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)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.
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
3.2.1 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
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
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
3.17 Rolling and yawing moments induced by the drag rudders. . . 60
3.18 LCDP. . . 62
3.19 CnβDYN. . . 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
5.23 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
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
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
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
TBA 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
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“ 12ρ8U82 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 “ πf x{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
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
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.
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].
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
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.
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 per-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 portfash-ion. Without any informatfash-ion about mass distribution of the SACCON (neither wind tunnel model, nor full-scale), these parameters were deducted by comparison to a comparable aircraft [29].
aerody-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.
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.
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
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].
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.
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.
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.
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.
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.
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 Cmversus 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
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 “ ´CLαpkn´ 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].
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:
9
x “ A x (2.3)
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 “ x0eλ 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.
2.2.3 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.
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.
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.
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.
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].
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].
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
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 “ qp0q9 Nzp8q « ωSP 2 Nα (2.5) where Nα“ ´ Zwu0 g ωSP “ b MqZw´ 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.
Figure 2.13: CAP requirements for Category B flight phase [35].
2.3
Flying wing design issues
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
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
