vehicle Engineering (Farkost och Flyg) at the KTH - Royal Institute of Technology
in Stockholm. There a research team, led by Prof. Arthur Rizzi, was conducting intensive
CFD research on the SACCON, in close collaboration with other teams worldwide, as
active Swedish member of the AVT-201 and former participant in the foregoing AVT task
groups. The work was eventually concluded at “Sapienza” - Università di Roma under the
supervision of Prof. Guido De Matteis.
The aerodynamic model and subsequent analyses presented in this dissertation are
based on the data provided by DLR (Deutsches Zentrum für Luft- und Raumfahrt e.V.)
and were obtained at DNW (German-Dutch wind Tunnel) NWB low speed wind tunnel.
6

33. 1.4: Thesis outline 1. Introduction
1.4 Thesis outline
Following this introduction, the subject matter of the dissertation is organized in sub-
sequent chapters summarized below.
Chapter 2 presents a detailed literature review. It introduce a brief history of the de-
velopment of flying wing aircrafts, together with a review of modern UCAV
designs. Then a concise review of the basic aspects of aircraft stability and
control analysis is reported. Finally, the advantages and drawbacks, espe-
cially stability issues, associated to this category of air vehicles are covered,
along with the attempts made to improve the design.
Chapter 3 describes the generation of the database of the SACCON. First the back-
ground of wind tunnel test is introduced, followed by the review of the
procedure adopted to merge all data into a single database. The results of
the aerodynamic and control authority analyses are then discussed.
Chapter 4 covers the definition of the flight envelope and the execution of stability and
trim analyses. The outcomes are then discussed and applied to update the
prototype envelope.
Chapter 5 presents the details of the linearization process of the model and the results
of subsequent analysis (both longitudinal and lateral aspects are covered).
It also discusses the control and handling qualities aspects of the considered
airframe and identifies the critical areas and drawbacks.
Chapter 6 concludes this dissertation. It summarizes findings, identifies limitations
and sets directions for further research work on the subject.
7

35. Chapter 2
Literature review
By virtue of the aim of the research and of the nature of its target configuration,
the literature review first presents a brief history of the flying wing design, followed by a
review of the latest stealth UCAV development projects. It then describes the distinctive
characteristics of the flying wing concept and their issues, especially regarding stability
and control. An overview of the main aspects of flight mechanics and handling qualities
analyses concludes the chapter.
2.1 Historical perspective
Tailless aircrafts have been experimented with since the earliest attempts to fly. By
merely looking at birds circling in the sky, the flying wing possibly represents the simplest
and most immediate configuration imaginable when venturing the design of a flying ma-
chine. Nature itself seems to reinforce the idea that a flying machine consisting of a single
wing is, in fact, a practicable design, as successfully adopted by soaring/gliding birds, bats,
plant seeds and, in prehistoric times, pterosaurs (see Fig. 2.1). It is no coincidence that
the later journals of Leonardo da Vinci contain a detailed study of the flight of birds and
several different designs for wings based, in structure, upon those of bats.
However, at a closer look, bird-like and aircraft flight mechanics differ radically, essentially
(a) Albatross. (b) Fruit bat. (c) Zanonia seed.
Figure 2.1: Nature’s noteworthy flying wing designs.
9

36. 2. Literature review 2.1: Historical perspective
because of the capability of living creatures to move their bodies, thus benefiting from
wing flapping and dynamic shifting of the center of gravity. The movable surfaces of air-
crafts, necessary for their flyability, somehow mimic the first feature, by locally changing
the shape of the machine, but only at an incredibly lower scale, with effects not even nearly
comparable to those achieved by the wings of animals.
In the early days of the aeronautical era the preeminent challenge dwelled in the design
of stable and controllable configurations, prior to focus on the improvement of pure perfor-
mance. However, despite all the amazing progresses and technological leaps made by the
aerospace industry in the past century, the appearance of the conventional aircraft, i.e. a
main wing paired to set of tail surfaces, all attached to a more or less cylindrical fuselage,
has not changed significantly since the beginning of the flight era. This is mainly due to
the extreme level of confidence reached in the aforesaid configuration, herein referred to
as conventional configuration, regarding stability and control issues, as well as structural
reliability and on-board system integration.
Many unconventional configurations have been explored over the last century, some of
which even proved (at least theoretically) better characteristics for certain aspects than the
standard one, but none has yet been able to replace it, nor become comparably widespread.
The reason dwells in the fact that unconventional configurations exhibit, in general, poorer
S&C characteristics and/or necessitate more complex structural designs. Moreover, con-
sidering that, until a few decades ago, the technical knowledge was not mature enough
to provide the tools and expertise to adequately support the development of unconven-
tional aircrafts, it is clear why such configurations invariably failed when compared to the
standard design.
Figure 2.2: The Penaud and Gauchot “Amphibian” - 1876 [46].
Nevertheless, for more than a century, there have been countless patents, projects,
and concepts relating to tailless airplanes. Many models and prototypes were constructed;
most enjoyed only a brief period of development and public interest, and then quickly
10

37. 2.1: Historical perspective 2. Literature review
disappeared. The lack of adequate financial backing, lack of government or public interest,
and politics often contributed to the premature end of a worthwhile project.
Whatever the source of inspiration, most designers persevered with their experiments and
research despite the lack of experimental facilities and financial backing. The critical period
was when experiments passed from the model and glider stages to powered flight. For the
reasons previously cited, projects were often terminated altogether at this stage. In other
cases, the problems of stability and control associated with the absence of the tail proved
insurmountable, so a conventional tail was added [45].
The pioneering age of aviation is replete with frustrated, brilliant men, to whom the
description “neglected genius” could have applied. The greater contributions of the period
came from European engineers, inventors or simple enthusiasts, who studied the flight
characteristics of every conceivable type of flying creature. They all attempted the trial
of sustained and controlled flight, often by means of flying wings, and, although they met
scarce success, their efforts laid the foundations for the development of aviation.
Some of the earliest contributions to flying wings came from English Lt. John William
Dunne, between 1907 and 1914. He started his work from a tailless glider and followed it up
by a series of powered bi-planes. Even at this early stage of development, he had realized
the advantage of wing sweep to increase the effective tail length. He also incorporated
wash out or twist at the wing tips to counteract the premature tip stall characteristics.
His D.5 flying wing biplane, depicted in Fig. 2.3, was perhaps the first tailless aircraft to
display inherent longitudinal stability.
Figure 2.3: Dunne’s D.8 flying wing biplane - 1912 [52].
It was not until the deep-chord monoplane wing became less experimental, after World
War I, that the opportunity to discard any form of fuselage arose and extensive studies
concerning the true flying wing took place between the 1930s and the 1940s.
Soviet designers such as Boris Ivanovich Chyeranovskii started research independently
11

38. 2. Literature review 2.1: Historical perspective
and in secret under Stalin after the 1920s. With significant breakthrough in materials and
construction methods, aircraft such as the BICh-17 (Fig. 2.4) became possible [45].
Figure 2.4: Chyeranovskii BICh-17 experimental fighter - 1934.
Several late-war German military designs were based on the flying wing concept, as a
proposed solution to extend the range of the otherwise very short-range jet engined aircraft.
Most famous contributors were Reimar and Walter Horten, often credited as the Horten
brothers, who served in the army during the World War II. They were aircraft pilots and
enthusiasts and, although they had little formal training in aeronautics, they designed some
of the most advanced aircraft of the 1940s [52]. Their extensive work on tailless airplanes
finally culminated in the design of the world’s first jet-powered flying wing, the Horten IX
Ho-229, which had its maiden flight in the year 1945.
Figure 2.5: The Horten Vc - 1941.
In the U.S., the most significant contribution toward the development of tailless aircraft
came from Jack Northrop, who became involved in the development of the cleanest possible
airplane early in his career as an aircraft designer in the late 1920s. One of his earlier
designs, the Northrop N-1M, flew in 1940, followed by the N-9M in 1942. Northrop’s
interest persisted after the war, when he proposed the concept as a design solution for
12

39. 2.1: Historical perspective 2. Literature review
long-range bombers. Such trend culminated in the piston-powered YB-35 in 1946 and its
jet-powered conversion YB-49 a year later. Unfortunately a series of technical problems
and a fatal crash during landing of the YB-49 doomed the future of that design, which was
eventually discarded in favor of more conventional solutions like the Convair B-36 and the
Boeing B-52 and never entered production.
Figure 2.6: The Northrop-Grumman B-2 “Spirit” - 1989.
In the mid-1970s, the search for a new U.S. strategic bomber to replace the Stratofortress
was underway, to no avail. Besides it was becoming clear that the best way to avoid
missiles and intercepts was the adoption of low detection measures, today known as stealth
technology. The increasing importance of stealth design features, together with the advent
of fly-by-wire technology, eligible for alleviating the stability and control flaws of all-wing
aircrafts, boosted again the interest in flying wing configuration in the 1980s. The approach
eventually led the most famous, as well as the only successful, flying wing of all times, the
Northrop B-2, shown in Fig. 2.6. By virtue of its advanced flight controls systems the B-2
shows Level 1 flying qualities throughout its flight envelope [9].
Since the 1990s, a peculiar type of tailless aircraft has emerged. Defined as the blended
wing body, or simply BWB, it features a flattened and airfoil-shaped body, which produces
most of the lift, the wings contributing the balance. The body form is composed of distinct
and separate wing structures, though the wings are smoothly blended into the body. With
the marked increase of composite materials use in airframe structures such non-cylindrical
shapes are nowadays considered feasible.
Currently, due to the excellent performance in the slow-to-medium speed range and its
stealth capabilities, the flying wing configuration is still regarded as a practical concept for
aircraft designers and there has been continuous interest in applying it both to military
and commercial aircraft design.
13

40. 2. Literature review 2.1: Historical perspective
2.1.1 Modern stealth UCAVs
Since the early years of the 1990s, thanks to the development of ever more reliable
communications links and to the wider use of automated systems, the military acquired
much more confidence with the concept of using uninhabited aircrafts for performing actual
combat missions. The idea was revived in the form of various designs generally designated
as Unmanned Combat Air Vehicles (UCAV).
The continuous pursue of the best achievable performance, supported by a robust
competence in the aforesaid technologies, has produced a series of remarkable aircrafts,
result of the inevitable synthesis of flying wing, stealth and UCAV technologies.
UCAVs missions would be conducted by an operator in a ground vehicle, warship, or control
aircraft over a high speed digital data link. Even so, the operator would fly the UCAV
(a) Boeing X-47 “Pegasus”. (b) Lockheed-Martin RQ-170 “Sentinel”.
(c) BAE Systems Taranis. (d) BAE Systems Corax.
(e) AVIC 601-s “Lijan”. (f) Dassault nEUROn.
Figure 2.7: Modern stealth flying wing UCAV designs.
14

41. 2.2: An overview on flight mechanics analysis 2. Literature review
with a merely supervisory role, rather than as an actual pilot. The robot would, in fact,
be able to handle the details of flight operations and complete its mission autonomously,
if communications were cut.
So far the U.S. have played a worldwide leading role in the development of UCAV
platforms of the first generation and they are, to all effects, laying the foundations for
the second one, which will likely correspond to the sixth of fighter aircrafts altogether [34].
However Europe and other countries are actively endeavoring to bridging the gap and their
efforts are finally paying off, even though most of the configurations are just technology
demonstrators and research prototypes.
The best representatives of this new breed of flying machines are depicted in Fig. 2.7.
2.2 An overview on flight mechanics analysis
A discussion on the underlying principles and equations that govern both static and
dynamic stability of an aircraft, as well as the estimation of its flying qualities is addressed
[9]. An appreciation of these aspects is doubly important. At the early design stages,
they lead the engineer to shape a design capable to generate adequate lift and control
forces and inherently stable. At more advanced design stages, they are considered when
the compliance of the aircraft with regulation requirements is tested.
It is important to point out that all the concepts exposed in this section are the result
of linear analysis and, as such, based on the assumption of linearity of the aerodynamic
coefficients and, ultimately, of the aircraft model. The latter, along with the linearization
procedure used to derive it, is described in Appendix C.
2.2.1 Static stability
The motion of an airplane can usually be broken into two parts: the first is the lon-
gitudinal or symmetric portion, which consists of motions inside the xz plane, with the
wings always leveled; the second is the lateral-directional portion, which consists of rolling,
yawing and sideslipping, at constant elevation angle. Such separation can be applied to
both static and dynamic analyses. However the results of greater importance in the context
of static analysis are those associated with the longitudinal portion of the aircraft motion
[2]. Hence the principles reviewed in the present section will be limited to longitudinal
stability, it being understood that the same approach is applicable in a similar fashion to
directional stability analysis.
The stability of a generic system is defined as its tendency to recover to the initial
condition after a disturbance without any external input. In aeronautical terms, longitu-
dinal static stability involves the generation of a restoring (nose-down) pitching moment
in response of an increase in the angle of attack, without any control action from the pilot.
Moreover, a steady flight condition is defined balanced, if the resultant force and moment
15

42. 2. Literature review 2.2: An overview on flight mechanics analysis
about the center of gravity are both zero, that is the aircraft is in equilibrium. In particular,
this requires the pitching moment to be zero.
(a) Balanced aircraft. (b) Unbalanced aircraft.
Figure 2.8: Pitching moment curves (fixed elevator) [2].
Thus, static analysis suggests that, for an aircraft to be statically stable in pitch, the
variation in pitching moment with angle of attack must be negative; then, for an equilibrium
condition to exist, the pitching moment at zero angle of attack must be positive [1].
Cmα ă 0 (2.1a)
Cm0 ą 0 (2.1b)
The derivative Cmα is occasionally called pitch stiffness, as it models a spring-like behavior
of the aircraft in the pitch axis. i.e. Figure 2.8 shows all the possible graphs of the pitching
moment coefficient Cm versus the angle of attack α, measured from the zero-lift line of the
aircraft. It is clear that a design can be considered practical only if its pitching moment
curve can be traced back to one those of Fig. 2.8a, i.e. if the signs of Cmα and Cm0 are
opposite, otherwise a trim condition is not guaranteed. In fact, an unstable aircraft can
be equipped with an appropriate FCS to stabilize its response, while it is never possible to
fly an aircraft that cannot be balanced. In other words, it is not the stability requirement,
taken by itself, that restricts the possible configurations, but rather the requirement that
the airplane must be simultaneously balanced and stable [2]. i.e. A positively cambered
airfoil exhibits a moment about its aerodynamic center always negative within the normal
range of angle of attack. Thus, in a conventional configuration, the value of Cm0 is made
positive by the contribution of an auxiliary surface, conveniently set with a slight negative
incidence.
The same surface also provides most of the negative component of Cmα , given that the
lift it generates, despite being modest in comparison to that of the wing, possesses a much
longer lever arm. A sketch of the standard arrangement is offered in Fig. 2.9.
16

43. 2.2: An overview on flight mechanics analysis 2. Literature review
Figure 2.9: Conventional wing-tail arrangement [2].
The identical argument, applied to the case of a canard, i.e. tail-first, arrangement, leads
to the conclusion that the auxiliary surface (the canard) must be set at a slightly positive
incidence. It is worth to point out that canards has the virtue of producing lift directed
consistently with that of the wing, thus alleviating its load, as opposite to the stabilizer.
A mathematical analysis of the longitudinal static stability of a complete standard aircraft
yields the position of the point at which the resultant lift is applied, called neutral point.
Since the pitching moment of an isolated surface about its AC can be safely considered
invariant with α, it follows that the resultant aerodynamic moment of the aircraft about
the that very point is constant with α. On this basis, it is possible to express the variation
in pitching moment due to changes in α, as:
Cmα “ ´CLα pkn ´ kq (2.2)
where the term in brackets denotes the dimensionless distance of the NP from the CG,
positive for CG fore of the NP. It follows that the neutral point corresponds to the AC of
the complete aircraft, that is to the position of CG at which Cmα is zero and static stability
is neutral. Thence the name. The larger the surface and the moment arm of the tail, the
further aft moves the neutral point. For example, the neutral point of the configuration
depicted in Fig. 2.9 would lie somewhere aft of the wing AC.
The term in brackets in (2.2) is called static margin K, usually quoted in percentage
of the mean aerodynamic chord, and it quantifies the margin of movement of the CG prior
to reach the stability limit. At first analysis, the SM is a measure of the static stability of
the airplane with respect to α disturbances [21].
It can be stated, the proof given in plenty of literature, that the pitch stiffness can be made
negative for virtually any combination of lifting surfaces and bodies by placing the center
of gravity far enough forward of the neutral point [2].
If the CG is behind the neutral point, the aircraft is longitudinally unstable (K ă 0),
and active inputs to the control surfaces are required to maintain steady flight. Though,
the trade-off of reduced stability is an increase in responsiveness to commands, i.e. an
improvement in maneuverability, a concept antithetical to stability. Indeed, an aircraft
with a large static margin is very stable, but also sluggish to respond to commands and
more prone to saturate the controls, due to their reduced effectiveness.
17

44. 2. Literature review 2.2: An overview on flight mechanics analysis
The value of the SM is of critical importance in the design of an aircraft, not only
because it represents the main indicator of the stability of the design, but also because,
ultimately, it determines the controllability and handling qualities of the vehicle.
2.2.2 Dynamic stability
The evaluation of static stability only provides a description of the reaction of the
aircraft immediately following a disturbance. This result, as crucial as it is, is not sufficient
to ascertain how the airplane will actually behave in time after a perturbation in steady
flight. The study of the dynamic response of the aircraft is of great relevance, especially
in the evaluation of flying qualities, as it defines its handling characteristics and measures
the level of ease and comfort with which it can be flown.
Figure 2.10: Dynamic response of a statically stable aircraft [52].
In general, static stability is a necessary, but not sufficient condition for the dynamic
stability of a system. The dynamic response of a system, as the static one, can be either
stable, neutral or unstable, depending on the evolution of the amplitude of its response.
Static stability analysis provides some useful, but rather crude measure of the airplane
dynamics, in the sense that neutral or negative static stability always implies dynamic
stability of the same type; while positive static stability admits any type of dynamic
behavior. This last case is clearly outlined in Fig. 2.10.
The response to a disturbance can be derived from the linearized six degrees of freedom
equations of motion of the aircraft. The approach is based on the method of representing
the aerodynamic forces and moments by means of stability coefficients, first introduced by
George H. Bryan in 1911. The technique assumes that the aerodynamic actions can be
expressed as a function of the instantaneous values of the perturbation variables [4]. Using
a first order Taylor series expansion, the approach finally leads to a set of linear differential
equations with constant coefficients, which in normal form reduces to:
9x “ A x (2.3)
By virtue of the already-mentioned decoupling between symmetric and asymmetric
motions, the problem can be broken into two distinct, easier to solve sets of differential
18

45. 2.2: An overview on flight mechanics analysis 2. Literature review
equations, namely longitudinal and lateral-directional dynamics.
Moreover, a useful facilitation is represented by the fact that, when applying eigenanalysis
to linear models in the form (2.3), the solution comes in the form of natural modes, which
decouple the response of the aircraft into a set of simpler motions, each dominated by a
limited number of states. In particular, the solution comes in the form:
xptq “ x0 eλ t
(2.4)
where λ is one of the eigenvalues or poles of the system and x0 is the eigenvector that
describes the associated natural mode.
Natural modes can be fully characterized by their frequency and damping ratio, which,
in turn, are determined by the value of the associated eigenvalue, in general a couple of
complex conjugate poles. In addition, information about relative amplitude and phase
shifting between the state variables associated to each mode, i.e. their dynamic behavior,
are incorporated in the eigenvectors.
Since damping ratio quantifies the time trend of the amplitude of the response, it is
the definitive parameter to assess the dynamic stability of an aircraft, or, more precisely,
of each distinct mode. Moreover, by applying the analysis to different flight conditions,
the tool provides a reliable prediction of the modification of dynamic stability properties
over the whole flight envelope.
The typical modes of motion of a conventional aircraft are listed below.
• Longitudinal modes
1) Phugoid: it can be described as a lazy interchange of kinetic energy and poten-
tial energy about the equilibrium flight condition. The motion has low damping
and very long period. It is usually easily manageable by the pilot.
2) Short period: it is a heavily damped pitch oscillation, with a very short period
and a time to half of the order of 1 s. Speed does not have time to change
significantly, hence it involves essentially an angle of attack variation.
• Lateral-directional modes
1) Roll subsidence: it consists of almost pure rolling motion and it is generally
non-oscillatory. It expresses the damping of rolling motion.
2) Spiral: it is a non-oscillatory motion, consisting of a slow turn with sideslip. It
is unstable in conditions of reduced dihedral effect and high directional stability.
3) Dutch roll: is a coupled roll and yaw motion, with a period of 3˜15 s, often not
sufficiently damped for good handling, especially in aircrafts with high dihedral.
19

46. 2. Literature review 2.2: An overview on flight mechanics analysis
2.2.3 Flying and handling qualities
Handling qualities are those characteristics of a flight vehicle that govern the ease and
precision with which a pilot is able to perform a flying task. They have a critical bearing
on the safety of flight and on the ease of controlling an airplane in steady flight and in
maneuvers. The way in which particular vehicle factors affect handling qualities has been
a matter of study in aviation for decades.
The problem of a preliminary and trustworthy estimation of the flying characteristics
and ease of operation of an aircraft arose since the first flights. Due to the increasing
frequency of aircraft crashes in the early twentieth century, aeronautical engineers became
aware of the primary importance of a design aimed at achieving specific handling qualities,
as well as adequate stability characteristics (two often antithetical concepts) [13].
Today, flying and handling qualities play a significant and necessary role in the design
of both civil and military, piloted and autonomous airplanes. In order to ensure the
accomplishment of the desired mission safely and successfully with the minimum amount
of workload for the pilot, the aircraft, whether it is augmented or not, must satisfy the
corresponding regulation.
Yet, what constitutes acceptable characteristics is often not obvious, and several at-
tempts have been made to quantify pilot opinion on acceptable handling qualities. Refer-
ence standards for the handling qualities of any category of air vehicle have been developed
and are now in common use [27].
Subjective flying qualities evaluations such as Cooper-Harper ratings are used to distinguish
between “good-flying” and “difficult-to-fly” aircraft. Moreover, quite useful and reliable fly-
ing qualities estimates may be provided on the basis of various dynamic characteristics, by
correlating pilot ratings to the frequencies and damping ratios of the aircraft’s modes of
motion, as in done in the U.S. military specifications. These standards essentially define a
subset of the dynamics and control design space that provides good handling qualities for
a given combination of aircraft type and flying task [6].
Nowadays new aircraft designs can be simulated way before actual flight testing to assess
their airworthiness. Nevertheless, such real-time, pilot-in-the-loop simulations are expen-
sive and require a great deal of information about the aircraft, which are not likely to be
available at early stages of design.
2.2.3.1 Cooper-Harper rating scale
The Cooper-Harper rating scale is a set of criteria formalized in the late 1960s and
ever since used by test pilots and engineers to evaluate the handling qualities of aircraft
during flight test. The scale ranges from 1 to 10, with 1 indicating the most desirable
handling characteristics and 10 the worst. The criteria are evaluative and, thus, the scale
is considered subjective.
It is important to note that a Handling Qualities Rating (HQR) can only be assigned
20

47. 2.2: An overview on flight mechanics analysis 2. Literature review
to a well defined combination of a repeatable task, a well trained pilot, that is actively
engaged in accomplishing that task, and a specific aircraft.
Figure 2.11: Cooper-Harper rating scale [52].
The scale cannot be applied straightforwardly for the purpose of evaluating the flying
qualities of an unmanned aircraft, for the very reason that it is based on the “sensations”
of a pilot physically located onboard the vehicle. Even though, it is arguable that it might
be adopted in the case of remotely piloted UAV. In that scenario a pilot is actually present
and his perceptions, however limited compared to those of a conventional pilot, could be,
with due caution, taken into account.
Recently an alternative version of the Cooper-Harper scale has been proposed by Cum-
mings, et al. [33]. Since in UAV operations displays are often the only information link
between operators and vehicles, a quasi-subjective display evaluation tool called the Modi-
fied Cooper-Harper for Unmanned Vehicle Displays (MCH-UVD) has been developed. The
tool, adapted from the Cooper-Harper aircraft handling scale, allows operators to evaluate
a display, rather than the dynamic behavior of the aircraft directly, by translating their
judgments on potential display shortcomings into a number corresponding to a particular
deficiency in operator support.
21

48. 2. Literature review 2.2: An overview on flight mechanics analysis
Figure 2.12: MCH-UVD diagnosis tool [33].
The intent of the redesign was to represent a severity scale that defines the ability to
complete the mission. Like the original Cooper-Harper scale that rated aircraft control-
lability on a scale of severity, the intent was to scale severity that reflected the UVD’s
ability to support safe mission completion. At the same time, the intent was to maintain
the concepts of the human information processing model within this new scale, as this is
a critical component to UV display designs [33].
2.2.3.2 MIL-HDBK-1797A
The first comprehensive military handling qualities specifications were issued in the
early 1940s by the Navy Bureau of Aeronautics and the U.S. Army Air Force (AAF-C-
1815), in acknowledgement of the demand of the military services of a unified standard,
less subjective than the Cooper-Harper scale and based on quantifiable parameters. More
importantly, the subsequent version MIL-F-8785B of 1954, began the precedence within
the handling qualities community that the true value in a specification document was
an elaborate Background Information and Users Guide (BIUG), wherein the data which
form the specification are contained, rather than the detailed requirements per se. The
BIUG forms the historical lessons-learned for handling qualities which provide a continual
improvement process for air vehicle handling qualities.
The use of military specifications fell out of favor in the 1980s. The last in this series
22

49. 2.2: An overview on flight mechanics analysis 2. Literature review
was MIL-F-8785C issued in 1980. MIL-F-8785C was then re-worked and updated into a
military standard (MIL-STD-1797A) in 1995, which was further re-designated in 1997 as
a handbook, the MIL-HDBK-1797A.
This latter specification is intended to assure flying qualities that provide adequate
mission performance and flight safety regardless of design implementation or flight control
system mechanization (although it primarily focuses on unaugmented piloted aircrafts).
The structure of the specification allows its use to guide these aspects in design, construc-
tion, testing and acceptance of the subject aircraft [35].
Under MIL-HDBK-1797A three levels of acceptability of the flight characteristics, re-
lated to the ability to complete the operational missions for which the airplane is designed,
are defined. These levels are presented in Tab. 2.1.
Level Degree Definition HQR
1 Satisfactory
Flying qualities clearly adequate for the mission
flight phase. Desired performance is achievable
with no more than minimal pilot compensation
ě 3.5
2 Adequate
Flying qualities adequate to accomplish the mis-
sion flight phase, but some increase in pilot work-
load or degradation in mission effectiveness, or
both, exists
ě 6.5
3 Controllable
Flying qualities such that the aircraft can be con-
trolled in the context of the mission flight phase,
even though pilot workload is excessive or mission
effectiveness is inadequate, or both
ě 9.5
Table 2.1: Definition of handling quality levels in MIL-HDBK-1797A [35].
For the purpose of handling qualities evaluation an aircraft is placed in one of the
following classes:
Class I small light aircraft;
Class II medium weight, low-to-medium maneuverability aircraft;
Class III large, heavy, low-to-medium maneuverability aircraft;
Class IV high-maneuverability aircraft.
The specification introduces a further subdivision of the analysis depending on the
flight phase, based on the experience with aircraft operations that certain flight phases
23

50. 2. Literature review 2.2: An overview on flight mechanics analysis
require more stringent values of flying qualities parameters than others do. A description
of the different flight phases defined within MIL-HDBK-1797A is summarized in Tab. 2.2.
Type Category Definition
Nonterminal
A
Phases that require rapid maneuvering, precision
tracking, or precise flight–path control
B
Phases that are normally accomplished using
gradual maneuvers and without precision track-
ing, although accurate light–path control may be
required
Terminal C
Phases normally accomplished using gradual ma-
neuvers and usually require accurate flight–path
control
Table 2.2: Definition of flight phase categories in MIL-HDBK-1797A [35].
The specification provides a comprehensive assortment of requirements, spanning all
modes of motion of a conventional airplane, that specify the limits of acceptability to be
met by the aircraft under study, according on the flight phase.
In terms of longitudinal modes, acceptable limits on the stability of the short period,
which defines the longitudinal control dynamic, are quantified by the range of damping
ratio for each flight phase categories and quality levels, as Tab. 2.3 shows.
Level
Category
A, C B
1 0.35 ď ζ ď 1.30 0.30 ď ζ ď 2.00
2 0.25 ď ζ ď 2.00 0.20 ď ζ ď 2.00
3 ζ ě 0.15 ζ ě 0.15
Table 2.3: Short period requirements in MIL-HDBK-1797A [35].
One can observe that the range constraints summarized in Tab. 2.3 identify a relatively
wide region of acceptability, relaxing the work of aircraft designers. Cook demonstrated
that the ideal damping ratio of SPO mode is 0.7, a value that ensure satisfactory margin of
stability, while minimizing the settling time after a disturbance [3]. Indeed, a value bigger
24

51. 2.2: An overview on flight mechanics analysis 2. Literature review
than 1.0, indicating an overdamped system, would imply a generally longer settling time.
The quality level of phugoid mode for all phases is characterized by its damping ratio,
as shown in Tab. 2.4 [35]. Note that such requirement applies with both free and fixed
pitch control.
Level All categories
1 ζ ě 0.04
2 ζ ě 0
3
unstable,
T2 ě 55 s
Table 2.4: Phugoid requirements in MIL-HDBK-1797A [35].
The requirements for the phugoid mode, compared to those of the SPO, are clearly relaxed,
because of the longer period, which leaves the pilot plenty of time to act.
Furthermore, it can be stated that a phugoid frequency approximately one tenth of that
of the SPO represent an ideal value [10].
The performance of the roll subsidence mode is evaluated by means of its time constant
τR, expressed in seconds, according to Tab. 2.5.
Category Class
Level (τR min)
1 2 3
A
I, IV 1.0 s 1.4 s
10 s
II, III 1.4 s 3.0 s
B all 1.4 s 3.0 s
C
I, II-C, IV 1.0 s 1.4 s
II-L, III 1.4 s 3.0 s
Table 2.5: Roll subsidence requirements in MIL-HDBK-1797A [35].
The requirements on spiral stability are aimed primarily at insuring that the aircraft
will not diverge too rapidly in bank from a wings level condition during periods of pilot
inattention. The criterion is formulated according to the requirement that, following a
disturbance in bank of up to 20 degrees, the time for the bank angle to double amplitude
25

52. 2. Literature review 2.2: An overview on flight mechanics analysis
shall be greater than the values reported in Tab. 2.6. This requirement must be met with
the aircraft trimmed in symmetric leveled flight and with free cockpit controls [35].
Category
Level (T2 min)
1 2 3
A, C 12 s
8 s 4 s
B 20 s
Table 2.6: Spiral requirements in MIL-HDBK-1797A [35].
Finally, the requirement specified for the dutch roll are aimed at attaining a sufficiently
stable and well damped lateral-directional oscillatory dynamic.
Level Category Class
Requirements (minimum value)
ζ [-] ζ ω [rad/s] ω [rad/s]
1
A (CO, GA,
RR, TF, RC,
FF, AS)2
all 0.4 0.4 1.0
A
I, IV 0.19 0.35 1.0
II, III 0.19 0.35 0.4
B all 0.08 0.15 0.4
C
I, II-C, IV 0.08 0.15 1.0
II-L, III 0.08 0.10 0.4
2 all all 0.02 0.05 0.4
3 all all 0 - 0.4
Table 2.7: Dutch roll requirements in MIL-HDBK-1797A [35].
The requirements are a bit more complex than those seen so far, due to the strong coupling
2
Indicating respectively: air-to-air COmbat, Ground Attack, in-flight Refueling (Receiver), Terrain
Following, ReConnaissance, close Formation Flying, Antisubmarine Search
26

53. 2.2: An overview on flight mechanics analysis 2. Literature review
of the mode and the key importance of its adequate controllability.
It is worth to point out that longitudinal requirements were empirically derived from
pilot comment. Specifically, they were established using criteria based on a human opera-
tor’s ability to act as the aircraft’s augmentation and control system.
Moreover, given that the primary guide to determine these values was pilot input on
unaugmented aircraft, their applicability to autonomous UAVs is limited to those that
are designed to match piloted aircraft dynamics for landing purposes or gust rejection,
irrespective of the UAV’s control system [13].
2.2.3.3 CAP criterion
The Control Anticipation Parameter (CAP), introduced by Bihrle in 1965, is one of
the earliest and most diffused flying qualities criteria, especially for unaugmented piloted
aircrafts. The CAP is defined as the ratio of the aircraft’s pitch acceleration to change in
steady state load factor. and it is used to correlate the sensitivity of the human vestibular
organ to pitch acceleration to a sensed g-loading of an aircraft.
CAP can be expressed as ratio of short period natural frequency ωSP and normal
acceleration derivative w.r.t. angle of attack Nα (see Appendix C), or equivalently as ratio
of instantaneous pitch acceleration and steady state normal acceleration [3]:
CAP “
:θ
∆ss
“
9qp0q
Nzp8q
«
ωSP
2
Nα
(2.5)
where
Nα “ ´
Zw u0
g
ωSP “
b
Mq Zw ´ Mw
`
Zq ` u0
˘
This expression gave rise to the short period frequency requirements found in the
military specification handbook, which are summarized graphically in charts such as the
one found in Fig. 2.13 (relative to Category B flight phase). CAP can then be evaluated
graphically using the parameters in (2.5), which, in turn, can be derived from the reduced
second order model for the short period mode.
In conclusion, it is important to notice that the application of flying qualities analysis
to UAVs presents a unique problem: the absence of a human pilot. Since flying quali-
ties analysis traditionally focuses on pilot opinion, a major component of flying qualities
analysis must be rethought. There is however, a large body of work of piloted criteria
that does provide guidance for the application of flying qualities analysis to UAVs. Any
application of a criterion to UAVs must follow some of the basic tenants used in piloted
analysis. In general, it must be simple enough to use, effective enough to make worth
using, and familiar enough so that there is a feeling of intuitive comfort in using it [13].
27

54. 2. Literature review 2.3: Flying wing design issues
Figure 2.13: CAP requirements for Category B flight phase [35].
2.3 Flying wing design issues
A flying wing is a tailless fixed-wing aircraft that has no definite fuselage, with most of
the crew, payload, and equipment being housed inside the main wing structure. It is often
regarded as, theoretically, the most aerodynamically efficient configuration for a fixed wing
aircraft. It also would offer high structural efficiency for a given wing thickness, leading
to light weight and high fuel efficiency. Because the airframe lacks conventional stabilizing
and/or associated control surfaces, in its purest form the flying wing can easily suffer from
the inherent disadvantages of being unstable and difficult to control.
28

55. 2.3: Flying wing design issues 2. Literature review
2.3.1 Longitudinal issues
The first challenge in the design of a tailless aircraft consists in obtaining a configura-
tion, if not stable, at least balanced, that is ending up with a pitching moment curve of
the type of Fig. 2.8a.
For straight winged tailless airplanes, the lack of horizontal tail makes the compliance
with the requirement expressed by (2.1b) only possible with the adoption of a reflex airfoil,
as done by the Horten brothers. Effectively, the same result is attained if a flap, deflected
upward, is incorporated at the trailing edge of a symmetrical airfoil [2]. However, this
solution, in both of its forms, comes with a reduction in maximum achievable lift, along
with a sensible increase in drag, together with the limited CG range, hence straight wing
flying wing configuration is seldom adopted.
The only feasible alternative for all-wing airplanes is the swept-back wing with twisted
tips (washout): when the net lift is zero, the forward part of the wing produces a positive
contribution and the rear part a negative one for a resulting positive couple [2]. However,
swept wings, especially if untapered (like that of the SACCON), tend to be subject to tip
stall, due to the high suction peaks on the leading edge in the proximity of the outer wing
sections, caused to trailing vorticity in the wake of the inboard wing sections [21]. If, on the
one hand, the advantage of such a behavior is a more progressive stall, its drawbacks are
represented by a de-stabilizing nose-up pitching moment, caused by the forward movement
of the center of pressure, the loss of aileron effectiveness and the risk of asymmetric stall,
leading to undesirable roll tendency. The use of tip slats was advocated by Donlan [28] as
being the most effective method for delaying the tip stall, as they may increase the angle
of stall as much as 10˝, if judiciously located. In addition, slats can also be employed to
adjust the Cm0 of the configuration, although at the price of increased drag.
Besides that, another typical phenomenon of highly swept wings is the development of
complex vortical flows on the upper surface of the wing. Such flow topology guarantees
lift up to higher angle of attack than straight wings, at a price of a reduced lift slope, but
it is responsible for undesirable behavior of the pitching moment as well, including sudden
dips and non-linearities concurrently with vortex breakdown dynamics [12].
Referring to the value of Cmα , as given by (2.2), it is clear that the position of the
NP of a tailless aircraft coincides with that of the AC of the wing. Thus the only possible
way to achieve negative pitch stiffness, i.e. positive static stability, is to locate the CG
ahead of the wing AC. Jones reports that an extreme reduction of thickness toward the
trailing edge may cause a backward displacement of the AC of 2 or 3 percent [26]. In
any case, this severely restricts the allowable CG range in comparison to a conventional
configuration. In particular, Donlan [28] suggests an optimal static margin range from 2%
to 8%, mainly due to limitations in control power for such aircrafts. Often the only feasible
way to provide stability to a tailless design is artificially, by means of suitable stability
augmentation systems based on modern fly-by-wire technology. In this regard, already
29

56. 2. Literature review 2.3: Flying wing design issues
in 1941 Northrop [25] advised that an intentionally unstable configuration augmented by
reliable and sophisticated fly-by-wire control system represents the best solution for a flying
wing, especially if sizable.
A second preeminent issue is represented by pitch damping, denoted by Cmq . For
conventional airplanes most of the contribution to pitch damping (actually nearly 90% of it)
comes from the horizontal stabilizer, the effect of the fuselage being negligible. Nevertheless
Jones [26] points out that, if the airplane is statically stable (AC aft the CG), the free
rotation in pitch couples with motions normal to the chord and the damping of such motions
is effective in contrasting the pitching. In fact, the lack of direct damping appears to alter
the sequence of the motions in such a way as to make this coupling more effective in the
case of tailless configurations. Remarkably, despite a rotary damping coefficient Cmq just
one tenth that of conventional aircrafts, the actual capability of tailless aircrafts to damp
pitch oscillations in flight is nearly as great. Northrop [25] further asserts that, despite the
low pitch damping, SPO results well damped due to the plunge damping parameter CZw ,
that absorbs most of the energy of oscillation.
Furthermore, as explained by Donlan [28], a relaxed or negative static margin may lead
to the development of an uncontrolled dynamic, called tumbling. It is a divergence motion
consisting of a continuous pitching rotation, capable of rendering conventional control
surfaces almost useless, once it is initiated. Indeed, tumbling was deemed responsible for
the accident that claimed the lives of Captain Glen Edwards and other four crew members,
during a low altitude stall test on board of the Northrop YB-49. According to Donlan, to
avoid tumbling dynamic, the static margin should never be permitted under any condition
to become negative. Nevertheless, it has been argued that a non-negative static stability,
might not be a guarantee against this phenomenon [9].
To exert the necessary control action, a tailless aircraft can only rely on large elevons
fitted at the trailing edge of the wing. Then, with the exception of delta wing designs, the
longitudinal distance between the control surface and the CG will be considerably smaller
than in a conventional aircraft. As a consequence, for the same static margin, the elevator
of a tailless aircraft will prove much less effective than that of a conventional configuration,
also implying larger deflections. Poor longitudinal control authority may become critical
during take-off, as the aircraft may not be able to generate a strong enough rotation
moment to overcome the combined action of the nose-down moment of its own weight
about the point of ground contact and that created by friction on wheels.
2.3.2 Lateral-directional issues
Despite possessing no direct stiffness in roll (Clϕ = 0), stable airplanes exhibit an in-
herent tendency to fly with leveled wings, called dihedral effect or, less frequently, roll
stiffness. The phenomenon is the consequence of the interaction of gravity with the deriva-
tive Clβ
and arises whenever a lateral velocity, thus sideslip β, is established due to an
30

57. 2.3: Flying wing design issues 2. Literature review
unbalanced lateral weight component. The value of Clβ
, whether positive or negative, is
generally kept small, to avoid excessive roll-yaw coupling (primary cause of undesirable
dutch roll characteristics, if negative) and lateral oscillations in rough air [26]. Since dihe-
dral effects originates from the configuration of the wing alone, the lateral static stability
for flying wing aircrafts is not much different from that of a conventional configuration and
satisfactory roll stiffness can be achieved using standard design practice.
The same principle applies to roll damping Clp , which is determined by the wing.
Lateral control is the only aspect that presents no apparent difficulty. It is achieved by
means of conventional ailerons or spoilers controls, placed at an appropriate span location.
The main difficulty in the design of a flying wing lies undoubtedly in the provision
of sufficient weathercock stability and yaw damping. Owing to the absence of any form
of vertical stabilizer, a flying wing shows poor directional stability, despite the beneficial
effect of the lack of fuselage. Wind tunnel tests showed that attainable values of Cnβ
for
a flying wing never exceed the 33% of those of a tailed design. Jones [26] states that the
fin surface necessary to realize the required degree of weathercock stability can be greatly
reduced by fitting lateral controls with zero or even favorable yaw coupling. Nevertheless,
Rahman [9] asserts that small fins fitted at the wing tips do not provide a valid solution,
especially for large configurations, and should in any case be avoided, on penalty of an
increase in drag and weight (potentially negating the advantage of the flying wing).
Jones [26] further suggests that the use of sweepback planform combined with downward
cranked (anhedral) wing tips could secure adequate directional stability, particularly at
high speed and angle of attack. The solution was actually implemented by Norhrop in its
N-1M, as it is clearly visible in Fig. 2.14, although he later abandoned it.
Figure 2.14: Northrop N-1M.
Another practical solution to increase directional stability of tailless aircrafts, advised by
Prandtl and later the Horten brothers, is represented by the proper use of wing twist
and airfoil sections, in order to establish a bell-shaped lift distribution, together with a
sweepback planform. Apparently this induces the wing tips to generate a forward-oriented
lift vector, that will effectively pull the trailing wing forward.
As in the case of pitching motion, the elimination of the tail has a negative impact on
yaw damping capability of the vehicle. In fact the damping action of the wing, caused by
31

58. 2. Literature review 2.3: Flying wing design issues
the distribution of drag along the span, has a marginal effect and it is subordinate to the
value of the local lift coefficient. In this case the adoption of fin-like surfaces at the tip
of the wing [26] produces marginal effects, quantified in Reference [26] in a Cnr derivative
roughly 10% of that of a conventional design.
The type of surface usually employed to control yaw is the rudder, which cannot be
fitted in a flying design. Northrop tested several rudder-like controls designs, the most
successful of which was the split aileron, or drag rudder, sometimes also referred to as
deceleron. The functioning is based on differential drag produced by the surface and by
virtue of their long moment arm from the CG (approximately half span).
With the implementation of a flight control system, integrated with adequate sensors, it
is possible to obtain an aircraft with directional flying and handling qualities as good as
those of a standard configuration.
Figure 2.15: The drag rudder deployed on the wing tip of the Northrop N-9M.
Moreover, if mounted far enough from the CG, the drag required to exert sufficient yawing
moment will be tolerable, making split aileron the most effective and ingenious method for
providing both directional stiffness and damping to a flying wing. Still, the use of drag
as a mean of control makes the design more suited for steady cruising in still air, while it
becomes less efficient when maneuvering or in turbulent air.
Finally, Donlan [28] infers that the thrust line should be kept as close as possible to the
centerline, so as to minimize the control power, i.e. additional drag, required in asymmetric
thrust conditions (engine failure).
32

59. Chapter 3
Aerodynamic database
The first subject that had to be addressed was the generation of the aerodynamic
database of the SACCON concept. The database constitutes the actual aerodynamic model
of the configuration under study as it will be used to estimate the S&C characteristics of
the final aircraft. The main aspects covered in the chapter are the collection of the wind
tunnel measurements, their processing and some remarks following their analysis. Finally
the evaluation of control effectiveness is discussed, along with some remarks concerning
the limitations of the configuration.
3.1 Foreword
In modeling an aircraft’s aerodynamic database the choice of the more suitable mathe-
matical structure is often a crucial challenge. This is mainly due to the complex, non-linear
functional dependencies of forces and moments on both present and past values of several
flow and control parameters. Addressing the problem remains difficult even after the de-
coupling of symmetric (longitudinal) and asymmetric (lateral) dynamics, motivated by the
symmetry characteristics of (most of) aircrafts, greatly reduced the intricacy of the former
issue. A reasonable simplification is that the airplane mass and inertia are significantly
larger than those of the surrounding fluid. Also the flow is often considered quasi-steady,
which implies that steady-state aerodynamic conditions are reached instantaneously after
a disturbance, de facto neglecting the memory effect of the flow field. In general, the static
dependencies of the aerodynamic coefficients upon steady parameters, chiefly incidence
angles and Mach number, constitute the so-called baseline database of the model, which
provides a fundamental overview of the aerodynamic loads throughout the flight envelope
of interest. Furthermore, it has the potential to represent non-linear phenomena such as
static stall, compressibility effects and the onset and breakdown of vortical flows [11]. Inte-
grating the baseline database with the portion accounting for the effects of control surfaces
deflections, yields the complete static model of the aircraft. Moreover, the data contained
in such a model can be directly measured from static wind tunnel tests.
33

60. 3. Aerodynamic database 3.1: Foreword
A quasi-steady model can be effectively employed to investigate the S&C characteristics
of the aircraft through simulation of flight maneuvers. On this regard, the results obtained
by Da Ronch et al. [29] indicate a fairly good agreement between time-accurate and quasi-
steady solutions, as long as the maneuver involves moderate angular rates and low angle of
attack (lazy eight). Da Ronch et al. [29] further report that such correspondence is lost as
the maneuver gets more rapid and involves higher angles of attack. In that case the onset
of non-linearities in the flow, such as vortices and separation, occurs and relevant time
history effects are triggered. Therefore the discrepancies observed between time-accurate
and quasi-steady solution were entirely ascribed to the lack of an accurate description of
lead/lag effects in vortex development.
As a matter of facts, a constantly increasing number of common interest application,
such as the SACCON, are dominated by non-linear, vortical flows, sometimes in the tran-
sonic regime and/or at high angle of attack, so much so that the assumption of quasi-steady
flows becomes narrow. In these cases the inclusion of dynamic behavior modelization, es-
pecially concerning reduced frequency effects, becomes mandatory for the attainment of
adequate correlation with time-accurate reference solutions.
With the introduction of the assumption that the influence of control surfaces deflection
on dynamic effects is negligible, the characterization of full functional dependencies of the
aerodynamic coefficients for a complete model can be broken down as:
Ci “ C0
i
´
α, β, M8
¯
looooooomooooooon
baseline
` ∆Cδ
i
´
α, β, M8, δe, δa, δr
¯
looooooooooooooomooooooooooooooon
control
` ∆Cω
i
´
α, β, M8, p, q, r
¯
loooooooooooooomoooooooooooooon
rotational
`
` ∆Ct
i
´
α, β, M8, 9α, 9β
¯
loooooooooooomoooooooooooon
unsteady
(3.1)
valid for i “ X, Z (or D, L), m, Y , l, n and where the contributions of dynamic effects and
controls are introduced as increments to the baseline values. The above decomposition fits
well with the common practice of wind tunnel experiments: forced oscillation and control
deflection data measurements are preceded by the determination of the baseline database,
which can be subtracted from the firsts, yielding the increments ∆Cω
i , ∆Cδ
i and ∆Ct
i with
all available dependencies. It is worth to note that rotational increments correspond to
quasi-steady contributions, since transients are not taken into account.
Usually, forces and moment coefficients are then tabulated as functions of the flight
states and control settings, covering the designed or expected flight envelope. The database
is formulated in a fashion that overcomes the general assumption of uncoupled longitudinal
and lateral dynamics, since every dependency and cross-effect is preserved within its data.
Da Ronch [11] points out that, if five values were to be used to provide a coarse resolution
for each state and control setting appearing in (3.1), the total number of table entries, i.e.
34