SAWE Report - Cal Poly San Luis Obispo

SAWE Paper No. 3232
Category No. 10
The Vendetta
Preliminary Design Report
by
Cal Poly, San Luis Obispo Design Team
consisting of
Kolby Keiser Chris Droney (Team Leader) Nathan Schnaible
Chris Atkinson Christopher Maglio Dan Salluce
For Presentation at the
61 st Annual Conference
of
Society of Allied Weight Engineers, Inc.
Virginia Beach, Virginia 20-22 May, 2002
Society of Allied Weight Engineers
Serving the Aerospace-Shipbuilding-Land Vehicles and Allied Industries
Permission to publish this paper, in full or in part, with credit to the author and the Society may be
obtained, by request, to:
Society of Allied Weight Engineers, Inc.
P.O. Box 60024, Terminal Annex
Los Angeles, CA 90060
The Society is not responsible for statements or opinions in papers or discussions at the meeting.

Abstract
Cal Poly proudly presents the Vendetta, a supersonic bomber designed to meet the criterion
specified by the AIAA 2001/2002 Undergraduate Team Aircraft Design Request for Proposal.
The mission to be flown by the Vendetta consists of a 1,750 nautical mile radius, all of which
must be flown at Mach 1.6 at or above 50,000 feet. The aircraft must have low frontal radar
cross-section and also be capable of dropping a 9,000-pound weapons payload. The Vendetta is
being designed to replace the stealthy F-117 Nighthawk and B-2 Spirit as well as the supersonic
F-15 Eagle and B-1 Lancer. The Vendetta’s current takeoff gross weight is 125,000 pounds and
its empty weight is 57,000 pounds. Furthermore, an analysis from a low observables standpoint
has been made possible by publicly accessible RCS code. The component weight buildup of the
Vendetta has been developed using both class I and class II methodologies as well as by
assigning mass properties to an actual solid model of the aircraft. Several challenges concerning
the balance of the aircraft have been introduced by the immediate and abrupt shift in the
aerodynamic center due to the acceleration from subsonic to supersonic Mach numbers.
Solutions to this problem have been developed and will be presented in the report.
i

Table of Contents
Abstract______________________________________________________________________ i
Table of Contents _____________________________________________________________ ii
List of Figures________________________________________________________________ iv
List of Tables _______________________________________________________________ vii
Nomenclature _______________________________________________________________viii
1 Introduction______________________________________________________________ 1
2 Defining the Design Domain ________________________________________________ 7
3 Configuration ___________________________________________________________ 10
4 Stealth Considerations ____________________________________________________ 16
5 Aerodynamics ___________________________________________________________ 22
5.1 Wing Sizing ________________________________________________________ 22
5.2 Wing Planform ______________________________________________________ 23
5.3 Wing Thickness _____________________________________________________ 26
5.4 Airfoil _____________________________________________________________ 28
5.5 Lift Curve __________________________________________________________ 28
5.6 Drag_______________________________________________________________ 31
6 Propulsion ______________________________________________________________ 34
6.1 Engine Selection _____________________________________________________ 34
6.2 Inlets ______________________________________________________________ 41
6.3 S-Duct _____________________________________________________________ 46
6.4 Nozzle _____________________________________________________________ 47
7 Materials and Structure____________________________________________________ 48
8 Landing Gear ___________________________________________________________ 52
9 Weight & Balance________________________________________________________ 55
10 Stability and Control____________________________________________________ 60
11 Performance __________________________________________________________ 71
11.1 Performance Requirements_____________________________________________ 71
11.2 Specific Excess Power Requirements_____________________________________ 74
11.3 Turn Rate Requirement________________________________________________ 76
11.4 Mission Requirements ________________________________________________ 78
11.5 Takeoff & Landing ___________________________________________________ 79
11.6 Performance Summary ________________________________________________ 81
11.7 Alternate Missions ___________________________________________________ 82
12 Payload ______________________________________________________________ 84
13 Cockpit ______________________________________________________________ 86
14 Systems ______________________________________________________________ 90
14.1 Auxiliary Power Generation System _____________________________________ 90
14.2 Vehicle Management System ___________________________________________ 91
14.3 Fuel System_________________________________________________________ 92
15 Manufacturing_________________________________________________________ 94
16 Cost Analysis _________________________________________________________ 96
Appendix___________________________________________________________________ 98
Threats Chart______________________________________________________________ 98
ii

Diffuser Efficiency _________________________________________________________ 99
Literal Factor Forms ________________________________________________________ 99
Foldout 1 ________________________________________________________________ 101
Foldout 2 ________________________________________________________________ 103
The Vendetta Design Team____________________________________________________ 105
References_________________________________________________________________ 107
iii

List of Figures
Figure 1.1 - Design Mission Profile _______________________________________________ 1
Figure 1.2 - F-111 Aardvark _____________________________________________________ 3
Figure 1.3 - F-15 Strike Eagle____________________________________________________ 4
Figure 1.4 - F-117 Night Hawk___________________________________________________ 4
Figure 1.5 - B-1B Lancer _______________________________________________________ 5
Figure 1.6 - B-2 Spirit__________________________________________________________ 5
Figure 2.1 - Historical Weight Fractions ___________________________________________ 7
Figure 2.2 - Constraint Plot______________________________________________________ 9
Figure 3.1 - Nergal ___________________________________________________________ 10
Figure 3.2 - Jackhammer ______________________________________________________ 10
Figure 3.3 - Interdictor ________________________________________________________ 10
Figure 3.4 - Big Paulie ________________________________________________________ 10
Figure 3.5 - Initial Configuration ________________________________________________ 11
Figure 3.6 - Radar Return of Initial Configuration___________________________________ 12
Figure 3.7 - Second Configuration _______________________________________________ 13
Figure 3.8 - Current Configuration _______________________________________________ 14
Figure 3.9 - Inboard Layout ____________________________________________________ 14
Figure 3.10 - Inboard Layout Continued __________________________________________ 15
Figure 4.1 - Stealth Considerations_______________________________________________ 16
Figure 4.2 - RCS Model Faceting________________________________________________ 17
Figure 4.3 - Radar Cross Section at 0º Lookup Angle ________________________________ 19
Figure 4.4 - Radar Cross Sections at 15º Lookup Angle ______________________________ 20
Figure 4.5 - Radar Cross Sections for a Radial Sweep________________________________ 21
Figure 5.1 - Optimization of Wing Area and Aspect Ratio ____________________________ 22
Figure 5.2 - Effect of Wing Leading and Trailing Edge Sweep on Aircraft RCS ___________ 24
Figure 5.3 - Wing Planform ____________________________________________________ 25
Figure 5.4 - Effect of Root Chord Thickness on Wing Weight and Cross Sectional Area ____ 26
Figure 5.5 - Effect of Root Chord Thickness on Fuel Burn ____________________________ 27
Figure 5.6 - Airfoil Section at MAC______________________________________________ 28
Figure 5.7 - Airfoil Section at Tip of Trailing Edge Flap______________________________ 28
Figure 5.8 – Variation in Lift Curve Slope with Mach Number_________________________ 29
Figure 5.9 - Lift Distribution of Wing with and without Twist _________________________ 30
Figure 5.10 - Subsonic Wing Lift Curve __________________________________________ 30
Figure 5.11 - Transonic Area Distribution _________________________________________ 31
Figure 5.12 - Supersonic Area Distribution (Mach 1.6) _______________________________ 32
Figure 5.13 - Drag Build-Up at 50,000 ft, Mach 1.6, Maneuver Weight, and 5% Static Margin 33
Figure 6.1 - VAATE Goals_____________________________________________________ 38
Figure 6.2 - Thrust Curves for Altitudes from Sea Level to 70,000 ft (21,300 m)___________ 39
Figure 6.3 - Military TSFC Curves for Altitudes from Sea Level to 70,000 ft (21,300 m) ____ 40
Figure 6.4 - Engine Sizing Plot__________________________________________________ 41
Figure 6.5 - Optimum Deflection Angle for Mach 1.6 Flow ___________________________ 42
Figure 6.6 - Pressure Recovery for a Two Shock versus Three Shock Inlet _______________ 42
Figure 6.7 - Inlet Area Ratio____________________________________________________ 43
iv

Figure 6.8 - Cost Association with Inlet Shocks_____________________________________ 44
Figure 6.9 - Off Design Area Required for Engine Mass Flow _________________________ 45
Figure 6.10 - Vendetta S-Duct __________________________________________________ 46
Figure 6.11 - Diffuser Angle to the Engine Face ____________________________________ 46
Figure 7.1 - Structure Buildup for Vendetta ________________________________________ 48
Figure 7.2 - Wing Attachment Detail _____________________________________________ 49
Figure 7.3 - Empennage Structural Layout_________________________________________ 50
Figure 7.4 - V-n Diagram for Vendetta____________________________________________ 50
Figure 8.1 - Main Gear Structural Attachment Point _________________________________ 52
Figure 8.2 - Landing Gear Configuration Trade Study________________________________ 53
Figure 8.3 - Main Gear Retraction Sequence _______________________________________ 53
Figure 8.4 - Nose Gear and Main Gear Retraction Schemes ___________________________ 53
Figure 8.5 - Completed Landing Gear ____________________________________________ 54
Figure 8.6 - Vendetta with MJ-1 Lift Truck and 2000lb JDAM_________________________ 54
Figure 9.1- Principle Axes _____________________________________________________ 56
Figure 9.2 - Center of Gravity Excursion __________________________________________ 58
Figure 10.1 - Longitudinal X-Plot at Mach 0.3 _____________________________________ 61
Figure 10.2 - Horizontal Area Required for Static Stability with Cant Angle ______________ 62
Figure 10.3 - Vertical Area Required for Static Stability with Cant Angle ________________ 63
Figure 10.4 - Radar Cross Section Impact of 20° vs. 30° Vertical Cant Angle _____________ 64
Figure 10.5 - Vendetta Empennage Configuration ___________________________________ 66
Figure 10.6 - Mach Tuck Illustrated ______________________________________________ 66
Figure 10.7 - Pitch Break Characteristics __________________________________________ 68
Figure 10.8 - Pheagle Simulator _________________________________________________ 70
Figure 10.9 - Flight Cab and Instruments __________________________________________ 70
Figure 10.10 - Graphics and Environment _________________________________________ 70
Figure 10.11 - Heads up Display ________________________________________________ 70
Figure 11.1 - Fuel Consumption Envelope at Average Climb Weight____________________ 72
Figure 11.2 - Drag on Aircraft in Loiter Conditions__________________________________ 73
Figure 11.3 - 1g Military Specific Excess Power Envelope at Maneuver Weight ___________ 75
Figure 11.4 - 1g Maximum Specific Excess Power Envelope at Maneuver Weight _________ 75
Figure 11.5 - 2g Maximum Specific Excess Power Envelope at Maneuver Weight _________ 76
Figure 11.6 - Maneuverability Diagram at 15,000 ft (4,572 m) and Maneuver Weight ______ 77
Figure 11.7 - Fuel Consumption over Mission ______________________________________ 78
Figure 11.8 - MPRL with 8 × 2,000 lb (907 kg) JDAM’s _____________________________ 82
Figure 12.1 - L to R configurations 1, 2, 3 _________________________________________ 84
Figure 12.2 - 180 inch MPRL___________________________________________________ 84
Figure 12.3 - Ballute and Sabot _________________________________________________ 84
Figure 12.4 - Bomb Bay Door Retraction Scheme___________________________________ 85
Figure 12.5 - 30in (76.2cm) Ejector rack __________________________________________ 85
Figure 12.6 - LAU-142A Ejection Sequence _______________________________________ 85
Figure 12.7 - (8) 2000lb JDAM + MPRL__________________________________________ 85
Figure 13.1 - Cockpit Width Trade Study _________________________________________ 86
Figure 13.2 - Forward fuselage Comparisons_______________________________________ 86
Figure 13.3 - Virtual Cockpit Model _____________________________________________ 87
v

Figure 13.4 – Rectilinear Vision Plot of Forward Cockpit Position______________________ 87
Figure 13.5 - Cockpit Display Arrangement________________________________________ 88
Figure 13.6 - Helmet Mounted HUD _____________________________________________ 88
Figure 13.7 - ACES II Ejection Seat______________________________________________ 89
Figure 13.8 - AFCPS__________________________________________________________ 89
Figure 14.1 - Sundstrand APS 3200 Location ______________________________________ 90
Figure 14.2 - Fuel Tank Locations in Vendetta _____________________________________ 92
Figure 14.3 - Fuel System Architecture ___________________________________________ 92
Figure 14.4 - Retractable in-flight refueling boom ports, F22, F-117, B-2 ________________ 92
Figure 15.1 - Routing Tunnel ___________________________________________________ 94
Figure 15.2 - Manufacturing Breaks______________________________________________ 94
Figure 15.3 - Assembly Line ___________________________________________________ 95
Figure 16.1 - Cost Analysis ____________________________________________________ 96
Figure 16.2 - Operating Cost ___________________________________________________ 97
Figure 16.3 - Lifecycle Cost ____________________________________________________ 97
vi

List of Tables
Table 1.I - Required Weapons Loadout ____________________________________________ 1
Table 1.II - Summary of Design Requirements ______________________________________ 2
Table 1.III - Comparison of the F-111, F-117, B-2, B-1B, and F-15E_____________________ 6
Table 2.I - Weight Fractions & Weights____________________________________________ 7
Table 2.II - Weight Fraction Assumptions __________________________________________ 8
Table 2.III - Constraint Assumptions ______________________________________________ 8
Table 4.I - Common Ground Radars______________________________________________ 18
Table 5.I - Wing Measurements _________________________________________________ 25
Table 6.I - Engine Specifications of RFP Supplied Engine ____________________________ 34
Table 6.II - RFP Dimensions Compared to the Snecma Olympus _______________________ 36
Table 6.III - IHPTET Goals ____________________________________________________ 36
Table 7.I - Materials Selection __________________________________________________ 51
Table 9.I - Initial Component Weight Buildup______________________________________ 55
Table 9.II - Final Component Weight Buildup______________________________________ 56
Table 9.III - Inertia Estimation __________________________________________________ 56
Table 10.I - Historical Aircraft Tail Volume Coefficients _____________________________ 60
Table 10.II - Pitching Moment Coupling with Rudder Deflection for Various Vertical Cant
Angles _________________________________________________________________ 64
Table 10.III - Rudder Control Power Results for OEI Condition________________________ 65
Table 10.IV - Longitudinal and Lateral Dynamic Mode Conformity with MIL-8785C ______ 69
Table 11.I - RFP Performance Requirements_______________________________________ 71
Table 11.II - RFP Design Mission _______________________________________________ 71
Table 11.III - Detailed Design Mission ___________________________________________ 73
Table 11.IV - Performance Measures of Merit______________________________________ 74
Table 11.V - Mission Results ___________________________________________________ 78
Table 11.VI - Fuel Consumption by Mission Segment _______________________________ 79
Table 11.VII - Takeoff Flight Profile _____________________________________________ 79
Table 11.VIII - Landing Flight Profile ____________________________________________ 80
Table 11.IX - Takeoff Results __________________________________________________ 80
Table 11.X - Landing Results ___________________________________________________ 81
Table 11.XI - Performance Summary_____________________________________________ 81
Table 11.XII - Alternate Mission Results__________________________________________ 83
Table 13.I - Military Vision Specifications ________________________________________ 87
Table 14.I - APU Selection Table________________________________________________ 90
Table 14.II - Fuel System Sizing Requirements _____________________________________ 93
vii

Nomenclature
AIAA American Institute of Aeronautics and Astronautics
A max Maximum Cross Sectional Area, ft 2
AR Aspect Ratio
C D Drag Coefficient
C D parasite Parasite Drag Coefficient
C D0 Zero Lift Drag Coefficient
C Di Induced Drag Coefficient
C Dwave Wave Drag Coefficient
Cf Skin Friction Coefficient
CG Center of Gravity, ft, in
C L Lift Coefficient
C l Section Lift Coefficient
C Lmax Maximum Wing Lift Coefficient
D Drag, lb
F F Form Factor
FS Fuselage Station
L Length of Fuselage, ft
L Lift, lb
L/D Lift to Drag Ratio
L HT Lever Arm of Horizontal Tail
L VT Lever Arm of Vertical Tail
M Mach Number
M Mach Number
MAC Mean Aerodynamic Chord, ft
M cd0 max Mach Number for Maximum Wave Drag
M cr Critical Mach Number
NPF Net Propulsive Force, lb
OEI One Engine Inoperable
P Pressure, lb/ft 2
P s Specific Excess Power, ft/s
Q Interference Factor
R Gas Constant, ft 2 /s 2 R
RCS Radar Cross-Section
Re L Reynolds Number Based on Length
Re Lcutoff Cutoff Reynolds Number
RFP Request for Proposal
S Sutherland’s Constant, R
S Wing Reference Area, ft 2
SFC Thrust Specific Fuel Consumption, lb/lb hr, lb/lb s
S HT Horizontal Tail Planform Area
SL, Sea Level
S ref Wing Reference Area, ft 2
viii

S VT Vertical Tail Planform Area
S w Wing Reference Area, ft 2
S wet Wetted Area, ft 2
T Thrust, lb
T Temperature, R
T/W Thrust to Weight Ratio
T available Available Thrust, lb
TOGW Takeoff Gross Weight, lb
T required Required Thrust, lb
TSFC Thrust Specific Fuel Consumption, lb/lb hr, lb/lb s
V Velocity, knots, ft/sec
V H Horizontal Tail Volume Coefficient
V stall Stall Velocity, knots, ft/s
V TD Touchdown Velocity, knots, ft/s
V TO Takeoff Velocity, knots, ft/s
V V Vertical Tail Volume Coefficient
W Weight, lb
W Fuel Flow Rate, slugs/s
F
W/S Wing Loading, lb/ft 2
a Speed of Sound, ft/s
a Temperature Lapse Rate in Troposphere, R/ft
c Chord
c Mean Aerodynamic Chord, ft
c Mean Aerodynamic Chord, ft
w
c root Root Chord, ft
c tip Tip Chord, ft
e Span Efficiency Factor
f Component Fineness Ratio
g Acceleration Due to Gravity, ft/s 2
h Altitude, ft
k Surface Roughness, ft
k 1 Induced Drag Coefficient
l Length of Drag Component, ft
n Load Factor
⎛ g ⎞
n Atmospheric Constant, ⎜n
= = 5.2561⎟
⎝ aR ⎠
q Dynamic Pressure, lb/ft 2
s Distance, n miles, ft
t Time, s
t Thickness of Wing to Chord Ratio
y mac y Location of Mean Aerodynamic Chord, ft
Λ LE Leading Edge Sweep, degrees, rad.
Λ t max Sweep of Position of Maximim Thickness, degrees, rad.
Trailing Edge Sweep, degrees, rad.
Λ TE
ix

α Angle of Attack, degrees, rad.
α eff Effective Angle of Attack, degrees, rad.
α i Angle of Attack Induced by Downwash, degrees, rad.
δ T Engine Correction Factor
γ Specific Heat Ratio
γ Climb Angle, degrees, rad.
λ Taper Ratio
µ Dynamic Viscosity, lb s/ft 2
µ brake Braking Coefficient of Friction
µ roll Rolling Coefficient of Friction
θ T Engine Correction Factor
ρ Density, slugs/ft 3
ψ Yaw Angle, degrees, rad.
Subscripts
0 Sea-Level Value
∗ Value at Tropopause
x

1 Introduction
Every year, the American Institute of Aeronautics and Astronautics (AIAA) sponsors collegiate
design competitions. The request for proposal (RFP) for the 2001-2002 team undergraduate
aircraft design competition outlined the requirement for a stealth supersonic interdictor. The
interdictor is introduced with a design mission as shown in Figure 1.1. The payload specified for
this design mission is shown in Table 1.I. Because multiple weapon load outs are specified, it is
clear that this, as with any modern aircraft, must be suited to more than one role.
The RFP also gives many requirements for the aircraft. These include operating constraints as
well as performance requirements. The design requirements are summarized in Table 1.II.
External tanks may be used but must be retained for the duration of the flight. Bomb pylons may
also be used suggesting the possibility of a non-stealth configuration. Another Important factor
is that the aircraft must cost less than 150 million dollars. This is a very small price tag for an
aircraft of this size and complexity.
Figure 1.1 - Design Mission Profile
Table 1.I - Required Weapons Loadout
Loading # (Quantity) Weapon
1 - Design (4) 2000 lb (907 kg) JDAM + (2) AIM-120
2 (4) Mk-84 LDGP + (2) AIM-120
3 (4) GBU-27 + (2) AIM-120
4 (4) AGM-154 JSOW + (2) AIM-120
5 (16) 250 lb (113 kg) Small Smart Bomb
1

Table 1.II - Summary of Design Requirements
Area Design Requirement Value (if applicable)
Misc.
Crew
500 lb (227 kg), 2 pilots, single pilot operation
Structure
Fuel
Stability
Observables
Operation
Positive g’s
7 (50% Internal Fuel)
Negative g’s
3 (50% Internal Fuel)
Dynamic Pressure
2,133 psf (102 kPa)
Factor Of Safety 1.5
JP-8
Self Sealing
Static Margin 10% to – 30%
Active Flight Controls for Unstable Aircraft
RCS (Front Aspect)
0.05 m 2 , frequency range 1 – 10 GHz
Balanced IR, Visual, Acoustical, RCS
Internal Stores
Runway Length 8,000 ft (2,438 m)
Operate from NATO Airports
All Weather Weapons Delivery
Cost
Max Cost
Minimize Life Cycle Costs
$150 Million, 2000 dollars
Performance
Supercruise Mission Radius
Specific Excess Power
1-g, Mach 1.6, 50,000 ft, Dry
1-g, Mach 1.6, 50,000 ft, Wet
2-g, Mach 1.6, 50,000 ft, Wet
Instantaneous Turn Rate, Mach 0.9, 15,000 ft
1,750 nm (3,240 km)
0 ft/sec (0 m/s)
200 ft/sec (61 m/s)
0 ft/sec (0 ft/s)
8 deg/sec
2

The RFP also lists several aircraft that collectively fulfill the mission of the proposed interdictor.
These are each outlined below.
F-111 - “Aardvark”
The F-111 (Figure 1.2, Table 1.III) is specifically mentioned as the predecessor to the aircraft
requested in the RFP. The F-111 officially entered service in 1967 and was retired in 1996; a
replacement is badly needed. It has been partially replaced by several aircraft, each outlined in
detail in the sections to follow. The F-111 is a very large aircraft capable of carrying a 31,000 lb
(14,061 kg) payload over 2,000 nm (3,704 km). Both the payload and combat radius are large
thus yielding a 91,000 lb (41,276 kg) aircraft. Though the F-111 is capable of Mach 2.2, but it
does not cruise supersonically. The F-111 was designed for a very different mission than the one
outlined in the RFP. The F-111 was actually designed to cruise subsonically to the target area,
dash in supersonically at low level, drop its payload, and fly out of the threat area quickly. After
retiring the aircraft, the air force decided a new aircraft was needed to drop precision weapons
from remote airfields with minimal support.
F-15E - “Strike Eagle”
Figure 1.2 - F-111 Aardvark
The F-15E Strike Eagle (Figure 1.3, Table 1.III) partially filled the role of the F-111 after it was
retired. The F-15E was designed to be capable of both air superiority and ground attack
missions. Superior maneuverability was achieved with the F-15E due to its high thrust-to-weight
ratio and low wing loading.
3

Figure 1.3 - F-15 Strike Eagle
F-117 – “Night Hawk”
The Night Hawk (Figure 1.4, Table 1.III) also aided in the replacement of the F-111. However,
it has a vastly reduced payload capacity and a limited range. The F-111 is also not capable of
supersonic speeds and is thus more vulnerable if it were detected. If a supersonic aircraft were
detected, the window of opportunity for an attack is relatively small. Thus, faster aircraft have a
tendency to be less venerable. Due to the small payload and high maintenance of the first
generation stealth technology, the F-117 is a poor replacement for the F-111.
Figure 1.4 - F-117 Night Hawk
4

B-1B – “Lancer”
The B-1 has never been a successful aircraft. Until the war against the Taliban, the B-1B was
never used in combat. Originally it was designed to have a mission much like the one specified
in the RFP. When the proposal for a high altitude supersonic bomber was withdrawn, politics
wouldn’t allow the B-1 to vanish. The aircraft was converted into the B-1B. The B model was
designated as a low level subsonic bomber much like the F-111 was. The B-1 is an expensive
aircraft to operate due to its vast size. A picture and table of data are also provided for this
aircraft (Figure 1.5, Table 1.III).
B-2 – “Spirit”
Figure 1.5 - B-1B Lancer
The B-2 (Figure 1.6, Table 1.III) is a large stealth subsonic bomber. It is designed to carry vast
amounts of payloads long distances undetected. The B-2 is a large aircraft that is very costly to
operate.
Figure 1.6 - B-2 Spirit
5

Table 1.III - Comparison of the F-111, F-117, B-2, B-1B, and F-15E
Manufacturer Lockheed
General
Dynamics
Boeing Northrop Rockwell
Type F-117 FB-111A F-15E B-2 B-1B
b - ft
(m)
43.6
(13.3)
32.0
(9.2)
42.8
(13.0)
172.0
(52.4)
78.2
(23.8)
AR – – 3 – –
L – ft
(m)
H – ft
(m)
Wing - ft 2
(m 2 )
W oe – lb
(kg)
W pl – lb
(kg)
W f – lb
(kg)
W to – lb
(kg)
66.6
(20.3)
12.5
(3.8)
913
(84.8)
29,500
(133,801)
5,000
(2,267)
73.5
(22.4)
17.1
(5.21)
–
46,171
(20,943)
31,500
(14,488)
– –
52,501
(23,814)
91,492
(41,500)
63.7
(19.4)
18.5
(5.6)
608
(56.5)
32,000
(14,515)
24,500
(11,113)
13,122
(5,952)
81,000
(36,740)
69.0
(21)
17.0
(5.2)
5274
(490.0)
153,700
(69,717)
40,001
(18,144)
200,003
(90,720)
375,998
(170,550)
147.0
(44.8)
34.0
(10.4)
1950
(181.1)
192,001
(87,090)
133,999
(60,780)
194,999
(88,450)
477,003
(216,365)
Max power loading – – 1.73 4.86 –
Max level (Mach) 0.9 2.2 2.5 0 1.25
Max combat radius
nm
Service Ceiling - ft
(m)
570 2,750 686 6,300 6,479
–
50,853
(15,500)
–
50,000
(15,240)
The solution to the RFP is not a trivial one. The aircraft will have to be well area ruled and have
a low frontal cross-section in order to minimize wave drag. The goal of this design is to meet or
exceed RFP requirements while minimizing manufacturing and operating costs.
–
6

2 Defining the Design Domain
The first estimation of aircraft weight used the iterative weight fraction method outlined in
Roskam. This method calculates the weight fraction for each mission segment. The initial
takeoff weight was guessed and the resulting empty weight was calculated. The resulting weight
generally yields an unrealistic total weight fraction. In order to check the realism of the aircrafts
weight fraction, a database of aircraft similar in mission was compiled. The aircraft gross
takeoff weight was iterated until the total weight fraction landed on the historical trend. This
trend is shown in Figure 2.1. The weights generated by this method include gross takeoff
weight, fuel weight, and landing weight. The results of the weight fraction method are shown in
Table 2.I. The assumptions made to generate the weight fractions are shown in Table 2.II.
Vendetta Estimated Weight Fractions
Cruise Back
16%
Reserve
6%
Misc.
5%
Warm -up &
Takeoff
6% Initial Climb
11%
Dash Back
14%
Dash Out
17%
Cruise Out
25%
Figure 2.1 - Historical Weight Fractions
Table 2.I - Weight Fractions & Weights
Mission Segment Weights
Weights
Start/Takeoff 6% Takeoff 108,400 lb (49,169 kg)
Climb To Cruise 11% Empty 51,600 lb (23,405 kg)
Cruise-Out 25% Fuel 47,600 lb (21,591 kg)
Dash-out 17% Payload 9,054 lb (4,107 kg)
Dash-Back 14% Fuel Weight Fraction 47.6%
Cruise-Back 16%
45 Minute Reserve 6%
Misc. 5%
Total 100%
7

Many of the equations buried in the weight fraction method depend on assumed parameters. In
many cases, values were assumed using figures and tables from Roskam, Nicolai, and Raymer.
The assumptions used are listed in Table 2.2. The weight fraction method is by no means an
accurate method for initial sizing. Inaccuracies of up to 10% are possible depending on the
quality of the initial assumptions, and 20% is not uncommon for unusual missions such as the
one outlined in the RFP. The weight fraction method may not be accurate for this type of aircraft
due to the lack of similar aircraft in the database. There are really only three supercruise aircraft,
the SR-71, YF-23, and the F-22. These aircraft all have a vastly different mission and may yield
invalid sizes for the interdictor. Though the method may be flawed, it was used anyway due to
the lack of a better method.
Weight fractions provide a starting point for the weight of a proposed aircraft; however, the
physical dimensions are not predicted. In order to determine the physical size, constraint plots
were created. A constraint plot examines the relationship between two variables based on given
requirements. Generally, the two variables used are wing loading and thrust to weight ratio.
The RFP gives many constraints as shown earlier in Table 1.II. The majority of constraints can
be written as functions of wing loading and thrust-to-weight ratio. This is the reason that it is a
popular type of constraint plot. The equations for range, takeoff distance, and many others were
found in Roskam, Nicolai, and Raymer. Many more assumptions were made to create the
constraint plot; these are shown in Table 2.III.
Table 2.II - Weight Fraction Assumptions
SFC _Cruise 1.11
SFC _Dash 1.11
SFC _Turn 1.11
SFC _Loiter 0.8
L/D Cruise 10
L/D Dash 10
L/D Turn 10
L/D Loiter 12
Table 2.III - Constraint Assumptions
C Lmax_TO 1.8
C Lmax_CR 1.2
C LCruise 0.2
AR 3
e 0.8
The constraint equations show how thrust to weight ratio and wing loading relate to a given
performance constraint. This allows engineers to determine which combinations of thrust to
weight ratio and wing loading are acceptable. The constraint plot for the RFP is shown in Figure
2.2. Note that any design point on the shaded side of a line would not meet the design
requirements. This again depends on the accuracy of the aforementioned assumptions. The
constraint plot clearly identifies a design domain, in which, any combination of thrust-to-weight
ratio and wing loading will satisfy the design requirements.
Combined with the weight fraction method, the constraint plot shows the physical size of the
airplane. Because a preliminary weight was determined from the weight fraction method, the
8

wing loading and thrust to weight ratio may be converted to wing area and thrust required. Once
this information is found the designer may start the configuration layout and choose a power
plant. In order to obtain the smallest aircraft, the design point should be at the highest wing
loading and lowest thrust to weight as possible on the constraint plot. As the design point moves
to the left on the constraint plot, the wing area increases. This yields a larger airframe. Large
airframes are expensive and cost more to maintain.
As the design point moves up on the constraint plot, the engines get bigger and heavier. This is
not desired as the aircraft will burn more fuel and have to fly at a higher lift coefficient for a
given Mach number. This would cause the fuel burn for a given mission and thus increase
operating costs. If the design point deviates from the lower right, an explanation is required.
The initial design point was chosen in the center of the design domain in order to allow for
aircraft growth. The design point was chosen away from constraint lines to make the aircraft
design performance less sensitive to changes in weight.
10
0.9
0.8
Current Design
Point
Takeoff
Landing
ψ
Range
Thrust to Weight Ratio
0.7
0.6
0.5
0.4
0.3
0.2
0.1
Initial Design
Point
Specific
Excess
Power
Specific
Excess
Power
40
60
80
100 120 140
Wing Loading, (psf)
160
180
200
Figure 2.2 - Constraint Plot
9

3 Configuration
The current configuration of the aircraft was developed through several iterations. The first
iterations were individual designs developed by each of the team members. Four different
designs were considered for the configuration the Vendetta would take and each was evaluated to
a similar level of detail.
The first of these designs was Nergal (Figure 3.1). This
was a concept that utilized thrust vectoring for stability
in the yaw axis. This configuration utilized a rotary
bomb bay configuration and a side-by-side cockpit
arrangement. This was the only design based on a
single PW-F119 as its power plant.
Figure 3.2 - Jackhammer
Figure 3.1 - Nergal
The Jackhammer shown in Figure 3.2 was another tailless
design. This aircraft differed from the Nergal only in its canard
arrangement and twin-engine approach. It included a rotary
bomb bay as well as F119 engines as its primary power plants.
The flight deck was a side-by-side layout that provided good
visibility for the pilots.
The Interdictor, shown in Figure 3.3, was based on the RFP engines
and sported good propulsive efficiency due to the lack of an S-style
duct. Like the previous aircraft, the Interdictor included a side-by-side
cockpit.
Figure 3.3 - Interdictor
The Big Paulie (Figure 3.4) had ample fuel volume and
utilized RFP engines. It was outfitted with a side by
side cockpit and ACES II ejection seats.
Figure 3.4 - Big Paulie
10

The downselect between these aircraft was not difficult. This early work clearly showed that the
engine provided by the RFP was far too large for the thrust it provided. The two aircraft
designed for the F119 were both smaller and more space efficient. This narrowed the downselect
to the Nergal and Jackhammer. Both of these aircraft were tailless and it was determined that
the weight and drag benefits associated with the lack of a vertical tail would be outweighed by
the costs associated with the thrust vectoring system. It was also determined that the aircraft
were too unstable laterally to be controlled by an inexpensive, low bandwidth, thrust vectoring
system. It was decided to begin with a new design incorporating the strong points of each
aircraft.
The first iteration of the aircraft is shown in Figure 3.5; it is a large aircraft that has many design
flaws. The first and most obvious is the above-chine mounted inlet. This is easily seen in the
aircraft’s front view. The chine causes a vortex roll-up that would be directly ingested by the
inlet at high angles of attack. This is not desired, as it would cause poor and unpredictable
performance at nearly any angle of attack. A low bypass ratio engine might tolerate these flow
disruptions without problems however; the design utilizes a new engine with a bypass ratio of
approximately 1.5. This type of engine will not tolerate the poor inlet location.
• Span = 50 ft
• m.a.c. = 23 ft
• S ref = 965 sq. ft
• TOGW = 121,600 lb
• Empty Weight = 62,000 lb
50’
19’
105’
23’
Figure 3.5 - Initial Configuration
Another downfall to the initial configuration was the weight distribution. The fuel center of
mass was not near the empty weight center of mass. This caused the aircraft to take off very
stable and land very unstable. This could not be remedied due to the small volume available for
fuel in the aft portion of the fuselage. The majority of the fuel volume in the aft portion of the
aircraft was located around the engines. This is undesirable due to the possibility of a
catastrophic failure of the engine fan disk or afterburner.
11

Another problem arises from the 20° facet on the bottom of the fuselage. This created a large
radar footprint underneath the aircraft, as shown in Figure 3.6. The vertical stabilizer also
created a very radar visible configuration. The final flaw that drives the aircraft to the new
configuration is the pitching moment characteristic of the fuselage. The side by side seating
arrangement of the first iteration caused the fuselage to be excessively large in the areas forward
of the aircraft’s neutral point. The pitch up tendencies of the aircraft grew very large with very
small angles of attack. The control power of the horizontal surfaces was deemed unacceptable to
combat this. The current configuration has a tandem cockpit arrangement, which will be detailed
later in the report, to help solve some of these issues.
Arc fuselage top creates
steps caused by facets
Less sensitive to
lower frequency radar
as wavelengths increase
and can’t distinguish facets
0
1 GHz
5 GHz
10 GHz
RCS in dBm 2
Flat bottom
Horizontal fuselage
intersection creates a
sharp cavity
20° bottom facet
Figure 3.6 - Radar Return of Initial Configuration
The second configuration, shown in Figure 3.7, is very different than the first. The
configuration features many changes that aide in solving the previously discussed problems. The
cockpit was changed to a tandem arrangement as well as the addition of two canted tails in the
place of the single vertical. The engines moved to the top of the fuselage to avoid detection from
infra red sensors. The take off gross weight was decreased to 114,000 lb due to an improved
engine deck and aerodynamics.
12

• Span = 53 ft
• m.a.c. = 32 ft
• S ref = 1500 sq. ft
• TOGW = 114,000 lb
• Empty Weight = 55,000 lb
35°
53’
19’
98’
18’
13°
Figure 3.7 - Second Configuration
This configuration was generated with a different mentality than the previous airframe. The
center of gravity was known before the first part was placed on the aircraft and every effort was
utilized to keep it in the appropriate place. The weight and balance issues, though still present,
were dramatically improved. The fuel load and payload compartment reside directly on the
desired center of gravity, however, the empty weight was too far aft. The low mounted wing
proved to be a structural challenge when incorporating a landing gear well. Another issue dealt
with the cruise angle of attack. It was shown that the aircraft would cruise at approximately 4
degrees. The forward chine on the fuselage would be shedding a vortex throughout the cruise
portion of the mission resulting in higher drag. The chine angle should meet the onset flow
angle. This prompted another revision to the aircraft.
The design was further refined and the current iteration was created. The current iteration has no
red flags and performs the mission well. The current iteration of the Vendetta is shown in Figure
3.8 and detailed in Foldout 1 of the Appendix. It has a forward chine of 4 degrees and sound
structural load paths, which will be discussed in detail later in this report. The Vendetta has
grown a small amount and currently weighs 124,000 lb (56,364 kg).
The aircraft has a tandem cockpit supported by a very long nose. The long nose offsets the mass
of the large engines and the massive structure required for the full flying horizontal stabilizers.
The APU is located in the engine compartment keeping the fuel and fire retardant systems as
redundant as possible. The inlets are under wing mounted to keep them in clean flow throughout
the flight envelope. The Vendetta has a 1500 ft 2 (139 m 2 ) wing area with a leading edge sweep
of 40 degrees. The design drivers will be discussed in detail throughout following sections. The
inboard layout can be seen in Figure 3.9 and 3.10.
13

• Span = 54.7 ft
• m.a.c. = 32 ft
• Length = 103 ft
• Root = 46.75 ft, 3% Thick
• Tip = 3%
• Λ c/4 = 27°
• Λ LE = 40°
• S ref = 1500 sq. ft
• TOGW = 129,000 lb
• Empty Weight = 60,000 lb
55’
103’
51°
20’
13°
Figure 3.8 - Current Configuration
APU
Weapons Bay
Retracted
Gear
Engines
Retracted
Gear
Figure 3.9 - Inboard Layout
14

Wing Tank X 2
6,366 lb each
70% Volume Usage
Forward Fuselage Tank
23,069 lb
80% Volume Usage
Aft Fuselage Tank
23,544 lb
80% Volume Usage
Tandem Cockpit
Figure 3.10 – Inboard Layout Continued
Full Flying
Horizontal
15

4 Stealth Considerations
As mentioned by the RFP, the aircraft is required to meet a stealth requirement. This is a very
important driver for the aircraft. The entire design is influenced by this consideration equally as
much as aerodynamics. There are many low observable considerations to be taken into account.
The first and most obvious is the radar cross section (RCS).
From the aspect of RCS, there are many drivers for an aircraft. The majority of the radar return
comes from the shaping of the aircraft. The fuselage is constructed from flat sides and constant
radius curves. The sides are kept at a 60° angle from the horizontal and the bottom is kept flat
(Figure 4.1). This is desired because, as later shown, the footprint of the aircraft remains small.
Another feature is the canted tails. This keeps the surfaces in the empennage section from
creating 90 degree angles. This is important because the 90 degree angle would radiate RF
energy directly back in the direction of the source. The leading edge sweep is 40°. This creates
spikes well off of the frontal aspect of the aircraft. All other leading edges are kept swept at this
same angle in order to minimize the magnitude of the frontal spoke. The 15° look up angle was
considered the most important aspect of the RCS. The majority of the threat encountered will be
below the Vendetta. This implies that they will be looking up at the aircraft, not from the front.
This is where the majority of the stealth considerations were taken into account. (See Figure 4.1)
The Low observability requirements are not only for RCS. In fact, the RFP specifically specifies
“Balanced Observables”. Aside form RCS, IR accounts for the next highest threat. Emissivity
matching can reduce the IR signature of the aircraft. The Vendetta will be coated with a material
with similar emissivity as the surroundings, aiding in the disappearance of the aircraft to an IR
sensor. The actual odds of becoming invisible to the IR sensor are fairly unrealistic due to the
cold surroundings. The aircraft is shadowed by what is essentially space at 50,000 feet. It is
hard to hide a warm object when backlit by a cold space. The other stealth consideration in this
area is the nozzles. These are axisymmetric nozzles that are proven to have lower IR signatures
than the axis symmetric option. This, combined with the frontal look-up threat direction
minimizes the impact on IR stealth.
40° LE Sweep
All other Surfaces
Matched
Hidden Canted
Verticals
60° Facet
Figure 4.1 - Stealth Considerations
16

Another area of concern is that of visual observability. This is not a very big problem as the
aircraft cruises so high that visually detecting the Vendetta would be near impossible. The only
threat here is the contrails left by the engines. The contrails can be minimized by contrail
avoidance techniques and don’t pose a large problem.
To quantitatively analyze the radar cross section of the Vendetta, Radbase2 software by Surface
Optics was utilized. First, a faceted model was generated from the 3D model. Faceting was
limited to only those necessary because of the demanding processing requirements. Facets were
limited to 10 degree tolerances at roughly 0.017 feet minimums. The facetted model is presented
as Figure 4.2.
Figure 4.2 - RCS Model Faceting
It can be seen that heavy facet optimization was needed to make sure that all facets followed
tangency requirements to leave smoothly curved and splined surfaces. The spline arc on the top
of the fuselage is modeled with facets every 10°. For the flat surfaces like the wings and
empennage 10° is more than adequate.
The Radbase2 RCS code calculates the radar returns based on Physical Optics and Chu-Stratton
integral methods. These are highly computationally intensive. Because of this, bounces off of
surfaces were limited to 2 after the initial bounce off the surface. This was deemed adequate for
this level of analysis. The vertical-vertical polarization of the return and transmission was
analyzed as it is the most relevant to how radar stations operate. Monostatic radars which both
broadcast and receive were used in the analysis as there would be too many possibilities to
calculate for bistatic radars.
The code was allowed to iterate on the model with 1° azimuth increments and for 0° and 15°
lookup angles. It was also run for 1, 5, 10, and 12 GHz radar frequencies. Most fast track and
search radar runs at the higher frequencies while long range threat radars utilize the lower
frequencies. The 1 to 10 GHz range covers most of the radars that are expected for the role of
this aircraft. A table of common ground and surface radars with their respective frequencies is
presented as Table 4.I.
17

Table 4.I - Common Ground Radars
Radar Manufacturer Frequency Image
AN/TPS-43E Mobile
Radar
Westinghouse
2.9 to 3.1 GHz
AN/TPS-70
Fixed Ground Radar
Northrop
Grumman
2.9 to 3.1 GHz
AN/SPS-49
Typical Long Range
Naval Radar
Navy Research
Labs
850 to 942
MHz
AN/SPS-55
Long Range Surface
Search Radar
ISC Cardion
9.05 to 10.0
GHz
Data is not readily available for radars made by foreign manufacturers. However, these radars
should be adequate for this level analysis because the properties for radar waves traveling
through the air over long distances are similar.
The 1 to 12 GHz range covers FM and XM radar bands which are most common threats. The
RFP specifically requires that the Vendetta has a frontal RCS of 0.05 m 2 . As the threat chart
shown in Appendix A shows, most threats will be from below and at shallow angles of about 15°
while at 50,000 during ingress. Because of this, the 0° and 15° lookup angles were analyzed. The
results of the Radbase2 software are illustrated first in Figure 4.3, which depicts the radar cross
section of the aircraft from a frontal, or 0° lookup angle.
18

30
20
10
0
-10
-20
-30
-40
-50
1 GHz
5 GHz
10 GHz
12 GHz
RFP Requirement
Figure 4.3 - Radar Cross Section at 0º Lookup Angle
Figure 18 shows that the vehicle does clearly meet the frontal RCS requirement of 0.05 m 2 (-12
dB) set forth in the Request for Proposal. It also shows that the measures taken at shaping the
aircraft are working. The leading edge and trailing edge of the wing come together closely. There
is a large return directly from the side of the aircraft due to the wing tip and fuselage side. It can
also be seen that although there are slight variations in the returns due to the different
frequencies, they do not vary that much. This is due to the fact that the Vendetta is a rather large
vehicle. None of the surfaces are small enough to interfere with the wavelengths of the radar.
The weakest azimuth angle for the Vendetta is the 40° angle where the leading edge sends a large
spike forward.
Looking at the equally crucial 15° lookup angle cross section in Figure 4.4 reveals a slightly
different picture.
19

30
20
10
0
-10
-20
-30
-40
-50
1 GHz
5 GHz
10 GHz
12 GHz
RFP Requirement
Figure 4.4 - Radar Cross Sections at 15º Lookup Angle
Figure 4.3 shows that the Vendetta meets and exceeds the 0° lookup angle returns. This is seen as
highly advantageous. The shape of the bottom of the aircraft is effective in keeping spikes at a
minimum. As mentioned earlier, this is a crucial area for the Vendetta. As most of its threats are
from the ground, it is important that the aircraft has a limited return in this orientation.
20

The software was also utilized to generate an RCS butterfly plot in a sweep around the vehicle to
determine the footprint that it will leave as it flies above its threats. Figure 4.5 shows this sweep.
60
50
40
30
20
10
0
-10
1 GHz
5 GHz
10 GHz
12 GHz
Figure 4.5 - Radar Cross Sections for a Radial Sweep
It can be seen that the 30° facets on the bottom of the fuselage are deflecting radar away from the
vulnerable lookup orientation. The aircraft is still producing a rather large return of almost 40 dB
in this position, however. Once again there is little variation in the returns for various
frequencies.
It is important to note that the addition of radar absorbing material (RAM) would further reduce
some of the returns on the aircraft. Also, currently the RCS software is treating the entire aircraft
and all of its parts as purely reflective metal surfaces. This is a conservative approach. RAM
could be applied in actuality to reduce some of the returns on the bottom and front of the aircraft.
21

5 Aerodynamics
The major aerodynamic aspects of the Vendetta are cruise lift-to-drag ratio and maximum
subsonic lift coefficients for landing performance. Wing design is critical to both of these
aerodynamic aspects as well as radar cross section. Area ruling was used to minimize the
supersonic wave drag on the aircraft and minimize the aircraft’s fuel consumption over the
mission.
5.1 Wing Sizing
The first aerodynamic aspects of the aircraft that were considered were the wing planform area
and aspect ratio. To select the optimum wing planform area and aspect ratio, the effect of these
two parameters on the specific excess power and fuel consumption over the design mission was
studied. The 1g military specific excess power at an altitude of 50,000 ft (15,240 m) and Mach
number of 1.6 was estimated using engine data and drag estimation based on component skin
friction drag and area ruling. The fuel consumption of the aircraft was estimated by numerically
integrating the engine fuel flow from over the design mission. The additional weight and
maximum cross sectional area of larger wing areas was considered in calculations, however the
mission profile was kept constant and the fuel weight at takeoff was kept constant at 59,250 lb
(26,875 kg). The results shown in Figure 5.1 indicate that a wing planform area of
approximately 1,500 ft 2 (139 m 2 ) and aspect ratio of 2 would maximize specific excess power
and minimize fuel consumption.
97
96
1,600 ft 2
Design Point
1,500 ft 2
Fuel Onboard
Specific Excess Power (ft/s)
95
94
93
92
1,700 ft 2
1,400 ft 2
1,300 ft 2
1,900 ft 2 1,200 ft 2
1,100 ft 2
Wing Area
91
1,800 ft 2 2,000 ft 2 Aspect Ratio 1.6
2.2
2.4
2
1.8
1,000 ft 2
90
57,000 58,000 59,000 60,000 61,000 62,000 63,000 64,000 65,000 66,000
Fuel Consumption over Mission (lb)
Figure 5.1 - Optimization of Wing Area and Aspect Ratio
22

5.2 Wing Planform
The next aspect of the wing that was considered was the leading and trailing edge sweep angles.
Because any edges on an aircraft reflect radar energy, the sweep angles of the edges of the wing
were chosen to minimize radar energy reflected back to the source, especially in the frontal
aspect of the aircraft where a specific RCS requirement is given by the RFP. To avoid reflecting
radar toward the front of the aircraft, the leading and trailing edges of the wing had to be highly
swept. In addition, the sweep angles could not be approximately 45º because a corner reflector
would be created. These requirements led to a diamond shaped wing planform with leading and
trailing edge wing sweeps of approximately 40º. Two initial designs were considered one having
a 40º swept leading edge and a 30º forward swept trailing edge and the other having matched
35.3º leading edge and trailing edge sweeps. A trade study was performed to select between
these two wing configurations by studying the effect of the two configurations on RCS and
aerodynamics. Figure 5.2 shows a comparison of radial sweeps of both configurations using
RadBase2. The return from the 40º and 35.3º leading edge sweeps can be clearly seen in the
plot. The leading edge spike on the matched leading and trailing edge configuration is
approximately 15 dB lower than the other configuration; however it is 5º closer to the frontal
aspect of the aircraft. The aerodynamic study of the two wing configurations indicated that
approximately 1,000 lb (454 kg) of additional fuel would be required due to the additional wave
drag of the lower leading edge sweep angle. Because of the aerodynamic benefits and because
the RFP only gives frontal aspect RCS requirements, the 40º leading edge and 30º trailing edge
configuration was chosen.
23

1 GHz. 40º LE Sweep
10 GHz. 40º LE Sweep
1 GHz. 35.3º LE Sweep
10 GHz. 35.3º LE Sweep
RFP Requirement (-12 dB)
50 dB
40 dB
30 dB
20 dB
10 dB
35.3º LE Sweep
40º LE Sweep
0 dB
-10 dB
-20 dB
-30 dB
-40 dB
-50 dB
Figure 5.2 - Effect of Wing Leading and Trailing Edge Sweep on Aircraft RCS
Once the wing area, aspect ratio, and sweep angles were chosen, the tip chord was kept at 8 ft
(2.4 m) to avoid an overly small tip chord that could interact with radar wavelengths
unpredictably. This resulted in the wing planform shown in Figure 5.3, with the measurements
given in Table 5.I. Leading and trailing edge flaps, and ailerons were added to the wing. The
chord high lift devices and control surfaces were kept at a constant percentage of the mean
aerodynamic chord so that the hinge lines would parallel to the wing edges. The trailing edge
flap chord is 20% of the mean aerodynamic chord and the leading edge flap and aileron are each
10% of the mean aerodynamic chord. The trailing edge flap extends from the fuselage to 65% of
the semi-span, the leading edge flap extends from the fuselage to 90% of the semi-span, and the
aileron extends from the edge of the flap to 90% of the semi-span. No moveable surfaces were
added to the last 10% of the semi-span so that radar obsorbing materials (RAM) could be added
in the wing tip to minimize any returns from that edge.
24

Figure 5.3 - Wing Planform
Table 5.I - Wing Measurements
Planform Area 1,500 ft 2 (139 m 2 )
Span 54.8 ft (16.7m)
Root Chord 46.8 ft (14.3 m)
Tip Chord 8.0 ft (2.4 m)
MAC 32.0 ft (9.8 m)
y Location of MAC 10.5 ft (3.2 m)
Aspect Ratio 2.0
Leading Edge Sweep 40.0º
Sweep at Quarter Chord 20.5º
Sweep at Half Chord 4.7º
Trailing Edge Sweep -30.0º
Taper Ratio 0.17
Leading Edge Flap Area 137 ft 2 (12.7 m 2 )
Trailing Edge Flaperon Area 238 ft 2 (22.1 m 2 )
Flapped Wing Area 935 ft 2 (86.9 m 2 )
25

5.3 Wing Thickness
The effect of wing thickness on the performance of the aircraft was studied so that the optimum
thickness could be chosen. Initially a wing thickness of 3% of the chord was chosen based on
existing supercruising aircraft. Increasing the root thickness of the wing was considered to
reduce the weight of the wing. The effects of wing root thickness on wing weight, cross sectional
area, and fuel consumption were studied. The weight of the wing was estimated using the
method presented in Raymer, and the additional cross sectional area was calculated numerically.
The resulting wing weights and cross sectional areas for wing root thicknesses from 3% to 6%
are shown in Figure 5.4. The effect of the resulting weights and cross sectional areas on the fuel
consumption during the mission were estimated using the same method used for the wing sizing.
The results in Figure 5.5 show that the larger cross section of a thicker wing root adds more
wave drag than the induced drag savings from the reduced wing weight. Based on this result, a
constant wing thickness of 3% was chosen.
8,500
8,000
t root = 3%
Weight of Wing (lb)
7,500
7,000
t root = 4%
t root = 5%
6,500
t root = 6%
6,000
14 15 16 17 18 19 20 21 22
Maximum Frontal Cross Sectional Area of Wing (ft 2 )
Figure 5.4 - Effect of Root Chord Thickness on Wing Weight and Cross Sectional Area
26

62,500
Fuel Consumption over Mission (lb)
62,000
61,500
61,000
60,500
60,000
59,500
59,000
58,500
58,000
Fuel Onboard
57,500
3.0% 3.5% 4.0% 4.5% 5.0% 5.5% 6.0%
Wing Root Thickness
Figure 5.5 - Effect of Root Chord Thickness on Fuel Burn
27

5.4 Airfoil
The NACA 65A-003 airfoil section was chosen for the aircraft, because a symmetrical airfoil
with the maximum thickness at 50% of the chord is optimum for supersonic flight. The airfoil
ordinates given in Theory of Wing Sections for an NACA 65A-006 were scaled and interpolated
using Lagrangian polynomials to define the geometry of the wing. The leading edge radius of
the airfoil is 0.1% of the chord, which is approximately 3/8 inch at the mean aerodynamic chord
and 1/10 inch at the tip. The airfoil sections at the mean aerodynamic chord and tip of the trailing
edge flap are shown in Figure 5.6 and Figure 5.7 respectively. Because the chords of the flaps
remain constant as the wing chord changes, each airfoil section has a different relative flap size.
0.1
0.05
0
-0.05
-0.1
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
Figure 5.6 - Airfoil Section at MAC
0.15
0.1
0.05
0
-0.05
-0.1
-0.15
-0.2
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
Figure 5.7 - Airfoil Section at Tip of Trailing Edge Flap
5.5 Lift Curve
To estimate the lift curve of the wing, first, the lift curve slope of the wing was estimated using
standard subsonic theory, compressibility corrections, and linear supersonic theory. The
resulting lift curve slopes are shown as a function of Mach number in Figure 5.8.
28

5
4.5
4
Lift Curve Slope (1/rad)
3.5
3
2.5
2
1.5
1
0.5
0
0 0.5 1 1.5 2 2.5 3
Mach
Figure 5.8 – Variation in Lift Curve Slope with Mach Number
Next, the stall angle of the wing was estimated under subsonic conditions by calculating the lift
distribution of the wing using LinAir. The section lift coefficient was calculated as a function of
the span-wise location of the section for different wing angles of attack. The wing was assumed
to stall when one of the section lift coefficients exceeded the maximum lift coefficient given in
Theory of Wing Sections. The stall angle of attack of the wing was determined to approximately
14º. Because the wing tip was shown to stall at a much lower angle of attack than the rest of the
wing, adding a –3º angle of incidence to the wing tip was considered. The resulting twist
extends the stall angle of attack to approximately 16º; however the twist decreased the lift
coefficient at a given angle of attack and could impact RCS and supersonic aerodynamics.
Ultimately, the non-twisted wing was chosen. The lift distributions of the wing with and without
twist are shown in Figure 5.9.
The effects of the trailing edge flap were estimated using the stall angle of attack and lift
coefficient increments given in Nicolai. The effect of the leading edge flap was estimated by
assuming that a 10º leading edge flap deflection would increase the stall angle of attack by
approximately 10º, and the decrease in lift coefficient was estimated based on the change in
effective angle of attack. The resulting subsonic lift curve at Mach 0.2 is shown in Figure 5.10.
29

1
Section Lift Coefficient
0.9
0.8
0.7
0.6
0.5
0.4
0.3
Max. Section Lift Coefficient
0º Tip Incidence
16º
15º 16º
15º
14º
13º
14º
12º
13º
12º
- 3º Tip Incidence
Calculated Using LinAir
0.2
0.1
0
0% 10% 20% 30% 40% 50% 60% 70% 80% 90% 100%
Spanwise Distance (percent semi-span)
Figure 5.9 - Lift Distribution of Wing with and without Twist
2
Lift Coefficient
1.5
1
0.5
0
C L = 1.20
C L = 1.15
C L = 0.56
C L α = 2.33 1/rad
C L = 1.51
30º TE Flap
Deflection
Clean
LinAir
10º LE Flap
Deflection
-0.5
Tail Strike Angle (14º)
-1
-20 -15 -10 -5 0 5 10 15 20 25 30
Angle-of-Attack (degrees)
Figure 5.10 - Subsonic Wing Lift Curve
30

5.6 Drag
The drag of the aircraft was divided into four parts: parasite drag, wave drag, induced drag, and
trim drag. The parasite drag was estimated using a component build method with form and
interference factors. The wave drag was calculated using the formula presented in Brandt &
Stiles. The wave drag efficiency factor was calculated from cross sectional area distributions
using the de Kármán integral and the theoretical wave drag of a perfect Sears-Haack body. The
cross sectional area distributions were measured at transonic and supersonic (Mach 1.6)
conditions. The transonic case was measured by passing vertical planes through a solid model of
the aircraft and measuring the intersecting area. The supersonic case was measured by passing
Mach cones through the model, measuring the intersecting area, and projecting that area onto the
vertical plane. For both cases, the engine capture area was subtracted from sections containing
the inlet, engine, and nozzle. The resulting area distributions shown in Figure 5.11 and Figure
5.12 match reasonably well with that of a perfect Sears-Haack body. Both distributions yield a
wave drag efficiency factor of approximately 2.14 (based on 80 ft 2 (7.4 m 2 ) max. area and 100 ft
(30.5 m) length).
90
80
70
Sears-Haack
Wing
Cross Sectional Area (ft 2 )
60
50
40
30
Fuselage
Vertical Tail
Horizontal Tail
20
10
0
0 200 400 600 800 1,000 1,200
Fuselage Station (inches aft datum)
Figure 5.11 - Transonic Area Distribution
31

90
80
Wing
Cross Sectional Area (ft 2 )
70
60
50
40
30
Sears-Haack
Fuselage
Vertical Tail
Horizontal Tail
20
10
0
0 200 400 600 800 1,000
Fuselage Station (inches aft datum)
Figure 5.12 - Supersonic Area Distribution (Mach 1.6)
Induced drag was estimated using standard subsonic theory and the supersonic equation
presented in Brandt & Stiles to calculate the induced drag term (k 1 ). Trim drag was calculated as
induced drag generated by the horizontal tail at the lift coefficient required to trim the aircraft
with a given static margin and moment coefficient. The resulting drag build-up for the aircraft at
an altitude of 50,000 ft (15,240 m), Mach number of 1.6, maneuver weight of 94,735 lb (42,971
kg), and 5% static margin is shown in Figure 5.13.
32

0.06
0.05
Drag Coefficient
0.04
0.03
0.02
Induced Drag
50,000 ft
Maneuver Weight
Trim Drag
0.01
Wave Drag
Parasite Drag
0
0 0.5 1 1.5 2 2.5 3
Mach
Figure 5.13 - Drag Build-Up at 50,000 ft, Mach 1.6, Maneuver Weight, and 5% Static Margin
33

6 Propulsion
In developing the propulsion system for the Vendetta, the RFP specifications of supersonic cruise
and stealth are the driving factors for the propulsions system. Due to the stealth criteria the fan
blades of the engine must remain hidden which drives the engine placement inside the airplane.
That combined with the supercruise criteria leads to four major aspects of the propulsion system:
the Engine, Inlets, S-ducts and Nozzles.
6.1 Engine Selection
The RFP specifies that the airplane should perform the mission with adequate installed thrust. A
Low-Bypass-Ratio Turbofan (LBR-TF) or a Turbojet (TJ) engine may be used to perform the
mission. Equations are provided for creating an engine deck, they are as follows:
LBR-TF:
TJ:
ρ
Tdry
= 0.9 TSL −dry
(0.88 + 0.24 ABS( M − 0.6) )( )
ρSL
2 ρ 0.8
Twet
= TSL −wet
(0.94 + 0.38 ABS( M − 0.4) )( )
ρ
1.4 0.8
ρ
Tdry
= 0.9 TSL −dry
(0.907 + 0.262 ABS( M − 0.5) )( )
ρSL
2 ρ 0.8
Twet
= TSL −wet
(0.954 + 0.38 ABS( M − 0.4) )( )
ρ
SL
1.5 0.8
However an installed engine deck for a LBR-TF with an axisymmetric center body inlet and a
mixed flow ejector nozzle was supplied with the RFP. Since it included physical dimensions and
fuel flow values the RFP engine deck was used instead of the above equations. The RFP engine
specifications are shown in Table 6.I.
Table 6.I - Engine Specifications of RFP Supplied Engine
Engine and Nozzle Length
Propulsion System Length
Fan Face Diameter
Maximum Diameter
Weight with Nozzle
SL
310 inches (787 cm)
425 inches (1080 cm)
50 inches (127 cm)
65 inches (165 cm)
7200 pounds (3266 kg)
The engine supplied by the RFP includes fuel flow and thrust data for part power, idle power,
and military power. All engine data supplied by the RFP is corrected to sea level and a Mach
number of zero. Therefore, every value for thrust and fuel flow at each altitude and Mach
34

number is given in corrected net propulsive force (NPF c ) and corrected fuel flow (WF c ). To find
the actual thrust (NPF) and fuel flow (WF) the following equations were used:
NPF = NPF ⋅ d
c T
0.6
F
=
F
⋅
c T
⋅
T
W W Q d
2 3.5 P
dT
= (1 + .2 M ) ( )
PSL
2 T
QT
= 1+
0.2 M ( ) T
Once the data was uncorrected the military thrust was found. The RFP supplied equations that
could be used to scale the engine based on a desired thrust. The scaling equations are as follows:
NPF
NewMeasurement = OldMeasurement( ) NPF
base
Axial length scaling exponent = 0.4
Diameter scaling exponent = 0.5
Weight scaling exponent = 1.0
SL
exponent
The RFP engine produced a military thrust of 26,350 pounds (117,210 N) and had a cruise thrust
specific fuel consumption (TSFC) of 1.19 1/hr (0.121 kg/h/N) for Mach 1.6 flow at 50,000 ft
(15,240 m). TSFC is calculated using the following equation:
TSFC =
W F
NPF
The Vendetta would require two engines to perform the desired mission. The size, weight, and
location of the engines have great effect on the size of the airplane. The larger the engines the
wider the aft portion of the fuselage and the longer the airplane. For the size and weight of the
RFP engine it produced merely too little thrust and burned too much fuel.
Other engines were sought out and analyzed in an attempt to find a better performing engine that
was smaller and lighter than that supplied. Through this research the Concorde’s Rolls-Royce
Snecma Olympus engine was found to be comparable to the RFP engine. However, the engine
was first manufactured and flown in the Concorde in the mid 60’s through mid 70’s. Table 6.II
compares the RFP engine to that of the Snecma Olympus. As can be seen the Snecma Olympus
is very close in size and weight to that of the RFP, however it produces even more thrust than
that of the RFP. Also the weight of the Snecma Olympus includes that of an afterburner whereas
the RFP engine is without an afterburner.
35

Table 6.II - RFP Dimensions Compared to the Snecma Olympus
RFP
Snecma Olympus
Fan Face Diameter 50 in (127 cm) 47.5 in (120.65 cm)
Length 310 in (787 cm) 280 in (711 cm)
Weight 7200 lbs (3266 kg) 7000 lbs (3175 kg)
Max Dry Thrust 26,356 lbs (117,237 N) 31,350 lbs (139,452 N)
Based on this data the RFP engine was assumed to be an older engine; a more efficient and
modern engine would be needed for the design of the Vendetta. The RFP engine deck was used
as a baseline for designing a newer better engine, as it was the only full engine deck available. It
was determined that an F119 engine would be the initial design engine for the airplane. This
engine is currently used in the F-22 and a derivation of the engine (the F135) is to be used in the
F-35.
All information regarding the F119 is classified except that it is in a 35,000 lbs (155,700 N)
weight class. Several methods were utilized to narrow in on the thrust produced by the F119.
Through the use of The Integrated High Performance Turbine Engine Technology (IHPTET)
program F119 characteristics were determined. IHPTET which began in 1988 and will culminate
in 2005 consists of a 3 phase plan, utilizing the most current advancements in industry. “IHPTET
is producing revolutionary advancements in turbine engine technologies due to the synergistic
effect of combining advanced material developments, innovative structural designs and
improved aerothermodynamics ” The 3 phases of the program are shown in Table 6.III.
Phase III (2005)
Phase II (1997)
Phase I (Completed)
Table 6.III - IHPTET Goals
+100% Thrust/Weight
-40% Fuel Burn
+60% Thrust/Weight
-30% Fuel Burn
+30% Thrust/Weight
-20% Fuel Burn
The Air Force Research Laboratory states that Phase I of the program has been completed and
that the technology has been applied to existing engines including the F100, F110, F414, and the
F119. Based on this data Phase I was applied to the RFP engine deck to yield an F119 engine.
This was done because both the RFP engine and the F119 are low bypass turbofan engines. The
20% decrease in fuel burn was applied and then the weight was decreased by 21% and the thrust
increased by 2% to account for the 30% change in thrust to weight. The resulting thrust produced
by the F119 is 27,000 pounds (120,101 N), has a cruise TSFC of 0.94 1/hr (0.096 kg/hr/N) and a
weight of 6,575 lbs (2,575 kg).
Vendetta’s design is to be frozen in 2010 and future advancements and technologies are to be
taken into account for its development. Phase II of IHPTET was to be completed by 1997 yet has
not been achieved. However, Pratt and Whitney has proven a 40% increase in thrust to weight.
36

The weight of the F119 was decreased down to 5,270 lbs (2,384 kg) to account for the additional
10% increase in thrust to weight. However there is no evidence of any other advancements with
the IHPTET program, and advancements in turbofan engines were sought out. The Versatile
Affordable Advanced Turbofan Engine (VAATE) is an industry projection to 2020. Even though
it builds upon IHPTET it uses the F119 as a base engine for its future goals. Figure 6.1 illustrates
the goals for turbofan engines thru 2020 and Phase I goals of a 25% decrease in TSFC and a 45%
decrease in cost by 2010. It is likely that this program will face similar problems in achieving its
goal by 2010, in which case a decrease of 13% in TSFC was taken and an estimated 25%
decrease in cost over the F119. The 13% change in TSFC was achieved by increasing the thrust
by 11% and decrease the fuel flow 2%. The new VAATE technology engine has a sea level
thrust of 30,000 lbs (133,500 N) and a cruise TSFC of 0.82 1/hr (0.084 kg/hr/N), however once
inlet and ducting losses are accounted for the cruise TSFC is 0.92. 1/hr (0.094 kg/hr/N).
37

38
Figure 6.1 - VAATE Goals

The data for the idle power of the engine was based off of the original RFP. The idle data has not
been changed since there is no distance credit or fuel credit for the descent portions of the
mission and that is the only time idle settings would be used for this mission. The military thrust
TSFC of the engine at various altitudes can be seen in Figure 6.2 and Figure 6.3 respectively.
The afterburner model was created by multiplying the military thrust by 1.65 and the military
fuel flow by 2. The equations provided by the RFP for calculating thrust from afterburner were
not utilized as the maximum thrust produced by those equations yielded the same amount of
thrust as the F119 at military power and altitude.
Thrust (lbs/eng)
35,000
30,000
25,000
20,000
15,000
10,000
Sea Level
Alt 5,000 ft
Alt 10,000 ft
Alt 20,000 ft
Alt 25,000 ft
Alt 30,000 ft
Alt 36,089 ft
Alt 43,000 ft
Alt 50,000 ft
Alt 55,000 ft
Alt 60,000 ft
Alt 70,000 ft
5,000
0
0 0.5 1 1.5 2 2.5
Mach Number
Figure 6.2 - Thrust Curves for Altitudes from Sea Level to 70,000 ft (21,300 m)
39

TSFC
3.0
2.8
2.6
2.4
2.2
2.0
1.8
1.6
1.4
1.2
1.0
0 0.5 1 1.5 2 2.5 3
Mach Number
Sea Level
Alt 5,000 ft
Alt 10,000 ft
Alt 20,000 ft
Alt 25,000 ft
Alt 30,000 ft
Alt 36,089 ft
Alt 43,000 ft
Alt 50,000 ft
Alt 55,000 ft
Alt 60,000 ft
Alt 70,000 ft
Figure 6.3 - Military TSFC Curves for Altitudes from Sea Level to 70,000 ft (21,300 m)
Once the engine deck had been established the engine dimensions were once again considered.
The fan face diameter of the new engine was assumed to be that of the RFP. Low bypass
turbofans typically have smaller fan face diameters however as the bypass ratio increases the fan
face diameter would increase as well. Since future technology is being taken into account it is
likely that the engine that would produce this thrust would have a larger bypass ratio but smaller
core keeping the fan face diameter comparable to that of the RFP. The length of the engine was
estimated based off of previous trends in engines. Engines used to plot this data include the
F100, F101, F110, F404 JT3D, JT8D and TF30. This data can be seen plotted in Figure 6.4
below. Logarithmic trendlines were fitted to the data and then extrapolated out past 30,000 lbs
(133,500 N). For a thrust of 30,000 lbs (133,500 N) the engine would have a maximum diameter
of 60 inches (152 cm) and a maximum length of 220 inches (559 cm).
40

Diameter (in)
70
60
50
40
30
20
10
300
250
200
150
100
50
Length (in)
0
0
12,000 17,000 22,000 27,000 32,000
Thrust (lbF)
Figure 6.4 - Engine Sizing Plot
6.2 Inlets
Sizing the inlet for supercruise flight at 1.6 Mach posed an interesting problem. A pitot inlet is
good up until about 1.6 Mach and it is by far the cheapest inlet possible. However the
performance of the inlet above Mach 1.6 is very poor. The pressure recovery of a two shock inlet
(one oblique and one normal shock) and a three shock inlet were analyzed. The optimum
deflection angle for Mach 1.6 flow was found for a two shock inlet by finding the stagnation
pressure loss across the oblique and normal shock for different deflection angles. The results
were graphed in Figure 6.5 and the resulting deflection angle for the greatest pressure recovery
was found to be 10.75 degrees yielding a pressure recovery of 97.65%. Finding the optimum
deflection angle for a three shock inlet is more involved therefore a rough estimate of a six
degree deflection angle followed by another 6 degree deflection angle was used to compare
against the two shock inlet. The difference in on design pressure recovery is about 1% however
the larger the deflection angles become the better the pressure recovery will become. The
pressure recovery comparison can be seen in Figure 6.6. The military specification for inlets is
given below and is represented in the graph.
Mil Spec MIL-E-5008B
η
rSpec
⎧ 1 M ≤ 1
⎩1
0.075( 1) 1 5
0
= ⎨ −
1.35
M0 − < M0
<
41

Deflection Angle Analysis for Mach 1.6
0.99
0.98
0.97
Pressure Recovery
0.96
0.95
0.94
0.93
0.92
Deflection Angle = 10.75
Pressure Recovery = 0.9765
0.91
0.9
0.89
0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15
Flow Deflection Angle
Figure 6.5 - Optimum Deflection Angle for Mach 1.6 Flow
Pressure Recovery vs. Inlet Shocks
0.99
Total Pressure Recovery
0.97
0.95
0.93
0.91
0.89
0.87
Design Point
Mil-E-5008B
0.85
1 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9 2
Mach Number
2 Shock 3 Shock angles 6 and 6 Mil Spec Mach 1.6
Figure 6.6 - Pressure Recovery for a Two Shock versus Three Shock Inlet
42

Other traits were looked at before choosing a two shock compression inlet. Figure 6.7 illustrates
the area ratio compared to Mach number for the two and three shock inlets. The jumps in the
curve, illustrate the Mach numbers at which the engine will work the hardest because they are
the locations at which the shocks occur. Both of which will occur before the cruise Mach number
of 1.6 therefore occurring during the climb portion of the mission.
Area Ratio vs. Mach
Area Ratio
1.4
1.35
1.3
1.25
1.2
1.15
1.1
1.05
1
0.95
0.9
0.8 1 1.2 1.4 1.6 1.8 2
Mach Number
2 Shock 3 Shock 6 deg and 6 deg
Figure 6.7 - Inlet Area Ratio
Figure 6.8 shows a rough cost relationship between different type inlets. As can be seen there is a
lower cost associated with a two shock inlet. There are drawbacks to having more shocks, in that
they drive the inlet to be larger, longer, and send multiple radar returns. The above traits do not
show enough of a benefit to go with a three shock inlet therefore a two shock inlet was chosen.
43

Figure 6.8 - Cost Association with Inlet Shocks
Figure 6.9 below shows the off design inlet area ratio that is required for the Vendetta. The
equations used to find the data are shown below. The actual capture inlet area is depicted by A 1 ,
with the area at the shock being A s , and the actual flow area being captured by the inlet as A 0i .
As the Vendetta increases in speed the engine requires a greater amount of inlet area for a
constant mass flow rate.
Mass Flow Ratio:
A A A
=
A A A
0i 0i s
1 s 1
Area Ratio:
A
A
0i s s
s
ρ V
=
ρ V
0 0
44

Mass Flow Performance
Inlet Area Ratio
1.3
1.2
1.1
1
0.9
0.8
0.7
Design Point
0.6
0.5
0.4
0.3
0.2
0.1
0
0 0.5 1 1.5 2 2.5 3
Mach Number
Figure 6.9 - Off Design Area Required for Engine Mass Flow
The inlet capture area was found by first estimating the mass flow rate required by the engine at
the design point. The mass flow of the engine could be estimated using the following equation.
Mass Flow Estimation:
m
e
= 26( FrontFaceDiameter)
2
The front face diameter of 4 ft (122 cm) was used; this yielded a mass flow rate of approximately
405 slugs/sec (5910 kg/sec). Now using the mass flow equation shown below, the area of the
inlet could be found for the design mission.
Mass Flow Equation:
m
= ρ AV
Once this was done the mass flow equation was used to calculate the area at different altitudes
based on conservation of energy. For the desired design point of 1.6 Mach and an altitude of
50,000 ft (15,240 m) this was found to be slightly larger than 5 ft 2 (4645 cm 2 ); however at
55,000 ft (16,764 m)t it was found to be about 6 ft 2 (5.6 m 2 ). Since different parts of the mission
take place at several different altitudes above 50,000 ft (15,240 m), the inlet area was sized to 6.5
ft 2 (6 m 2 ). By sizing the engine to 6 ft 2 (5574 cm 2 ) air could be bypassed from the inlet to cool
the fuel.
45

The inlet has a boundary layer diverter for high speeds and auxiliary doors for low speed flight,
since the required inlet area at take off will be twice what it is at cruise. The final inlet sizing for
Mach 1.6 is:
• Inlet capture area = 6.5 ft 2 (6 m 2 )
• Inlet compression angle 10.75 degrees
• Inlet Pressure Recovery is 97.6%
• Speed after Normal Shock, M=0.82
The inlet is located on a boundary layer diverter on the lower side of the wing. This keeps any
vortices produced off of the wing or side of fuselage from being ingested by the inlet, as well as
aid in inlet capture at high angles of attack.
6.3 S-Duct
S-ducts were used to move
the flow from the inlets to
the engine faces so that the
compressor face of the
engine could not be seen.
Stealth is a requirement for
the mission and the
Figure 6.10 - Vendetta S-Duct
compressor face is a large
contributor to radar return. The s-duct goes from a minimum area just after the inlet to a
maximum area at the compressor face as can be seen in Figure 6.10. The s-duct shape
progressively goes from a square at the inlet to an oval
and then a circle at the engine face.
3 o
4 o
The portion of the s-duct closest to the fan face is used
to straighten and slow the flow before it hits the
compressor. This is done by having that portion of the
duct be fairly long and gradually diffuse up to the
compressor face through an upper deflection angle of 3
degrees and a lower angle of 4 degrees as shown in
Figure 6.11.
Figure 6.11 - Diffuser Angle to the Engine
Face
The s-duct is 28 feet (853 cm) in length with an average
duct height of 2.7 feet (82 cm). The efficiency of the s-
duct is found by calculating the average diameter of the
duct and dividing that by the length of the duct. The
figure presented in Appendix B was used to calculate
the efficiency of the duct. This yielded a length over
diameter of just over 10 and an engine area to inlet area of 2 which yielded a duct efficiency of
91.5%.
46

6.4 Nozzle
The nozzle of the Vendetta will have an afterburner and thrust
reversers incorporated. It is a converging-diverging nozzle which
allows for backpressure control at off design Mach numbers
when the afterburners are on. That leads into a low signature
axisymmetric advanced nozzle similar to that seen in Figure
6.12. The advanced nozzle is being used because it has
comparable signature to that of a 2-D nozzle however it weights
50% less, costs 60% less and requires 300 fewer parts.
Figure 6.13 - Clam Style Thrust Reversers
The nozzle will have
thrust reversing
capabilities to enable
the aircraft to land on
an icy runway and
Figure 6.12 - Low-Signature
Axisymmetric Advanced Nozzle
stop within the required 8,000 ft (2483 m) as
specified by the RFP. Clam style thrust reversers,
similar to those seen in Figure 6.13 will be utilized to
reverse 25% of the thrust. Thrust vectoring is will not
be incorporated as the Vendetta is not required to
maneuver like a fighter.
47

7 Materials and Structure
The overall layout of the Vendetta’s structural layout is shown in Figure 7.1. The wing structure
is similar to that of an F-15 and the material selection is similar an F-22. The main load path is
in the form of a central keel that runs from between the nozzles and engines to the nose gear
attachment point. The weapons bay splits the keel in the center of the aircraft. The load is
shifted from the keel to the aft weapons bay wall and back into the keel at the forward end of the
weapons bay. A close up of the weapons bay is also shown in Figure 7.1.
Main Wing Spars
Weapons Bay
Stiffeners
Figure 7.1 - Structure Buildup for Vendetta
48

The layout of Vendetta’s inlets and landing gear allow for a continuous structural member, in the
form of a bulkhead, to carry the aerodynamic loads from each wing directly to the central keel.
This ideology breaks down as the bulkheads move away from the main wing load paths. The
weapons bay splits the forward wing attachment bulkheads. This occurs only well in front of the
aerodynamic center of the wing. Just forward of the aerodynamic center is the main forward
load path for the wing. The aft load paths are a ring structure around the engines and inlets. The
Important thing to note is that where the primary loads are being distributed, between 25 to 50
percent of the mean aerodynamic chord, the bulkheads are continuous.
It is also important to note that the landing gear attaches to a bulkhead just forward of the aft
closure to the weapons bay. This is important because it locates the airborne and ground laden
load paths on top of each other. This allows for some redundancy in the structure and allows for
a lighter aircraft. Another redundant feature is the aft main load path. This bulkhead acts as the
main forward engine attachment point. Again this allows for a minimum of large structural
bulkheads and thus creates a lighter aircraft. The wing attachment points are shown in Figure
7.2.
The empennage structure follows the same methodology as the wing attachment structure. The
vertical tails attach to the aft primary carry through of the wing. The aft vertical attach point is
the same as the primary load path for the horizontal tails. The horizontal tail is an area of
concern for the Vendetta. The horizontal surfaces are capable of producing tremendous forces on
the aircraft. The horizontal surfaces are full flying and thus must attach at a single point in the
aircrafts structure. This bulkhead is a ring carry through type that distributes the load from the
pivot point to the central keel. Two secondary bulkheads back up this main bulkhead. The
empennage structure is shown in Figure 7.3.
Aft Primary
Bulkhead & Main
Engine Attachment
Forward Secondary
Bulkheads
Forward Primary
Bulkhead
Aft Secondary
Bulkheads
Main Gear
Attachment
Figure 7.2 - Wing Attachment Detail
49

Vertical Attachment
Points
Horizontal Pivot
10” Diameter Shaft
Horizontal Structural Load
Paths
Figure 7.3 - Empennage Structural Layout
The structure of Vendetta was created with the RFP load requirements in mind. A V-N Diagram
shown in Figure 7.4 was created using the required maximum and minimum g’ limits, and
knowing the maximum dynamic pressure the aircraft should withstand. This diagram shows the
load envelope the aircraft can operate in. The diagram also shows the standard gust lines for 1-g’
flight.
8
Max g Limit
6
Max q
4
Max Lift
Gust Lines
60 ft/sec
g's
2
0
0 ft/sec
-2
-4
-60 ft/sec
Min g Limit
0 500 300 1000 600 1500 900 2000
1200
Equivlent Knots Equivalent Velocity (ft/sec) Airspeed
Figure 7.4 - V-n Diagram for Vendetta
50

The materials selection for Vendetta was difficult. Vendetta takes advantage of the benefits of
modern composites while relying on the proven durability of more conventional metals. The
materials selection for different components is shown in Table 7.I.
Table 7.I - Materials Selection
Forward fuselage: • Skins and Chine - composite
• Bulkheads / Frames - resin transfer, molded composite, and
aluminum
• Fuel Tank Frame and Walls - resin transfer, molded composite
• Side Array Doors and Avionics - formed thermoplastic
Mid-fuselage: • Skins - composite and titanium
• Bulkheads and Frames - titanium, aluminum, and composite
• Fuel Doors - composite
• Weapons Bay Doors:
o Skins - thermoplastic
Hat Stiffeners - resin transfer, molded composite
o
Aft fuselage • Forward Boom - welded titanium
• Bulkheads and Frames - titanium
• Keel Web - composite / HC Cocure
• Upper Skins - titanium and composite / HC Cocure
Empennage: • Skin and Closeouts - composite
• Core - aluminum
• Spars and Ribs - resin transfer, molded composite
Wings:
• Pivot Shaft - tow-placed composite
The materials used in the construction of the main portion of the wing
are titanium alloy (42%), composites (35%), aluminum alloy (24%),
steel alloy, and some other materials. The following materials are used
in the construction of the wing:
• Skins - composite
• Side of body fitting - titanium HIP cast
• Spars
o front - titanium
o intermediate - resin transfer, molded composite and
titanium
o rear - composite and titanium
Miscellaneous: • Duct Skins - composite
• Landing Gear – Airmet 100 steel alloy
51

8 Landing Gear
Landing Gear design for the Vendetta has eight significant design drivers.
1) Tire selection due to the high 150 knot takeoff and landing speed (277.8 km/hr)
2) 120,000 lb gross weight (54,300 kg)
3) Ease of loading and reloading weapons
4) Tail Strike Angle
5) Ground Handling Characteristics
6) Structural Location
7) Minimal Internal Volume Usage
8) Low Weight
Suitable structural attachment points dictated the main
gear be positioned near the subsonic center of pressure on
the main wing (near the main spar) shown in Figure 8.1.
This placement, as well as limited internal volume, good
ground handling characteristics, minimal frontal area, and
ease of unloading and loading weapons led to the
adoption of a tricycle landing gear configuration. The
main gear configuration was then approximated as a 737
type main gear, (near our GTOW) for volume purposes.
Initial sizing began with tire selection. Following
methodology outlined in Raymer the main gear of the
Vendetta should carry 88% of the GTOW and the nose gear
should carry 12%. Starting with a database of tires and
wheels developed by the Aerospace and Ocean Engineering
Department at Virginia Polytechnic Institute and State
Figure 8.1 - Main Gear Structural
Attachment Point
University and a listing of tires in Raymer’s Aircraft Design a Conceptual Approach the initial
listing was narrowed to the choice of 36in x 11in (91.4cm x 27.9cm) tires for the main gear and
24in x 7.7in (61cm x 19.6cm) for the nose gear. The tires selected allowed a 1.5 factor of safety
over the dynamic landing load of the aircraft.
Now knowing the approximate volume of the 737 landing gear configuration with usable tires a
rough solid model of the fuselage and internal components was produced to determine the exact
gear location. The main gear was then designed to fold into the allotted space. The initial design
considered the smallest internal volume as well as smallest frontal area for a given load, as
shown in Figure 8.2. After analyzing both the internal position the gear would have to fold into,
behind and under the main inlet ducts, the tandem configuration was chosen.
The next challenge presented was obtaining the necessary gear height for easy loading and
unloading as well as a tip back angle which did not exceed the tail strike angle, and having that
gear fit into the limited internal volume available. The gear retraction scheme adopted produced
52

a landing gear retraction scheme similar to an
XB-70 Valkarie as shown in Figure 8.3. The
complexity was necessary due to overall
configuration drive of low supersonic maximum
cross sectional area.
Due to the Sears-Haack ideal of supersonic area
distribution the nose gear could be easily placed
in the forward fuselage. No need for complex
folding arrangements led to the selection of a
standard side-by-side tire configuration. The
complete retraction schemes and nose wheel
configuration can be seen in Figure 8.4. The
completed landing gear configuration can be
seen in Figure 8.5.
The braking system for both the nose gear and
main gear configuration would use a standard
rotor disk braking mechanism. The heat-sink
Figure 8.2 - Landing Gear Configuration Trade
Study
Figure 8.3 - Main Gear Retraction Sequence
Figure 8.4 - Nose Gear and Main Gear Retraction Schemes
will be made of carbon rather than steel because of the fact that carbon is superior to steel as it
has a higher thermal conductivity and temperature limit, and the thermal expansion of carbon is
lower. The only drawback to carbon brakes is the lower density of carbon compared to that of
53

steel. This means that more
braking material is required for
a carbon braking system than
what would be required for a
steel system. The superceding
benefit is that carbon offers a
higher service life and has
lower
maintenance
requirements than steel brakes.
The sizing of the shock
absorption system was
designed around a hydraulic
fluid pressure limit of 1,500 psi
(10.3 GPa). The maximum
load acting on each strut was
then calculated and the
corresponding piston area
required to support this load
was then calculated to be
approximately 7 inches
(17.8cm).
Figure 8.5 - Completed Landing Gear
The Vendetta’s landing gear, as seen
in Figure 8.6 allows for drive up
loading utilizing a MJ-1 or MHU-83
lift truck. Landing gear sizing took
account maximum lift truck reach to
place weapons on the MPRL within
the Vendetta’s main weapons bay.
Figure 8.6 - Vendetta with MJ-1 Lift Truck and 2000lb JDAM
54

9 Weight & Balance
Weight and balance is one area often overlooked in aircraft design. The weights engineer
develops a buildup that heavily influences aircraft performance engineer in terms of wing sizing,
propulsion system selection, and in predicting aircraft performance. The stability and control
engineer must rely on the weights engineer in order to size control surfaces, to develop flight
control systems, and to develop stability derivatives. The configurator must work carefully with
the weights engineer in order to develop a feasible inboard layout. In all aspects, the weights
engineer plays a major role coordinating the design of any aircraft.
After having sized the aircraft using the weight fractions technique, and after having developed
an initial configuration, the next step is to develop a more accurate, class II weight buildup of the
aircraft. The class II method used in the design of the Vendetta was developed from those
methods found in the Nicolai, Raymer, and Roskam texts in order to obtain a collaborative and
unbiased perspective. These methods involved defining several physical and geometric
parameters of the aircraft. These parameters were inputs into a series of equations developed
from historical weight trends. The weight estimations for various components as well as the
level of agreement between authors are shown below in Table 9.I.
Table 9.I - Initial Component Weight Buildup
Roskam Nicolai Raymer Average Roskam Nicolai Raymer
Component
Weight
(lbs)
Weight
(lbs)
Weight
(lbs)
Weight
(lbs)
Accuracy
(%)
Accuracy
(%)
Accuracy
(%)
Structures
Wing Group 9,687 11,466 7,870 9,674 0% -31% 26%
Horizontal Tail 1,135 1,694 958 1,262 14% -62% 32%
Vertical Tail 801 1,538 1,497 1,279 47% -34% -28%
Fuselage 10,681 16,031 10,398 12,370 19% -52% 22%
Main Landing Gear 2,742 2,969 1,156 2,289 -33% -52% 60%
Nose Landing Gear 387 405 408 400 5% -2% -3%
Propulsions 10,636 10,878 11,199 11,209 4% 0% -4%
Systems 18,649 14,506 14,350 20,574 -29% 12% 13%
Payload 9,500 9,500 9,500 9,500 0% 0% 0%
Fuel 58,974 58,974 58,974 58,974 0% 0% 0%
TOGW 123,653 128,435 128,435 127,531 4% -2% -2%
The detailed weight buildup of the structures, control surfaces, systems, payload, and fuel groups
has been compacted in order to save space and can be viewed in its entirety in Foldout 1. The
table indicates that all three authors tend to disagree to some extent in their weight estimates of
certain components, and for other components, one author may have no way of estimating that
components weight at all. The most accurate way to develop the most detailed component weight
buildup was to consider all three methods and take the average shared between them. An
author’s estimation was discarded if it did not agree to within ±30% of the average of the other
authors’ estimations. Once the estimations were in agreement to within ±30%, they were
averaged in order to develop a weight buildup for the entire aircraft. The class II weight buildup
for the Vendetta after the downgrading process is shown below in Table 9.II.
55

Table 9.II - Final Component Weight Buildup
Roskam Nicolai Raymer Average
Component
Weight
(lbs)
Weight
(lbs)
Weight
(lbs)
Weight
(lbs)
Structures
Wing Group 9,687 XXXXXX 7,870 8,779
Horizontal Tail 1,135 1,694 958 1,262
Vertical Tail 801 1,538 1,497 1,279
Fuselage 10,681 XXXXXX 10,398 10,540
Main Landing Gear 2,742 2,969 1,156 2,289
Nose Landing Gear 387 405 408 400
Propulsions 10,636 10,878 11,199 11,209
Systems 18,649 14,506 14,350 20,574
Payload 9,500 9,500 9,500 9,500
Fuel 58,974 58,974 58,974 58,974
TOGW 124,800
Inertias were calculated using guidelines outlined by the Society of Allied Weight Engineers
(SAWE). Each components mass and location in reference to the aircraft center of gravity was
used to calculate that components inertia. The sums of these inertias were then used in
calculating the total moments of inertia about the
Vendetta’s principal axes shown in Figure 9.1. In order to
determine whether or not these values were accurate, the
moments of inertia were transformed into non-dimensional
radii of gyration coefficients. These coefficients were then
compared to typical values for a jet bomber provided by
SAWE. These comparisons are shown below in Table
9.III.
Table 9.III - Inertia Estimation
Figure 9.1- Principle Axes
Inertias (slug*ft 2 )
Ix Iy Iz
Vendetta 35,248 807,819 838,575
Nondimensional Radii of Gyration
Rx Ry Rz
Vendetta 0.11 0.29 0.38
SAWE 0.31 0.33 0.47
Accuracy 63% 13% 18%
The table indicates that the inertias are well
within the typical values for a jet bomber
except about the roll axis. This is because the
Vendetta is similar to a typical jet bomber in
length; however, it has a much shorter
wingspan. This would constitute a smaller
moment of inertia about the roll axis.
After having developed an initial configuration and a more detailed class II weight buildup, the
next step was to balance the aircraft. This was done for two types of payload, the first being
fixed equipment and the second being variable payload. The variable payload aboard the
Vendetta consists of both fuel and weapons because the weight and center of gravity of these
56

items varies throughout the mission. The fixed equipment aboard the Vendetta is considered in
this case to be everything other than the variable payload.
Vendetta was first balanced it with the fixed equipment and then with the additional variable
payload. This was done by first allowing the configurator to develop an inboard configuration.
The weights engineer then calculated the center of gravity location resulting from this inboard
arrangement. This process was iterative in that the weights engineer and configurator had to
continuously modify the inboard arrangement until the center of gravity location was at the
desired location.
In order to minimize the trim drag on the aircraft, it was opted that the aircraft’s center of gravity
location stay as close to the aerodynamic center as possible. This was a difficult task because of
the dramatic shift, 12% MAC, in the location of the aerodynamic center when transitioning from
subsonic to supersonic flight conditions. A trim tank was considered in order to allow the center
of gravity to follow the aerodynamic center during this dramatic shift in order to maintain a
neutrally stable condition at both subsonic and supersonic flight conditions; however, this idea
was discarded because the trim tank would only require additional fuel volume in an already
congested aircraft. Without a trim tank, in order to minimize drag by keeping the center of
gravity as close as possible to the aerodynamic center the aircraft would have to fly with an
unstable static margin, subsonically, and with a stable static margin, supersonically.
The current arrangement of the fixed payload is such that it provides for a 5% unstable static
margin at subsonic flight conditions. A center of gravity monitor makes use of fuel burn control
in order to keep the aircraft as close as possible to the neutrally stable flight condition.
Furthermore, with full fuel tanks, full weapons load, and subsonic flight conditions, i.e. takeoff,
the aircraft is balanced such that it provides for a 5% unstable static margin. With the empty
weight and takeoff gross weight balanced to provide a 5% unstable static margin, and with an
aerodynamic shift of 12%, the aircraft immediately goes to a 7% stable static margin upon
transitioning to supersonic flight. The center of gravity monitor then controls the fuel burn in
such a way that the center of gravity follows the aerodynamic center and the Vendetta maintains
neutral stability.
Because the fuel load is constantly changing throughout the mission, balancing the fuel load
throughout the mission can be a challenging task. Furthermore, the deployment of various
weapons at any point during the mission makes this balancing process even more difficult.
Because of the complexity involved in developing a center of gravity monitor, a computer code
was developed in order to simulate the center of gravity monitor. The first step in developing
this code was to obtain the best solution to balance the fuel payload throughout the mission. The
code required four inputs including; the weight and location of the fixed equipment, the location
and weight of the fuel at any given time, the amount of fuel burned at intervals throughout the
mission profile, and the desired center of gravity location at that interval. With these inputs, the
code can then determine which tank to burn fuel from in order to obtain the center of gravity
location closest to that corresponding to the desired static margin. The code then outputs the
center of gravity location and the remaining fuel payload. This is done at 10-second intervals
57

throughout the 5-hour mission. Using this data, the center of gravity path can then be plotted as
it tracks that path corresponding to the desired static margin.
The next step was to balance the weapons payload. Because the weapons payload was placed in
a rotary launcher, the center of gravity of the payload was concentrated in one location. If it had
been placed in a more conventional arrangement spread across the belly of the aircraft, the center
of gravity of the weapons would have also been spread across the belly of the aircraft. By
concentrating the center of gravity of the weapons payload in one location and placing the
weapons payload on top of the aircraft’s empty weight center of gravity location, deployment of
the weapons payload did not generate any problems in balancing the aircraft or in disturbing the
static margin. The center of gravity is shown tracking along the path of the desired static margin
by means of fuel monitoring and pumping in Figure 9.2.
Figure 9.2 - Center of Gravity Excursion
The figure indicates that the center of gravity location at takeoff gross weight is slightly aft of
the neutral point; however, the center of gravity tracks the desired static margin shortly after
the aircraft has transitioned to supercruise. Notice the path of the aerodynamic center as it
shifts during the transition from subsonic to supersonic flight. It is clear that the aircraft flies
supersonically shortly after takeoff, or when the aircraft’s gross weight is just below takeoff
gross weight. Furthermore, near the zero fuel weight, the aircraft flies subsonic for the
remainder of the flight. The figure also indicates that with the current fuel tank arrangement,
the desired static margin cannot be tracked during the final portion of the supercruise because
there is not enough fuel available to properly trim the aircraft. At this point, the center of
58

gravity is influenced by only the fixed weight of the aircraft and again the aircraft remains at
a 5% unstable static margin during landing. This plot indicates that the center of gravity
monitor works together with the control system in order to minimize trim drag while at the
same time maintaining the aircraft’s controllability.
59

10 Stability and Control
To initially size the horizontal tail, tail volume coefficients from historical aircraft were
analyzed. This was done in an attempt to determine the rough size of the horizontal and vertical
tail surfaces prior to addressing stability and control issues. The tail volume coefficients are
unitless parameters defined by geometric values relating the size of the empennage surface to the
aircraft. The horizontal and vertical tail volume coefficients are defined in the following
equations.
V
H
=
SHTL
c S
W
HT
W
V
V
=
SVT
L
b S
W
VT
W
Because the demands for most supersonic cruising aircraft are considered similar to a certain
extent, the historical values of tail volume coefficients are used to back out the planform areas
for the horizontal and vertical surfaces. Similar aircraft and their tail volume coefficients are
presented in Table 10.I.
AIRCRAFT
Table 10.I - Historical Aircraft Tail Volume Coefficients
TAIL VOLUME COEFFICIENTS
V H
V V
Boeing SST (2707-300) 0.36 0.049
Concorde n/a 0.080
GD F-111A 1.28 0.064
Rockwell B1B 0.80 0.039
TU-22M 1.11 0.087
TU-144 n/a 0.081
AVERAGE 0.58 0.067
Using the average tail volume coefficient for these similar aircraft yielded a horizontal stabilizer
area of 386 ft 2 . This is rather large and may be attributed to the fact that these vehicles require
large robustness in CG travel without the use of a flight control augmentation system (CAS).
Likewise, the vertical tail would require 196 ft 2 of area. This number is driven slightly larger due
to the fact that some of the larger historical tail volumes are inflated because these aircraft’s
verticals are mounted on booms which extend aft. These booms allow for greater moment arms
and make the vertical more effective.
The effects of horizontal tail area on longitudinal static stability were looked at in an attempt to
determine what the driving factors for horizontal tail area are. A Roskam class II method was
60

used to see how the increased weight of a bigger horizontal affects the longitudinal static margin.
This “X-plot”, as it is commonly known, is shown as Figure 10.1.
Figure 10.1 - Longitudinal X-Plot at Mach 0.3
It is notable that only about 110 ft 2 of horizontal area is required to keep the Vendetta neutrally
stable at Mach 0.3. It can be seen that as the tail grows, the CG of the entire configuration shifts
aft. This also shifts the effective neutral point (center of pressure) of the aircraft aft at a faster
rate than the CG shifts aft. A horizontal that is bigger than 100 ft 2 yields a stable aircraft but will
pay the price in trim drag if the aircraft is too stable.
A stable static margin is necessary in flight without the use of a digital flight control margin. The
RFP mandates this as well as adherence to MIL-8785C, the military specification for handling
qualities of aircraft. A statically unstable aircraft would have a tendency to pitch up in a static
level condition. The purpose of the horizontal tail is to apply a force which counteracts this
offending moment. This comes at the price of trim drag, however. As the elevator is deflected,
drag is created and this hurts the overall aircraft performance in cruise. It is because of this that a
neutrally stable or marginally stable (1-3%) aircraft is desired in cruise.
To complicate matters, it can be seen from Figure 6.1 that areas above 110 ft 2 are required for a
Mach number of 0.3. The aerodynamic center (center of pressure) on the wing and most surfaces
propagates aft as the Mach number passes the transonic regime. This shift effectively leaves the
61

difference in neutral point and center of gravity greater. This difference means the aircraft is
actually more stable in a supersonic cruise. The fact that the center of gravity is so far forward in
relation to the neutral point causes the aircraft to pitch down. More trim is required which causes
drag. This phenomenon is known as Mach tuck.
It is because of this that the weight and balance of the aircraft must be closely in synch with the
control system. Trim drag will be minimized and controllability will be enhanced with
completely integrated systems.
Canting the horizontals in a v-tail configuration was investigated in an attempt to shape the
empennage in a stealthy manner. The effective area of the vertical and horizontal are functions of
the square of the cosine of the cant angle. These effects are reflected in Figure 10.2.
Figure 10.2 - Horizontal Area Required for Static Stability with Cant Angle
It can be seen from the plot that as the cant angle increases the total planform area of the
horizontal must increase to maintain the nominally desired static stability of 5%. Five percent
was chosen because at this stage in the sizing it was uncertain what the dynamic characteristics
of the aircraft would be. Attempting to maintain a minimally statically stable aircraft would ease
the job of control system design if future needs warrant one. Angles up to 30° were looked at
because it would be unwise from an RCS point of view to approach a 90° angle created by larger
cants near 45°. Beyond 45° the trend would be the same; however the horizontal would drive the
area instead of the vertical.
This plot shows that only 118 ft 2 of horizontal area is required to maintain the desired static
margin. This is far off from the historical class I method and by initial inspection appears small.
62

This leads us to believe that the area required maintaining static stability is not the driving factor
in the size of the horizontal. Control power required to rotate the aircraft, dynamic
considerations, and high angle of attack recovery will most likely drive this size.
A similar study was conducted on the vertical stabilizer to see what area would required for
varying cant angles to maintain 0.001 (1/degree) lateral weathercock stability. This is illustrated
in Figure 10.3.
Figure 10.3 - Vertical Area Required for Static Stability with Cant Angle
From Figure 6.3 it can be seen that at 30°, 165 ft 2 of vertical area is required to maintain 0.001
(1/degree) of lateral weathercock stability. Although the 30° cant angle on the verticals was
initially selected to match the bottom fuselage facets for RCS considerations, lowering that angle
to 20° would allow other advantages. Shallower cant angles are easier to manufacture, require
less structure, weigh less, and have less coupling with pitch modes. For these reasons, the impact
on RCS was investigated for the 20° cant angle as well as the pitch coupling term for rudder
deflection, C
δ
.
m r
The RCS code was run on two aircraft configurations. The same wing, fuselage, and horizontal
were modeled with the vertical planforms mounted at both 20° and 30°. The results of that study
are shown as Figure 10.4.
63

40
30
20
10
0
-10
-20
-30
-40
-50
20° Canted Vertical
30° Canted Vertical
RFP Requirement
Figure 10.4 - Radar Cross Section Impact of 20° vs. 30° Vertical Cant Angle
Figure 69 clearly shows that there is an impact on the RCS for changing the cant angle. The RFP
required -12 dBm 2 return is shown in red. As mentioned in the RCS section, this requirement is
only mandated for the frontal 0° azimuth angle. Going to a 20° cant does not violate this
requirement and yields the aforementioned benefits.
The effective area of a rudder sized to 27% mean aerodynamic chord of the vertical was
calculated in the horizontal plane of the aircraft. In normal non-canted configurations, Cm
δ
is
r
nonexistent. Table 10.II shows the values for this coupling term and various cants.
Table 10.II - Pitching Moment Coupling with Rudder Deflection for Various Vertical Cant Angles
C
δ
Vertical Cant Angle
(165 ft 2 27% m.a.c. Rudder) m r
0° 0.0000
10° 0.0004
20° 0.0009
30° 0.0021
64

The extra 10° cant resulted in a substantially larger pitch coupling term. In addition to the
complications of canting more, a 30° angle would mean that a more complex mixer and control
system would be required. This would add to the cost and is avoided.
It is important to note that the previous static methods do not take into account the dynamic
characteristics or modes of this aircraft. With such a large amount of the fuselage in front of the
center of pressure, the Vendetta may require a complex yaw damper or larger vertical to
compensate.
The size of the vertical could potentially be driven by the one engine inoperative (OEI) control
power requirements. Because the engine nozzle centerlines are mounted considerably offset from
the centerline at 3 feet (0.9144 m), a large yawing moment will be created if the Vendetta loses
an engine during takeoff. The engines produce roughly 45,000 pounds of thrust and would
generate a 135,000 foot-pound moment. Table 10.III shows the results of the rudder control
power analysis for this critical OEI condition.
Table 10.III - Rudder Control Power Results for OEI Condition
PARAMETER NOTATION VALUE
Side Force due to Rudder C yδr 0.0105
Rolling Moment due to Rudder C lδr 0.0072
Rudder Effectiveness C nδr -0.0070
OEI Critical Yawing Moment
135,000 ft-lbs
Rudder Deflection Required in OEI Condition at Takeoff 13.6°
With a rudder effectiveness of -0.0070 (1/deg), a 13.6° rudder deflection is required to keep the
aircraft flying straight in the OEI condition on takeoff. This is not too large, and would suffice by
allowing approximately another 10° of rudder deflection for the pilot to yaw the aircraft beyond
the straight condition for controllability. In this condition, the aircraft would be susceptible to
large amounts of sideslip, β.
This rudder deflection would be substantially higher if a higher cant angle were used. In these
critical situations where the aircraft is in danger, the added drag created by the mixing is desired
to be as little as possible.
A separate 4 surface empennage was now made necessary because going to a purely V-tail was
shown to be ill-advised at this stage because of the aforementioned studies. If a pure v-tail was
chosen, it would have to be full-flying due to the demand placed on the surface and hinge lines in
supersonic flight. This would require a large actuator and large structural members in the aft
portion of the aircraft. This would considerably drive the configuration away from initial RCSfriendly
layouts as well as increasing complexity and cost.
65

The Vendetta configuration utilizes a 20° cant on the
verticals and a separate full-flying horizontal as seen
in Figure 10.5. It was mentioned earlier that one of
the reasons the horizontal tail volume coefficient
was larger in the historical aircraft was because
those aircraft did not utilize control augmentation
systems or digital fly-by-wire control systems. Not
only did they have to account for wide shifts in CG,
they also had to combat the muck tuck problem
associated with breaking the sound barrier.
Figure 10.5 - Vendetta Empennage Configuration
Figure 10.6 shows that as the aircraft
Mach trim
exceeds the critical Mach number,
the center of pressure of the wing
and other control surfaces travels aft.
In the case of the Vendetta, this
leaves the CG an extra 12% m.a.c. in
front of the neutral point; this makes
it 12% more stable. This 12% shift
was calculated with the Air Force’s
Data Compendium (DATCOM)
mg
12% m.a.c.
methods.
Figure 10.6 - Mach Tuck Illustrated
The shift in the neutral point of the
wing means that the horizontal
would have to deflect to keep the Vendetta from “tucking” under. The trim drag created could be
avoided by shifting the CG, by altering the neutral point, or designing the aircraft to be unstable
subsonic and stable supersonic.
The use of a trim tank was investigated to pump fuel aft and shift the CG closer to the neutral
point in supersonic cruise. This notion was dismissed because the tank would be a vacant waste
of space and would complicate ground procedures where refueling would have to leave the tank
vacant.
The use of an extra flying surface such as a canard could be used as well. The canard would
destabilize the aircraft by moving the neutral point forward and closer to the CG but it would
make the Vendetta even more uncontrollable in the subsonic landing and takeoff conditions. This
extra control surface would add to the cost and complexity.
A fuel management system could be used to burn fuel from certain tanks to keep the CG travel in
check. After analyzing the abrupt shift in the neutral point when the Vendetta climbs to its cruise
condition, it was decided that the fuel management system could not pump fuel fast enough to
66

trim the aircraft. The likewise was true when decelerating. The aircraft would suddenly go
unstable. Because of this, use of a digital flight control system (DFCS) which is provided as GFE
would allow the aircraft to fly unstable subsonic. The DFCS could easily allow a 0% - 7%
unstable aircraft takeoff and land.
The wing was placed and the empennage sized for the Vendetta to be 5% unstable in the
subsonic regime and 7% stable in the supersonic regime without CG modification due to the
12% shift. The fuel management system could then be used to enhance cruise performance by
pumping fuel in a way which results in neutral or marginal static stability.
A DFCS will not impact the design too much because complex navigation and autopilot systems
will already have to be incorporated into the design. In addition to this, the DCFS will be used to
enhance the dynamic modes of the aircraft. These may require it due to the large fore body and
unstable pitch break exhibited by the Vendetta. Also, the 2010 delivery date will mean that next
generation control laws and hardware could be implemented. All modern fighters being designed
today utilize such systems already. The DFCS along with the fuel management system would
maintain the static and dynamic stability.
DATCOM and the compiled Digital DATCOM Fortran code proved to be useful tools in
calculating many of the aerodynamic stability and control derivatives for the Vendetta. This was
done in an attempt to identify problematic behaviors and to adhere to MIL-8785C.
It was calculated that the Vendetta’s fuselage fore body will destabilize the aircraft an additional
3.1% in subsonic cruise and 5.0% in supersonic cruise. The wing was placed to account for this.
This is much improved over previous configurations where the fuselage destabilized the aircraft
up to 16%. This is due to the fact that so fuselage with a large mean width was in front of the CG
and NP.
Figure 10.7 shows the Vendetta’s pitch break characteristics in the subsonic low speed and
supercruise regimes given a CG location that would yield a statically stable aircraft.
67

-0.6
-0.4
UNSTABLE
NEUTRAL
-0.2
Lift Coefficient (C L )
-0.25 -0.2 -0.15 -0.1 -0.05 0 0.05 0.1 0.15 0.2 0.25
0
0.2
0.4
M = 0.2
M = 1.6
0.6
0.8
Moment Coefficient (C m )
Figure 10.7 - Pitch Break Characteristics
This figure shows that as the Vendetta rotates and has some angle of attack in the low speed
subsonic (M=0.2) regime, it will want to continue to rotate and break away. In the supercruise,
the aircraft behaves mush more linearly. The subsonic characteristics are of some concern, but
even simple feedback schemes in the DFCS could solve this problem. The supersonic
characteristics are actually more desirable because the maneuvering required is very light and the
control system will not be oscillating the control surfaces, which creates unnecessary drag, to
keep the aircraft flying straight.
A full state-space based model for the aircraft driven by a Taylor expansion and fit into equations
of motion was developed for flight simulator validation. These forms are too complex for simple
dynamic analysis, so the literal factor forms of the dynamics modes were used to determine
flying qualities and conformity with MIL-8785C.
The literal factors are nothing more than simplifications of the transfer function forms for
longitudinal and lateral modes of interest. These forms omit insensitive stability derivatives. The
literal factor forms for the modes of interest are provided in the Appendix. The conformity with
the military specifications for handling quality is shown in Table 10.IV.
68

Mode
Table 10.IV - Longitudinal and Lateral Dynamic Mode Conformity with MIL-8785C
Damping ratio (ζ) Natural Frequency (ω n )
Vendetta MIL-8785C Vendetta MIL-8785C
MIL-8785C
Level
Phugoid 0.094 > 0.04 0.091 - I
Short Period 0.921 0.35 – 1.3 4.721 - I
Dutch Roll 0.103 > 0.08 1.960 > 0.4 I
Table 10.IV shows that the Vendetta satisfies all of the military specifications for these three
important modes while in a subsonic cruise with the CG monitor. The only thing of concern
regarding these results is high value for undamped natural frequency in the Dutch Roll mode. It
is not uncommon for aircraft of this size and type to incorporate fairly simple yaw dampers
operating on the yaw rate. With the use of the DFCS, the Vendetta would have no problem
keeping that mode in control. Because there is a large amount of robustness available with CG
excursion and the DFCS, the longitudinal modes are well within the Type I military
specifications and would remain there in the supercruise.
69

Validation of an aircraft design such as this is fairly hard to accomplish with traditional models.
Traditional wind tunnel testing is limited to the availability of supersonic wind tunnels. Because
of this, the Cal Poly Pheagle Flight Simulator was used to evaluate handling qualities, ground
handling due to landing gear placement, up-and-away tasks, and low speed performance.
A non-linear table lookup scheme was incorporated in the computer simulation to lookup the
DATCOM derived stability and control derivatives. These derivatives were used to develop
tables for the major derivatives as functions of Mach number and angle of attack. The simulator
and flight cab are shown as Figure 10.8 and Figure 10.9, respectively.
Figure 10.8 - Pheagle Simulator
Figure 10.9 - Flight Cab and Instruments
The static flight cab allows the pilot to see the a full panel of operational instruments as well as
feel up to 50 pounds of force on the stick. The stick used was that from an F-15 Eagle. Graphics
models were incorporated for pilot cues and situational awareness as seen in Figure 10.10. A full
HUD was also incorporated and used as illustrated in Figure 10.11. Stick forces and dynamics
were approximated based on existing information from current fighter aircraft.
Figure 10.10 - Graphics and Environment
Figure 10.11 - Heads up Display
Results of validation proved that the DFCS control laws would need to be developed for the
aircraft to be flyable in all flight regimes.
70

11 Performance
11.1 Performance Requirements
The performance requirements listed in the RFP, shown in Table 11.I, consist of specific excess
power requirements, a turn rate requirement, mission requirements, and takeoff and landing
requirements. The specific excess power and turn rate requirements are measured at maneuver
weight, which is defined as 50% internal fuel, and a payload of 2 × AIM-120 AMRAAM’s and 4
× 2,000 lb (907 kg) JDAM’s. The mission requirements are defined as the ability to perform the
design mission listed in Table 11.II.
Table 11.I - RFP Performance Requirements
Supercruise Mission Radius
1,750 nm (3,240 nm)
Supercruise Mach Number 1.6
1g Military Specific Excess Power at 50,000 ft & Mach 1.6 0 ft/s (0 m/s)
1g Maximum Specific Excess Power at 50,000 ft & Mach 1.6 200 ft/s (61.0 m/s)
2g Maximum Specific Excess Power at 50,000 ft & Mach 1.6 0 ft/s (0 m/s)
Maximum Instantaneous Turn Rate at 15,000 ft & Mach 0.9 8.0 deg/s
Takeoff Field Length 8,000 ft (2,438 m)
Landing Field Length (Icy) 8,000 ft (2,438 m)
Table 11.II - RFP Design Mission
1. Takeoff and acceleration allowance
a. Fuel allowance for warm-up
b. Fuel to accelerate to climb speed at maximum thrust (no distance credit)
2. Climb from sea-level to optimum supercruise altitude
3. Supercruise 1,000 nm (1,852 km) at Mach 1.6 and optimum altitude
4. Climb above 50,000 ft (15,240 m)
5. Dash 750 nm (1,389 km) at Mach 1.6 above 50,000 ft (15,240 m)
6. Weapons delivery
a. 180º turn at 50,000 ft (15,240 m) and Mach 1.6
b. Drop air-to-surface weapons
7. Dash 750 nm (1,389 km) at Mach 1.6 above 50,000 ft (15,240 m)
8. Supercruise 1,000 nm (1,852 km) at Mach 1.6 and optimum altitude
9. Descend to sea-level (no distance credit or fuel used)
10. Reserve for 30 min at sea-level and speed for maximum endurance
The RFP design mission explicitly defines some aspects of the required mission, while other
aspects of the mission such as cruise altitudes and loiter speed are arbitrary. Within the
constraints of the design mission, a detailed mission was created and optimized to minimize fuel
71

consumption. The main aspects of the mission that were optimized were the initial climb
sequence, the cruise and dash altitudes (dash altitude must be greater than 50,000 ft (15,240 m)),
and the loiter speed. The optimum climb sequence was found by creating a flight envelope with
lines of constant climb rate to fuel flow ratio (dh/dW F ) at the average climb weight of the
aircraft. The climb profile that minimizes the fuel required to climb the aircraft to a given cruise
condition is then found by drawing a flight path between the initial and cruise conditions that
follow the maximum climb rate to fuel flow ratio. The resulting flight path and fuel
consumption envelope are shown in Figure 11.1.
50,000
dh /dW F = 0 ft/lb
100 ft/lb
40,000
Stall Limit
50 ft/lb
Altitude (ft)
30,000
20,000
100 ft/lb
150 ft/lb
200 ft/lb
250 ft/lb
10,000
300 ft/lb
q Limit
0
0 0.5 1 1.5 2
Mach
Figure 11.1 - Fuel Consumption Envelope at Average Climb Weight
The optimum cruise and dash altitudes were found by running a series of missions at different
altitudes and finding the mission with the lowest fuel consumption. Because the aircraft’s
weight decreases as fuel is burned, the optimum cruise altitude increases over the mission
profile. It was found that the optimum sequence of cruise altitudes began at 52,000 ft (15,850 m)
for the initial cruise and increased by 2,000 ft (610 m) for each successive cruise or dash segment
resulting in a final cruise altitude of 58,000 ft (17,678 m). The optimum loiter speed was found
as the speed at which the minimum drag occurred under loiter conditions (sea level and 61,000 lb
(27,669 kg) weight). The drag on the aircraft under these conditions is plotted as a function of
Mach number in Figure 11.2 showing that the minimum drag occurs at Mach 0.35 or 391 ft/s
(119.2 m/s). The resulting detailed mission is listed in Table 11.III.
72

30,000
25,000
Wave Drag
Drag (lb)
20,000
15,000
Sea-Level
Loiter Weight
10,000
5,000
Induced Drag
Parasite Drag
0
0 0.2 0.350.4 0.6 0.8 1 1.2
Mach
Figure 11.2 - Drag on Aircraft in Loiter Conditions
Table 11.III - Detailed Design Mission
1. Warm-up 2 min at idle thrust
2. Takeoff – Accelerate to takeoff speed 266 ft/s (81.1 m/s)
3. Accelerate to Mach 0.86 at maximum military thrust
4. Climb to 15,000 ft (4,572 m) and accelerate to Mach 0.9
5. Climb to 17,500 ft (5,334 m) and accelerate to Mach 1.25
6. Climb to 33,000 ft (10,058 m) and accelerate to Mach 1.72
7. Climb to 45,000 ft (13,716 m) at Mach 1.72
8. Climb to 52,000 ft (15,850 m) and decelerate to Mach 1.6
9. Cruise 1,000 nm (1,852 km) at 52,000 ft (15,850 m) and Mach 1.6
10. Climb to 54,000 ft (16,459 m) at Mach 1.6
11. Dash 750 nm (1,389 km) at 54,000 ft (16,459 m) and Mach 1.6
12. Descend to 50,000 ft (15,240 m) at Mach 1.6
13. Turn 180º at n = 1.55
14. Drop 4 × 2,000 lb (907 kg) JDAM’s
15. Climb to 56,000 ft (17,069 m) at Mach 1.6
16. Dash 750 nm (1,389 km) at 56,000 ft (17,069 m) and Mach 1.6
17. Climb to 58,000 ft (17,678 m) at Mach 1.6
18. Cruise 1,000 nm (1,852 km) at 58,000 ft (17,678 m) and Mach 1.6
19. Descend to Sea-Level and decelerate to loiter speed 391 ft/s (119.2 m/s)
20. Loiter 30 min at Sea-level and 397 ft/s (121.0 m/s)
21. Decelerate to landing speed 266 ft/s (81.1 m/s)
22. Land – decelerate to zero speed
23. Unload non-fixed equipment (2 × AMRAAM’s and crew 1,280lb (5,806 kg))
73

In addition to the performance requirements, the RFP includes performance measures of merit
listed in Table 11.IV. The measures of merit are not specific requirements, but are measures by
which competing aircraft will be judged.
Table 11.IV - Performance Measures of Merit
Mission duration, radius, and fuel burn by mission segment for design mission
Takeoff and landing distance for design mission on dry and icy runways at sea-level
Maximum Mach number at 36,000 ft (10,973 m)
1g Maximum thrust specific excess power envelope
2g Maximum thrust specific excess power envelope
5g Maximum thrust specific excess power envelope
Maximum thrust sustained load factor envelope
Maximum thrust maneuvering performance diagrams at sea-level and 15,000 ft (4,572 km)
11.2 Specific Excess Power Requirements
Compliance with the specific excess power requirements is best shown using specific excess
power envelopes. Figure 11.3 shows the 1g military specific excess power envelope. The RFP
requirement of 0 ft/s (0 m/s) at Mach 1.6 and 50,000 ft (15,240 m) is met with 95.7 ft/s (29.2
m/s) specific excess power. The 1g maximum (afterburner) specific excess power envelope,
shown in Figure 11.4, shows that the RFP requirement of 200 ft/s (61.0 m/s) at Mach 1.6 and
50,000 ft (15,240 m) is met with a value of 316 ft/s (96.3 m/s). This envelope also shows that
the maximum Mach number at 36,000 ft (10,973 m) is 2.1. The 2g maximum specific excess
power envelope, shown in Figure 11.5, shows that the RFP requirement of 0 ft/s (0 m/s) at Mach
1.6 and 50,000 ft (15,240 m) is met with a value of 108 ft/s (32.9 m/s).
74

70,000
60,000
50,000
Stall Limit
P s = 0 ft/s
RFP Requirement 0 ft/s
P s = 200 ft/s
Altitude (ft)
40,000
30,000
P s = 100 ft/s
Flaps
P s = 200 ft/s
20,000
P s = 300 ft/s
q Limit
10,000
P s = 400 ft/s
0
0 0.5 1 1.5 2 2.5
Mach
Figure 11.3 - 1g Military Specific Excess Power Envelope at Maneuver Weight
70,000
P s = 0 ft/s
60,000
Stall Limit
50,000
RFP Requirement 200 ft/s
Altitude (ft)
40,000
30,000
20,000
Flap
P s = 200 ft/s
P s = 400 ft/s
P s = 600 ft/s
P s = 600 ft/s
q Limit
10,000
P s = 800 ft/s
0
0 0.5 1 1.5 2 2.5
Mach
Figure 11.4 - 1g Maximum Specific Excess Power Envelope at Maneuver Weight
75

70,000
Altitude (ft)
60,000
50,000
40,000
30,000
20,000
10,000
Stall Limit
P s = 200 ft/s
P s = 400 ft/s
P s = 600 ft/s
P s = 0 ft/s
RFP Requirement 0 ft/s
P s = 600 ft/s
q Limit
0
0 0.5 1 1.5 2 2.5
Mach
Figure 11.5 - 2g Maximum Specific Excess Power Envelope at Maneuver Weight
11.3 Turn Rate Requirement
The maximum instantaneous turn rate requirement of 8.0 deg/s at 15,000 ft (4,572 m) and Mach
0.9 is shown in the maneuverability diagram at 15,000 ft (4,572 m) in Figure 11.6. The
maneuverability diagram shows that the required turn can be sustained with military power. The
maximum sustainable turn rate under maximum power is 13.1 deg/s.
76

30
n
2
3
4 5 6 7
r = 2,000 ft
25
r = 4,000 ft
Turn Rate (deg/s)
20
15
10
5
Stall Limit
Max. P s = 0
Mil. P s = 0
RFP Requirement
8 deg/s
r = 6,000 ft
r = 8,000 ft
r = 10,000 ft
q Limit
0
0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2
Mach
Figure 11.6 - Maneuverability Diagram at 15,000 ft (4,572 m) and Maneuver Weight
77

11.4 Mission Requirements
By completing the design mission, the requirements for a supercruise Mach number of 1.6 and
mission radius of 1,750 nm (3,240 km) are met. To determine the fuel capacity required to
perform the design mission, the mission was simulated by numerically integrating the fuel burn
rates over the mission profile. The mission simulation also allowed certain aspects of the
mission such as cruise altitudes to be optimized. Table 11.V lists the results of the mission
simulation, Figure 11.7 shows a breakdown of fuel consumption by mission segment, and Table
11.VI lists the fuel consumption by mission segment.
Reserve
4%
Misc.
5%
Accelerate &
Climb
12%
Cruise Back
18%
Cruise Out
28%
Dash Back
15%
Dash Out
18%
Figure 11.7 - Fuel Consumption over Mission
Table 11.V - Mission Results
Total fuel consumption
57,940 lb (26,280 kg)
Mission radius
1,750 nm (3,240 km)
Total distance traveled over mission 3,980 nm (7,370 km)
Total mission duration
4 hr. 57 min.
Takeoff weight
124,800 lb (56,610 kg)
Empty weight
56,270 lb (25,520 kg)
Fuel weight (total fuel onboard) 59,250 lb (26,880 kg)
Maneuver weight
95,830 lb (43,470 kg)
Landing weight
57,550 lb (26,100 kg)
Average cruise lift to drag ratio 6.6
78

Table 11.VI - Fuel Consumption by Mission Segment
1. Warm-up 2 min at idle thrust 83 lb (37.6 kg)
2. Takeoff (Accelerate to takeoff speed 266 ft/s) 174 lb (78.9 kg)
3. Accelerate to Mach 0.86 at maximum military thrust 741 lb (336.1 kg)
4. Climb to 15,000 ft and accelerate to Mach 0.9 724 lb (328.4 kg)
5. Climb to 17,500 ft and accelerate to Mach 1.25 935 lb (424.1 kg)
6. Climb to 33,000 ft and accelerate to Mach 1.72 3,528 lb (1,600.2 kg)
7. Climb to 45,000 ft at Mach 1.72 805 lb (365.1 kg)
8. Climb to 52,000 ft and decelerate to Mach 1.6 214 lb (97.1 kg)
9. Cruise 1,000 nm at 52,000 ft and Mach 1.6 15,893 lb (7,208.9 kg)
10. Climb to 54,000 ft at Mach 1.6 262 lb (118.8 kg)
11. Dash 750 nm at 54,000 ft and Mach 1.6 10,589 lb (4,803.1 kg)
12. Descend to 50,000 ft at Mach 1.6 53 lb (24.0 kg)
13. Turn 180º at n = 1.55 658 lb (298.5 kg)
14. Drop 4 × 2,000 lb JDAM’s 0 lb (0 kg)
15. Climb to 56,000 ft at Mach 1.6 381 lb (172.8 kg)
16. Dash 750 nm at 55,000 ft and Mach 1.6 8,879 lb (4,027 kg)
17. Climb to 58,000 ft at Mach 1.6 139 lb (63.0 kg)
18. Cruise 1,000 nm at 58,000 ft and Mach 1.6 10,532 lb (4,777.2 kg)
19. Descend to Sea-Level and decelerate to loiter speed (391 737 lb (334.3 kg)
ft/s)
20. Loiter 30 min at Sea-level and 397 ft/s 2,453 lb (1112.6 kg)
21. Decelerate to landing speed 266 ft/s 132 lb (59.9 kg)
22. Land (decelerate to zero speed) 25 lb (11.3 kg)
23. Unload non-fixed equipment (2 × AMRAAM’s and crew
1,280lb)
0 lb (0 kg)
11.5 Takeoff & Landing
The RFP requires that the aircraft be able to takeoff and land on an icy standard NATO runway
8,000 ft (2,438 m) long. Takeoff and landing calculations were done according to MIL-C5011A.
Takeoff and landing were simulated by numerically integrating velocity and rate of climb to
determine distances and altitudes over a standard flight profile. Additional drag due to flaps and
landing gear was taken into account for takeoff and landing as well as -25% military thrust from
a thrust reverser during landing. The takeoff and landing profiles used in the simulation are
listed in Table 11.VII and Table 11.VIII.
Table 11.VII - Takeoff Flight Profile
1. Ground roll under max. power up to 1.2 V stall
2. Rotate for 3 s
3. 1.15g pull-up to transition to climb
4. Climb out at a 10º climb angle over 50 ft obstacle
79

Table 11.VIII - Landing Flight Profile
1. Approach from 50 ft obstacle on 3º glide slope
2. 1.15g pull-up for flare – touchdown at 1.2 V stall
3. Roll for 3 s
4. Brake and apply thrust reversers until stopped
MIL-C5011A defines field length to be the distance required to takeoff and clear a 50 ft (15.2 m)
obstacle or the distance to land from a 50 ft (15.2 m) obstacle. Takeoff and touchdown speed are
defined as 1.2 times the aircraft’s stall speed, and the speed over the 50 ft (15.2 m) obstacle must
be greater or equal to 1.3 times the stall speed for both takeoff and landing. The takeoff gross
weight for the design mission of 124,800 lb (56,610 kg) was used for the aircraft weight for both
takeoff and landing calculations. This allows the aircraft to land immediately after takeoff
without the need to jettison fuel or weapons. Takeoff and touchdown speeds were always greater
than the required 1.2 stall speed because of the acceleration during the 3 second rotation and roll
periods. During takeoff, due to the high speeds of the aircraft, the 50 ft (15.2 m) obstacle was
cleared before the climb angle was reached, so the climb segment of the profile was ignored.
Also, to simplify the calculations, the landing simulation was run backward so that the
touchdown point could be found without having to calculate the altitude and speed at the
beginning of the flare necessary to have the touchdown occur at the correct altitude and speed.
The results of the takeoff and landing simulations that are listed in Table 11.IX and Table 11.X
show that the RFP requirements for takeoff and landing on an icy 8,000 ft (2,438 m) runway are
met.
Table 11.IX - Takeoff Results
Weight
124,800 lb (56,610 kg)
Maximum lift coefficient 1.2
Stall speed 242 ft/s (73.7 m/s)
Takeoff speed 266 ft/s (81.1 m/s)
50 ft obstacle speed ≥ 290 ft/s (88.4 m/s) 326 ft/s (99.4 m/s)
Rolling friction coefficient 0.025
Runway length 3,842 ft (1,171 m)
Field length over 50 ft obstacle 5,280 ft (1,609 m)
80

Table 11.X - Landing Results
Weight
124,800 lb (56,610 kg)
Maximum lift coefficient 1.2
Stall speed 242 ft/s (73.7 m/s)
Takeoff speed 266 ft/s (81.1 m/s)
50 ft obstacle speed ≥ 290 ft/s (88.4 m/s) 294 ft/s (89.6 m/s)
Dry braking friction coefficient 0.3
Icy braking friction coefficient 0.1
Thrust reverser effectiveness
25% Mil.
Dry runway length 3,857 ft (1,176 m)
Dry field length over 50 ft obstacle 5,256 ft (1,602 m)
Icy runway length 5,707 ft (1,739 m)
Icy field length over 50 ft obstacle 7,107 ft (2,166 m)
11.6 Performance Summary
As shown above, the Vendetta meets all performance requirements in the RFP. A summary of
the performance requirements is shown in Table 11.XI.
Table 11.XI - Performance Summary
Requirement
Value
Supercruise Mach Number 1.6 1.6
Supercruise Mission Radius 1,750 nm (3,240 km) 1,750 nm (3,240 km)
Dash Altitude > 50,000 ft (15,240 km) > 54,000 ft (16,460 km)
1g Mil. Ps M = 1.6, 50,000 ft 0 ft/s (0 m/s) 95.7 ft/s (29.2 m/s)
1g Max. Ps M = 1.6, 50,000 ft 200 ft/s (61.0 m/s) 316 ft/s (96.3 m/s)
2g Max. Ps M = 1.6, 50,000 ft 0 ft/s (0 m/s) 108 ft/s (32.9 m/s)
Max. Turn Rate M = 0.9, 15,000 ft 8 deg/s 13.1 deg/s
Takeoff Field Length 8,000 ft (2,438 m) 5,280 ft (1,609 m)
Landing Field Length 8,000 ft (2,438 m) 7,107 ft (2,166 m)
Fuel Burn Over Mission ─ 57,940 lb (26,280 kg)
Maximum Mach Number at 36,000 ft ─ 2.1
81

11.7 Alternate Missions
In addition to the design mission, the Vendetta shows great performance on alternate missions.
The multi purpose rotary launcher (MPRL), shown in Figure 11.8, carried by the Vendetta allows
it to carry a total of 8 × 2,000 lb (907 kg) bombs compared to the 4 required for the design
mission (no AMRAAM’s can be carried in this configuration.) Compared to an earlier custom
designed rotary launcher, the MPRL has only a 4 in (10 cm) greater diameter. The extra cross
sectional area due to the MPRL only adds approximately 1,300 lb (590 kg) of fuel consumption
for the design mission. The performance of the Vendetta over three alternate missions was
calculated. Fully loaded missions and subsonic missions flown at Mach 0.8 and an altitude of
approximately 30,000 ft (9,144 m) were considered. The results shown in Table 11.XII indicate
that only a small loss of range occurs due to the additional weight of 8 × 2,000 lb (907 kg)
bombs, and the range of the aircraft can be greatly extended by flying subsonic (although it
extends the mission to 11 hours long.)
Figure 11.8 - MPRL with 8 × 2,000 lb (907 kg) JDAM’s
82

Table 11.XII - Alternate Mission Results
Design Mission
Mission Radius 1,750 nm (3,240 km)
Takeoff Weight
124,800 lb (56,610 kg)
Mission Time
4 hr. 57 min.
8 × 2,000 lb (907 kg) bombs – Supersonic
Mission Radius 1,690 nm (3,130 km)
Takeoff Weight
132,800 lb (60,240 kg)
Mission Time
4 hr. 51 min.
4 × 2,000 lb (907 kg) bombs – Subsonic
Mission Radius 2,380 nm (4,410 km)
Takeoff Weight
124,800 lb (56,610 kg)
Mission Time
11 hr. 16 min.
8 × 2,000 lb (907 kg) bombs – Subsonic
Mission Radius 2,300 nm (4,260 km)
Takeoff Weight
132,800 lb (60,240 kg)
Mission Time
10 hr. 56 min.
83

12 Payload
Weapon internal layout drove the Vendetta’s size and layout. For small C.G. excursion due to
weapons deployment all stores
were initially positioned as close to
the C.G as possible. As shown in
Figure 12.1, three configurations
were produced. Configuration one
utilizes a standard weapons bay
configuration. The large weapons
bay drove the configuration to over
120 ft (36.576m) in length when
room for landing gear and
weapons targeting systems were
integrated. In an effort to decrease
overall size a small rotary launcher
was designed and integrated into a
second configuration. This step
increased the maximum cross
sectional area by 5ft 2 (0.464m 2 Figure 12.1 - L to R configurations 1, 2, 3
)
and shortened the length of the
aircraft to 95ft (28.9m).
Figure 12.2 - 180 inch MPRL
The next iteration of the design utilized the
existing 180in (4.57m) MPRL out of the B-1B
and shown in Figure 12.2. This caused the final
configuration to grow to 105ft (32m) in length
and maximum cross sectional area of 55 ft 2
(5.11m 2 ). Utilizing the MPRL allowed for a wide
selection of alternate missions and configurations
to be discussed later.
In an effort to ascertain the feasibility of RFP delineated weapons as supersonic deployment
candidates research was done for each proposed system. As can be seen
in the foldout only one of all of the weapons has even been wind tunnel
tested for supersonic deployment. This led to the proposed systems for
possible supersonic deployment.
Retrofitting many of the weapon systems with a ballute and sabot,
shown in Figure 12.3, would aide in supersonic stability. The use of a
bomb bay supersonic flow deflector and acoustical resonance damping
system as well as flow modification system would aide in deployment.
Figure 12.3 - Ballute
and Sabot
84

Standard ten degree fall clearance is maintained for all weapons
stowed and weapons bay doors were designed to rotate into the
bomb bay, shown in Figure 12.4, and not into the free stream.
Rotating the bomb bay doors into the fuselage has no detrimental
effects on lateral stability, allows for the usage of lighter bomb bay
doors, and lowers the radar cross section of the aircraft when the
bomb bay is open versus a door which opens into the free stream
such as those found on the F-117.
Figure 12.5 - 30in (76.2cm)
Ejector rack
Alternative
uses for the Vendetta is the main
reasoning behind the choice of the
MPRL. According to MIL-A-8861B a
fighter/attack aircraft must be capable of
withstanding, under “basic flight design
gross weight” +7.5, -3 g’s. under
“maximum design gross weight” +5.5,
-2 g’s. According to the RFP we must
Figure 12.4 - Bomb Bay
Door Retraction Scheme
In an effort to minimize undesirable underbody flow the entire
underside of the aircraft was kept as flat as possible. The MPRL
chosen allows the use of 30in (76.2cm) ejector racks shown in Figure
12.5. The rack has electrically fired impulse cartridges, a gas operated
mechanism, and is designed to forcibly eject conventional or nuclear
stores up to 4000 lb (1810 kg) weight class. The LAU-142A ejector
was used with the AIM-120C shown in Figure 12.6.
be able, with the design load, and 50% of fuel volume structurally be able to withstand +7 -3 g’s.
Thus it is conceivable to design the aircraft to structurally withstand the RFP’s mandated +7 -3 g
loading with the RFP design load of (4) 2000lb (905 kg) JDAMs and (2) AIM-120’s.
Alternatively the aircraft could then be loaded with (8) 2000 lb (905 kg) JDAM’s shown in
Figure 12.7, and still meet the military specification of +5.5 -2 g’s at “maximum design gross
weight”. This fully satisfies the RFP and military specifications.
This doubling of weapons load shortens mission radius by
150 miles (241 km) (see performance section). Alternative
mission configurations might also include eight 2000lb (905
kg) fuel tanks for increased ferry range or 8 JASSM next
generation cruise missiles specifically designed for the B-1’s
MPRL.
Weapons guidance is accomplished with the on board
INS/GPS system (all weapons), a AN/APG-77 radar system
(AIM-120 guidance) as well as an on-board second
generation tessa LANTIRN system (laser designated
weapons). More detailed information on weapons can be found in Foldout 2 of the Appendix.
Figure 12.7 - (8) 2000lb JDAM + MPRL
Figure 12.6 - LAU-142A Ejection Sequence
85

13 Cockpit
Cockpit Design began with the RFP requirement of a crew of two. Comparison trades of tandem
versus side-by-side seating configurations were performed as shown in Figure 13.1. Analysis of
the two configurations was based upon a cockpit solid model where instrumentation, controls,
circuit breakers and military aft pilot vision requirements (MIL-STD-850B) were used to
construct the tandem and side by side arrangements.
Figure 13.1 - Cockpit Width Trade Study
Tandem vs. side by side configuration has very little effect on frontal area of the cockpit
configuration in military aircraft. It was found due to instrumentation and control placement,
which were the major cockpit width contributors, that other factors must be taken into account
before a final decision could be made on pilot placements. Weapons configuration, the use of the
180 in (457 cm) MPRL, favored side by side seating placement. The rational was the width of
the rotary launching system allowed for the use of a side-by-side seating arrangement. This
arrangement allowed greater pilot communication as well as
the elimination of many redundant circuit breakers as well as
instrumentation. However preliminary stability and control
analysis revealed a need to narrow the forward fuselage, as
shown in Figure 13.2, due to adverse C mα characteristics of a
wide nose section (see stability and control section).
Figure 13.2 - Forward fuselage
Comparisons
Therefore the decision was made to utilize a tandem seating
configuration. This configuration offered a smaller frontal
area, a much more ideal supersonic (M=1.6) area plot, and a
better field of vision for the primary pilot.
86

Table 13.I - Military Vision Specifications
Forward Pilot
5.1.1 Vision
azimuth
(°)
up
(°)
down
(°)
0 10 11
20 20
30 25
90 40
135 20
11°
5°
5.1.2 Aft Pilot Position
0 5
Figure 13.3 - Virtual Cockpit Model
Utilizing this information the virtual cockpit model shown in Figure 13.3 was generated. The
solid model also took into account the use of an ejection seat, room for instrumentation, controls,
switch placement as well as the above military vision specifications.
Further vision refinement produced rectilinear vision plots as shown in Figure 13.4.
Canopy reinforcing structure was removed from the areas between 25° and 40° up to aide in inflight
refueling vision. Runway vision areas are defined and every effort was made to increase
downward vision to aide in ground handling as well as takeoff and landing.
Takeoff and Landing vision is inherently limited in supersonic aircraft. In an effort to reduce
pilot workload and increase flight ability, multifunction displays (MFD) in modern aircraft can
double as vision aid tools. MFD 1 incorporated into the glare shield and upper instrument panel
Figure 13.4 – Rectilinear Vision Plot of Forward Cockpit Position
87

that could be used in landing and takeoff to increase downward
vision, meshing seamlessly with the actual cockpit over nose
view shown in Figure 13.5.
MFD’s 2, 3, and 4 display moving map imagery, flight critical
data, and mission critical information. The moving map
display could also double in landings as another artifical vision
aide. Perhaps utilizing infra-red or other electromagnetic
spectrums for poor weather penetration and increased all
weather capabilities. The standard dash mounted HUD was
dropped in favor of a current helmet mounted HUD systems
under development shown in Figure 13.6. The HUD allows far
superior situational awareness as well as more aerodynamic
canopy configuration.
Figure 13.5 - Cockpit Display
Arrangement
Figure 13.6 - Helmet
Mounted HUD
Aircrew safety was a primary concern in the design of Vendetta. Due to
RFP requirements the majority of the Vendetta’s mission will occur
above the military specified ceiling for flight without a full pressure suit
(50,000ft). Further research revealed the reasoning behind the
specification. The NASA Bioastronautics study SP-3006 shows that
animal and human life functions become critically affected by the lower
oxygen content and lower pressure of the upper atmosphere. The study
outlines how physiological effects such as the bends and hypoxia as
well as the extremely low temperatures of high altitude within seconds
render a human unconscious and dead in a mater of minutes. Also
outlined is the Armstrong Line (63,000 ft), or the altitude at which
water, at room temperature, will freely boil. In the study it shows how
animals survived momentary exposure to altitudes higher than 63,000ft
due to intravenous pressure keeping the blood within their veins liquid.
Balancing this information against the economics and long prep and turn around time associated
with full pressure suits the decision was made to opt for a partial pressure suit configuration. The
advanced fighter crew protection system is shown in Figure 13.8. A partial pressure suit system
was developed specifically for this altitude mission. It represents the next step beyond current
systems and offers low unit cost in comparison to full pressure suits as well as low turn around
time due to no necessity for a suiting procedure which involves lowering of blood nitrogen levels
such as those used in the U-2.
88

Ejection seat trade
studies will need to be
performed for the
final configuration.
The ACES II ejection
seat shown in Figure
13.7 was initially
chosen as the premier
American seat
available. However
due to global politics
the availability of
many superior
Russian seats would
produce a necessary
Figure 13.7 - ACES II
Ejection Seat
trade between increased cost of bi-lingual
design teams, language retrofit on new
equipment, product per system cost and the
unknown safety superiority of the Russian
seat(s).
Figure 13.8 - AFCPS
Further pilot safety issues were found in the
canopy design. The use of an ejectable canopy
was thrown out over the choice of a shape
charge cutting system due to the possibility of
aircrew and canopy striking each other after
ejection.
89

14 Systems
The systems of the Vendetta will closely follow the design architecture of the F-22. Technology
advances by 2020 will render most of the electronics aboard the F-22 antequated, thus the next
generation of this system should be implemented. Design trades on communications, processor
I/O (such as unified vs. federated), system redundancy, and actuation will have to be performed
on the new system.
14.1 Auxiliary Power Generation System
The auxiliary power generation
system consists of two components:
an auxiliary power unit (APU) aand
a self-contained energy storage
system (SES). APU selection
involved examining a number of
mid-sized, gas turbine, generators
with output exceeding the minimum
350 Hp (261.1 kW) estimated as
needed for the Vendetta. A
shortened list appears in Table 14.I.
The RFP lists an APU but research
Company
Table 14.I - APU Selection Table
points to the cost, volume and weight of the APU specified as being a Ram Air Turbine. A Ram
Air Turbine (RAT) was eliminated due to the need for ground power and previous service
experience. Modern designs utilize a SES for power backup due to its higher reliability,
invulnerability to aircraft maneuver position, and airspeed.
Model
Honeywell 131-9A
Honeywell 36-300
Honeywell 331-200
Pratt and
Whitney PW901
Sundstrand
Sundstrand
APS2100
APS3200
Startup
Ceiling
Dry
Weight Rating Power/Weight
ft (m) lb (kg) kW Hp Hp/lb (kW/kg)
41000 350
(12497) (158) 343 460 1.31 (2.17)
35000 300
(10668) (136) 291 390 1.3 (2.14)
43000 500
(13106) (226) 432 579 1.16 (1.91)
25000 835
(7620) (378) 1145 1535 1.84 (3.03)
37000 270
(11278) (122) 376 504 1.87 (3.08)
39000 305
(11887) (138) 450 603 1.98 (3.26)
Due to common unreliability of in-flight APU startup, startup ceiling
was not seen as a major driver in APU selection. Overall high power
to weight ratio as well as a rating above 350 Hp (261.1 kW) and small
size was seen as the main APU selection criteria. With this in mind the
Sundstrand APS 3200 was selected. Placement of the APU can be
seen in Figure 14.1.
Figure 14.1 - Sundstrand
APS 3200 Location
The SES is a design point to be focused on in the next level of design.
Current hypergolically fuelled systems offer high power to weigh ratio
but fuels are hazardous and expensive. Next generation fuel cell
systems offer reasonably high power to weight ratios with standard
JP-8. Trade studies considering cost of development, service life cost
savings, and internal volume usage will have to be performed.
90

14.2 Vehicle Management System
The vehicle management system (VMS) of the aircraft provides flight and propulsion control and
includes the following systems: Control stick, Throttle controls, Rudder pedals, and actuators,
Air data probes, Accelerometers Leading-edge, flap drive actuators, Primary flight control
actuators .
The choice of primary actuator type was the first trade examined. Four choices are commonly
available for these systems:
1) Electrohydrostatic
2) Electric
3) Pneumatic
4) Hydraulic
Pneumatic systems were eliminated due to low power to weight, large comparative size, and low
power transmission efficiencies. Electrohydrostatic actuators, although offering many benefits
such as a “line replaceable unit”, optimized per service dynamic impedance shaping, and
optimized K factor suffered from low observability and weight problems as a consequence of
electric power transmission throughout the aircraft( EM emissions in the IR (actuator heating)
and Radio (cables) spectra). This low observability problem to an even greater degree affected
the all electric system which utilizes a comparatively large and heavy actuator.
Thus the decision was made to utilize an all hydraulic system with digital fly-by-wire control.
The F-119 utilized in the F-22 has a PTO driving two 72 gpm (273 lpm) pumps (main line), four
pumps total to supply two independent 4000 psi (27.6 GPa) systems. The choice of the high
pressure systems was due to weight and volume considerations. Peak hydraulic system demand
will be satisfied via an air pressurized hydraulic reservoir allowing for a constant energy bleed
from the engines.
Further detail was then examined into the electric system of the aircraft. Borrowing the electrical
system from the F-22, a Smiths Industries 270 volt, direct current (DC) electrical system was
chosen. It uses two PTO driven 65 kilowatt generators.
Next level design trade studies need to be completed on the redundancy level of the electrical
and flight control systems. Flight control system design should be aimed at Operation-
Operational-Operational Fail (OOOF) or better, safety design. INS/GPS and other navigational
systems should be examined and a proper redundancy level chosen.
91

14.3 Fuel System
Initial fuel system design began with
configuration placement of fuel tanks
symmetrically about the CG laterally and
longitudinally. The two fuselage tanks (Figure
14.2) serve to trim the aircraft in flight,
necessary for the shift in neutral point location
due to supersonic flight.
Aircraft sizing began with fuel load
internal volume necessary to complete the
mission. The final configuration provides for
Figure 14.2 - Fuel Tank Locations in Vendetta 68,000 lb (30,770 kg) of fuel to be carried
within 80 %(fuselage) and 75 %(wings) volume
usage tanks. All tanks in the aircraft are
pressurized with nitrogen gas from the on-board inert gas generating system (OBIGGS).
Pressurizing is minimal due to structural constraints and JP-8’s low vapor pressure of 0.029 psia
@ 100 °F (200 Pa @ 42.3 °C). Nitrogen reduces fuel fumes and thus the chance of an accidental
explosion. All tanks on the aircraft are self sealing and feature flame resistant overflow and
exhaust venting.
Single point fueling and de-fueling can be
performed from the starboard side of the
forward fuselage Figure 14.3. This fueling
point shares common lines with the Air Force
retractable fueling boom port (Figure 14.4)
located on the upper portion of the same
segment of forward fuselage. Both ports offer
fueling rates as high as 1100 gpm (4164 lpm),
the maximum KC-10 Fuel Probe refueling rate
(Table 10.II).
Fuel is then power transferred to tank three
and four. From tank four it is distributed to
wing tanks one and two.
Figure 14.3 - Fuel System Architecture
Figure 14.4 - Retractable in-flight
refueling boom ports, F22, F-117, B-2
92

Gravity ground refueling is provided via ports on the upper surface of the wing into tanks one
and two. From there fuel can gravity feed into tanks three and four. This same gravity feed
system can be utilized in flight given power feed system loss.
All major thermal transfer within the aircraft is performed by the fuel system. The air cooled fuel
cooler utilizing inlet duct boundary layer and flow control diversion air, placed ahead of the main
gear bays dissipates all kinetic heating experienced by the airframe as well as systems heat
generated. Dual heat exchangers are utilized for combat survivability. Fuel flow rate system
requirements (Table 14.II) will be used to size fuel lines and pumps. All fuel lines are redundant
to provide for fuel circulation and system combat survivability. The Vendetta is fitted with an
onboard fire suppression system, utilizing a halon suppressant in the engine bay and fuel cutoff
valves at all fuel tanks. The system is designed so it can be retrofitted with a more
environmentally friendly suppression chemical in the future.
Table 14.II - Fuel System Sizing Requirements
Fuel System Sizing Requirement
Fuel Flow Rate
GPM (LPM)
CG Shift Requirement Between #3 and #4 Fuel Tanks 100 (379)
KC-10 Probe Maximum Refueling Rate 1100 (4164)
(2) F-119 Turbofan @ Max Thrust With Reheat 294 (1113)
Air Cooled Fuel Cooler (ACFC) 400 (1514)
93

15 Manufacturing
Several manufacturing considerations have been considered and planned for in the design of the
Vendetta. One of these considerations is part commonality, which reduces the total number of
separate parts that must be manufactured; thereby lowering manufacturing costs. The Vendetta
is symmetrical in that both left and right wings,
landing gear, horizontal and vertical stabilizers,
etc. will be manufactured almost exactly the
same. Furthermore, all flying surfaces are
symmetrical and would provide an additional
improvement in manufacturing costs.
The design has also taken into consideration the
routing of electrical lines, hydraulic lines, etc.
These systems would be interconnected through a
routing tunnel through the fuselage of the aircraft.
This would reduce installation complexity and
therefore reduce the amount of labor involved in
the installation process. The routing tunnel as it
passes beside the forward fuel tank, and aft past
the weapons bay is shown below in Figure 15.1.
Figure 15.1 - Routing Tunnel
Manufacturing breaks include the wings, empennage, forward, center, and aft portions of the
fuselage, as well as the landing gear itself. These breaks are shown in Figure 15.2.
Figure 15.2 - Manufacturing Breaks
94

The entire propulsion system will be capable of being dropped out the bottom of the aircraft
which will provide for an easy installation during the manufacturing process as well as allowing
for easy access during routine maintenance. Computer-aided manufacturing will enable more
complex parts to be machined by computer-numerically-controlled (CNC) machining tools.
Large items such as bulkheads can be easily machined from a single piece of metal. This would
be required in order to meet the stringent structural load limits mandated by the AIAA RFP.
Inspection and maintenance panels will be placed wherever possible throughout the aircraft
without compromising the low-observability requirements. Furthermore, access panels will be
built as “structural doors” able to carry through the skin loads that will also be required to meet
the stringent structural load limits. These access panels will ease maintenance and reduce
maintenance hours required per flight hour.
The assembly line would allow for major components, such as the wing, fuselage, and
empennage to be pre-fabricated at possibly other site locations and brought in to a central
assembly line as shown below in Figure 15.3.
Figure 15.3 - Assembly Line
95

16 Cost Analysis
The final and perhaps most important issue in the purposed development of the Vendetta is the
cost analysis. The methodologies used in developing this analysis were found in the Raymer and
Nicolai texts. Despite the fact that the Nicolai text was written in 1974, when adjusted for
inflation, the method was accurate to within 5% of that method found in the 1999 Raymer text.
Both of these analyses are adjusted for inflation to 2000 dollars. The methods used in the cost
analysis were based on the DAPCA IV model developed by the RAND Corporation. This model
provided a means of calculating the operating cost, life cycle cost, flyaway cost, and the cost
required for research, development, test, and evaluation, or RDT&E.
The RDT&E cost was predicted to be approximately $6.5 billion; whereas, the flyaway cost for a
200 unit buy was calculated to be $128.5 million. This cost approximately 15% under that cost
required by the AIAA RFP set at $150 million dollars per 200 unit buy. The cost per aircraft
based on the number of aircraft purchased is shown below in Figure 16.1.
Figure 16.1 - Cost Analysis
The figure indicates that the cost per aircraft at a 600 unit buy is significantly less at $80.5
million. Note the cost of engineering, development, manufacturing, and materials in the cost
breakdown per unit at a 600 unit buy in comparison to the cost breakdown per unit at a 200 unit
buy; the percentages associated with development and engineering decreases while the
manufacturing and materials percentages increase. This is due to the fact that at a 600 unit buy,
96

there are more aircraft available to help pay the $6.5 billion cost associated with RDT&E.
Furthermore, there is a learning curve associated with the development of a large quantity of
aircraft and the airplane become even less costly to produce.
Four factors were considered when determining the operating cost of the Vendetta. These factors
included the cost of the fuel and pilots, as well as the cost of parts and maintenance personal.
Raymer estimates that a bomber flies approximately 400 hours per year and requires 40
maintenance man hours per flight hour. In addition, because the Vendetta is designed to fly very
fast at high altitudes, the fuel cost is a large percentage of the total operating cost. The operating
cost of the Vendetta is calculated to be $13,000 per flight hour. This cost breakdown is shown
below in Figure 16.2.
Figure 16.2 - Operating Cost
One final cost that must be considered beyond the cost of RDT&E, flyaway, and operations is
the lifecycle cost. This cost considers the cost of RDT&E, flyaway, and operations over a 30
year period at 400 flight hours per year, as well as the cost of disposal. This cost is totaled at
$293 million per aircraft at a 200 unit buy. A breakdown of the lifecycle cost is shown below in
Figure 16.3.
Figure 16.3 - Lifecycle Cost
97

Appendix
Threats Chart
Common Surface to Air Missiles and Threats to Airborne Aircraft
98

Diffuser Efficiency
Literal Factor Forms
Mode Damping ratio (ζ) Natural Frequency (ω n )
Phugoid
ζ
P
−X
u
=
2ω
Zα
Mq
+ M α
+
Short Period u
0
ζ =
SP
2ω
Dutch Roll
ζ
DR
1
β 0 r
=− ⎜ ⎟
2ωn
u0
DR
n
n
⎛Y
⎝
p
SP
+ u N
⎞
⎠
ω
n
DR
ω
n
P
=
Z M
−Zg
u
u
α q
ω
n
= −
SP
u0
=
0
M
α
YN − NY+
uN
β β β r 0 β
u
0
99

Wing Break Point
Fuel
Aft Fuselage
Break Point
56°
14°
13°
Forward Fuselage
Break Point
54.7' [16669.5]
20.7'
[6294.8]
5.3'
[1621.3]
Fuel
Fuel
102.8' [31337.8]
5°
11°

(4) Mk-84 LDGP + (2) AIM-120
Weapon Weight
1,967 lb (890 kg)
Configuration Weight 10,222 lb (4,625 kg)
Weapon Length
12.6 ft (3.84m)
Weapon Diameter
18 in (45.7cm)
Tail Span 2 ft (0.61 m)
Max Drop Height
Unlimited
Max Tested Drop Velocity M=1.3
Guidance
Ballistic
Weapon Information:
Development of the Mk 84 Low Drag General
Purpose Bomb for use by the United States armed
forces began in the 1950’s. The Mk 84 bomb, which is
fitted with 30 in (0.762m) spaced suspension lugs, is
packed with 942 lb (426 kg) of Tritonal or H-6. The
known inventory of Mk 81, 82, and 84 bombs is 1.13
million.
(4) GBU-27 + (2) AIM-120
Weapon Weight
2,165 lb (980 kg)
Configuration Weight 11,014 lb (4,984 kg)
Weapon Length 13.9 ft (4.24 m)
Weapon Diameter
14.6 in (37 cm)
Tail Span 2 ft (0.61 m)
Max Drop Height
Unlimited
Max Tested Drop Velocity Unknown
Guidance
Semi-Active Laser
Weapon Information:
The GBU-27 is a modified GBU-24 Paveway III
designed for internal carriage in the F-117A. This LGB
carries the designation GBU-27 /B and uses a BLU-
109 /B penetrator bomb for its warhead. The main
modifications made to the GBU-24 were to have
shorter adaptor rings and to use the GBU-10’s rear
wing unit to decrease the bomb’s length, and to clip the
canards in order to make the weapon fit into the small
F-117A Bomb Bay. The other major difference was the
use of radar absorbing materials in order to prevent the
bombs from being picked up by enemy radar once the
aircraft’s bomb doors were opened. As a result of these
modifications, the GBU-27 has a shorter range than the
GBU-24, which can also be launched at lower
altitudes.
Guidance is by semi-active laser, the scanning
detector assembly and laser energy receiver being
mounted in the front of the canister behind the glass
dome. After the bomb is released the laser error
detector measures the angle between the bomb’s
velocity vector and the line between the bomb and
target. Steering corrections are made by moving the
nose mounted canard control fins to adjust the bomb’s
trajectory to line up with the target. The tail fins/wings
are for stabilization purposes only. Target illumination
for the system may be either by an aircraft-mounted
laser marker (not necessarily the parent aircraft) or a
ground-based laser transmitter.
(4) 2000lb JDAM +(2) AIM-120
Weapon Weight 2,100 lb (950 kg)
Configuration Weight 10,754 lb (4,866 kg)
Weapon Length 13.2 ft (4.02 m)
Weapon Diameter 18 in (46 cm)
Tail Span 2 ft (0.61 m)
Max Drop Height Unlimited
Max Drop Velocity M=1.3 tested
Guidance
GPS / INS
Weapon Information:
A parallel program to the AGM-154 JSOW the
GBU-31 JDAM program began in the late 1980’s. The
goal of the program was to produce a low cost guided
munition. Interesting to note is the GBU-31 is soon to
be replaced by the GBU-32/35. This new weapon, will
utilize an I-1000 1000 lb (452.5 kg) penetrator warhead
and is intended for future use in the F-22 raptor. This
weapon, the GBU-32/35 is being used to size the
raptor’s weapon bays.
The GBU-31 utilizes both the Mk 84 and BLU-109
warheads. Due to the Mk 84’s low cost, and
commonality, it was chosen for the solid model seen
above. The GBU-31 consists of three major
subassemblies. The warhead (Mk 84), Saddleback stub
wing assembly (attaches at hardpoints, three
components), and a bolt on tail cone guidance kit.
The guidance kit, contained within the replacement
bolt-on tail cone consists of the following key
elements: combined inertial measuring unit and GPS
receiver; flight control computer; battery and power
distribution unit; tail actuators and four movable
clipped delta fins in a cruciform configuration. In
keeping with other GPS guided weapons, the unit is
believed to be fitted with two GPS antennas, one on
top of the unit for initial flight and one in the tail for
good reception during terminal maneuvering.
Prior to bomb release the guidance unit will be fed
with aircraft position, velocity and target coordinates
through the aircraft to bomb interface. After release the
bomb will guide itself to the target by means of rear fin
deflection, which are driven by commands from an
onboard computer that is constantly being updated by
the GPS. The combination of the INS/GPS is expected
to allow the bombs to hit within 32.8 ft (10 m) to 49.2
ft ( 15 m) of their targets. Wind tunnel tests in 1996 are
reported to have cleared JDAM for release at up to
Mach 1.3.
(4) AGM-154 JSOW + (2) AIM-120
Weapon Weight 1,064 lb (481 kg)
Configuration Weight 6,610 lb (2,991 kg)
Weapon Length 14 ft (4.26 m)
Weapon Diameter 21 in (53 cm)
Tail Span 24 in (0.61 m)
Max Drop Height Unlimited
Max Drop Velocity Subsonic
Guidance
GPS / INS
Weapon Information:
In the late 1980’s the US Navy began a review of
conventional weapons with the intention of reducing
the number of weapon types. New systems were
selected for future development: JDAM, TSSAM,
JASSM, and the advanced interdiction weapon system
to be later named Joint Standoff Weapon (JSOW).
The JSOW program is intended to replace six
existing weapons: the AGM-65 Maverick, AGM-123
Skipper, AGM-62A Walleye, Rockeye and APAM
(Anti-Personnel/Anti-Material) submunition
dispensers, and laser- and TV- guided bombs.
Of particular attention on the previous list is:
1) All weapons are air to ground.
2) This weapon is designed to replace the
GBU-27, one of the weapons on the RFP
attachment 3 list.
The JSOW is an aerodynamically shaped,
unpowered glide dispenser with a rectangular crosssection
body shape. It is made up of three major
sections: a streamlined nose fairing that houses the
guidance and control system, a rectangular center
section payload container for holding the bomblets
(this is fitted with two folding high aspect ratio wings
on its upper surface, and two standard 30 in (0.762 m)
spaced suspension lugs); and the tail section which has
six fixed, sweptback rectangular fins positioned
radially on the boat tail and contains the flight control
system.
(16) 250 lb Small Smart Bomb
Weapon Weight 250 lb (113 kg)
Configuration Weight 5,500 lb (2,489 kg)
Weapon Length 8.2 ft (2.5 m)
Weapon Diameter 6 in (0.15 m)
Max Drop Height Unlimited
Max Drop Velocity Unknown
Guidance
GPS / INS
Weapon Information:
The Small Smart Bomb is a 250 lb (113 kg) weapon
that has the same penetration capabilities as a 2000lb
(905 kg) BLU-109, but with only 50 lbs (22.6 kg) of
explosive. With the INS/GPS guidance in conjunction
with differential GPS (using all 12 channel receivers,
instead of only 5) corrections provided by GPS SPO
Accuracy Improvement Initiative (AII) and improved
Target Location Error (TLE), it can achieve a 5-8m
(16.4 to 26.3 ft) CEP. The submunition, with a smart
fuze, has been extensively tested against multi-layered
targets by Wright Laboratory under the Hard Target
Ordnance Program and Miniature Munitions
Technology Program. The length to diameter ratio and
nose shape are designed to optimize penetration for a
50lb (22.6 kg) charge. This weapon is also a potential
payload for standoff carrier vehicles such as
Tomahawk, JSOW, JASSM, Conventional ICBM, etc.
The Swing Wing Adapter Kit (SWAK) is added to give
the SSB standoff of greater than 25 nm (48.6 km) from
high altitude release. The wing kit is jettisoned at a
midcourse way point if penetration is required so that
velocity can be increased after wing release. For soft
targets the wing kit continues to extend the glide range
until small arms threat altitude is reached. At this point
the wings are released. With INS/GPS guidance,
coupled with AII, a 6-8 m (19.7 to 26.3 ft) CEP can be
achieved. This wing kit allows the SSB to be directly
attached to the aircraft at any 300 lb (135.75 kg) store
station. The major advantage to the 250 lb (113.125
kg) small smart bomb is an improved number of targets
per pass capability.
AIM-120 C AMRAAM
Weapon Weight 327 lb (148 kg)
Configuration Weight 5,500 lb (2,489 kg)
Weapon Length 12 ft (3.657 m)
Weapon Diameter 7 in (0.1778 m)
Fin Span 1 ft 6 in (0.457 m)
Max Drop Height Unlimited
Max Drop Velocity Supersonic
Guidance
Command from
Launch Aircraft
INS
Monopulse Radar
Seeker
Weapon Information:
The Advanced Medium-Range Air to Air Missile
(AMRAAM) AIM-120 development program was
started in 1975. It was designed to follow on and better
the performance of the Aim-7 Sparrow and be carried
on the F-14, F-15, F-16 and F/A-18 aircraft. In the late
90’s a modified(smaller) version of the missile, the
AIM-120C was developed to be fitted to the F-22
Raptor. This newer version also incorporates a dual
mode active and passive radar seeker. The AIM-120C
is deigned to be rail, ejector or trapeze launched. On
the F-22 the AIM-120C is launched using an EDO
corp. LAU-142/A hydraulic / pneumatic ejector.
In a typical engagement the missile is launched and
first guided by on-missile inertial navigation, with
command guidance updates from the launch aircraft.
The missile then goes into the mid-course autonomous
mode and continues to guide by inertial navigation
only. Finally, the terminal mode is automatically
initiated by the missile itself when the target is within
rage of the missile’s active monopulse radar seeker,
which then guides the missile onto the target aircraft.
Foldout 2 – Weapons Information

The Vendetta Design Team
Chris Atkinson is a bachelor’s candidate in aerospace engineering
with a computer science minor. His responsibilities for the Vendetta
included aerodynamic and performance analysis as well as validation
modeling. He participates in the Cal Poly Flight Simulation Group and
enjoys fencing and hiking.
Chris Droney is a blended program master’s candidate in aerospace
engineering. Chris was the team leader and lead configurator for the
Vendetta. He designs, builds, and flies R/C planes and gliders. Chris also
has his private pilot’s license and enjoys flying when he gets the chance.
Kolby Keiser is a 23 year old master’s candidate in aerospace
engineering. Her responsibility in the Vendetta design group was aircraft
propulsion. In her spare time, Kolby enjoys hiking, exercising, and other
outdoor activities, in addition to swing dancing. She also participates in
diabetes education in the community.
105

Chris Maglio is a bachelor’s candidate in aerospace engineering. His
responsibilities in the Vendetta design group encompassed configuration,
weapons, and systems. Chris is an active member of Central Coast Polo
and the USPA. Besides pursuing the hobby of mountain biking
throughout the academic year, Chris, working as a consultant, has
designed and manufactured numerous nautical, aerospace, civil, and
automotive products.
Dan Salluce is a 23 year old blended program master’s candidate in
aerospace engineering. His responsibilities in the Vendetta design
included stability and control analysis, radar cross section determination,
and flight simulator validation. Dan is interested in control system
design. He participates in the Cal Poly Flight Simulation Group, teaches
Aerospace classes at Cal Poly, and enjoys R/C gliders and the San Luis
Obispo nightlife and nearby beaches.
Nathan Schnaible is a 5 th year aerospace engineering student who
will graduate with a bachelor’s degree in June of 2002. Nathan worked
on the weights & balances, manufacturing, and cost analysis areas of the
Vendetta. His background in the aerospace industry includes
maintenance duties on refurbished Albatross’, ground service on both
corporate and commercial aircraft, as well as piloting experience.
Nathan will be attending the United States Navy Flight School in the
spring of 2003.
106

References
1. “Ejection Systems,” www.bfg-aerospace.com, BF Goodrich Corporation, 2001.
2. “Military Standard – Aircrew Station Controls and Dispalys: Location, Arrangement, and
Actuaion of, for Fixed Wing Aircraft,” United States Department of Defense, 1991.
3. “Military Standard – Aircrew Station Geometry for Military Aircraft,” United States
Department of Defense, 1976.
4. Abbott, I. H., Von Doenhoff, A. E. Theory of Wing Sections, Dover Publications, INC.
New York 1959.
5. Cummings D. Boeing Long Beach, Advanced Aircraft Design
6. Currey, N. S. Aircraft Landing Gear Design: Principles and Practices, AIAA,
Washington DC, 1988.
7. Dillenius, M. F. E. Perkins, S. C., Nixon, D., “Pylon Carriage and Separation of Stores,”
Tactical Missile Aerodynamics: General Topics, Edited by Michael J. Hemsch, Vol. 141,
Progress in Aeronautics and Astronautics, AIAA, New York, 1992, pp. 575-666.
8. Goodall, J. C. Americas Stealth Fighters and Bombers, Motorbooks International
Publishers and Wholesalers, Osceola, WI, 1992
9. Jane’s All The World’s Aircraft 2000-2001, Janes Information Group Inc. Alexandria,
Virginia, 2000
10. Jane’s Avionics 2001-2002. Janes Information Group Inc. Alexandria, Virginia, 2001
11. Lennox D. Jane’s Air-Launched Weapons Issue 35. Janes Information Group Inc.
Alexandria, Virginia, 2000
12. Mattingly, Jack D., Elements of Gas Turbine Propulsion, McGraw-Hill, Inc. New York,
NY, 1996.
13. MIL-A-8860B
14. MIL-A-8861B
15. MIL-STD-850B
16. MIL-E-5008B
107

17. NACA-TN-3182, “Manual of the ICAO Standard Atmosphere Calculations by the
NACA”, NASA, 1976
18. Nicolai, L. M. Fundamentals of Aircraft Design, METS Inc., California, 1984.
19. Oates, G. C. Aircraft Propulsion Systems Technology and Design, AIAA, Washington
DC, 1989.
20. Raymer, D. P. Aircraft Design: A Conceptual Approach – Third Edition, AIAA,
Washington DC, 1999.
21. Roskam, J. Airplane Design, Part I: Preliminary Sizing of Airplanes, DARcorporation,
Kansas, 1997.
22. Roskam, J. Airplane Design, Part II: Preliminary Preliminary Configuration Design and
Integration of the Propulsion System, DARcorporation, Kansas, 1997.
23. Roskam, J. Airplane Design: Part III, Roskam Aviation And Engineering Corporation,
Ottawa, KS, 1989, pp 1-34.
24. Roskam, J. Airplane Flight Dynamics and Automatic Flight Controls, DARcorporation,
Kansas, 1979.
25. Wilcox, F. J., Baysal, O., Stallings, R. L., “Tangential, Semisubmerged, and Internal
Store Carriage and Separation,” Tactical Missile Aerodynamics: General Topics, Edited
by Michael J. Hemsch, Vol. 141, Progress in Aeronautics and Astronautics, AIAA, New
York, 1992, pp. 667-721.
26. www.aeronautics.ru/nws002/f22/diagram05.jpg
27. www.aeronautics.ru/nws002/f22/diagram06.jpg
28. www.aeronautics.ru/nws002/f22/systems.htm
29. www.af.mil/news/efreedom/bombs.html
30. www.af.mil/news/factsheets/KC_10A_Extender.html
31. www.af.mil/news/factsheets/KC_135_stratotanker.html
32. www.arfl.afr.mil
33. www.aoe.vt.edu/aoe/faculty/Mason_f/M96SC.html
34. www.batnet.com/mfwright/spacesuit.html
108

35. www.dfrc.nasa.gov/PAO/PAIS/HTML/FS-061-DFRC.html
36. www.eureka.findlay.co.uk/archive_features/Arch_Automotive/n-push/n-push.html
37. www.fas.org/man/dod-101/sys/ac/equip/lau-142.htm
38. www.fas.org/man/dod-101/sys/ac/equip/lau-142.htm
39. www.fas.org/man/dod-101/sys/missle/amraam-5.jpg
40. www.fas.org/man/dod-101/sys/smart/agm-154.htm
41. www.fas.org/man/dod-101/usaf/docs/mast/annex_f/part06.htm
42. www.globalsecurity.org/military/systems/aircraft/f-22-fcas.htm
43. www.sff.net/people/geoffrey.landis/vacuum.html
44. www.skf-linear.co.il
109