Vendetta Final Proposal Part 2 - Cal Poly

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 a landing gear similar but smaller to an XB-70. The
complexity was necessary due to overall configuration drive of low supersonic maximum cross sectional area.
Figure 8.2 - Main Gear Retraction Sequence
The forward fuselage has ample volume to accommodate the nose gear thus no complex folding arrangements
were utilized. This facilitated the use a standard side-by-side tire configuration. The complete retraction schemes and
nose wheel configuration can be seen in Figure 8.3.
Figure 8.3 - Nose Gear and Main Gear Retraction Schemes
The braking system for both the nose gear and main gear configuration will use a standard rotor disk braking
mechanism. The rotors as well a pad material will be made of carbon rather than steel. Carbon offers superior thermal
conductivity, upper temperature limit, and lower thermal expansion. 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. 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. Vendetta’s landing gear allows for drive up loading utilizing either a
MJ-1 or MHU-83 lift truck. Landing gear sizing took account maximum lift truck reach to place weapons on the
Multipurpose Rotary Launcher (MPRL) within the Vendetta’s main weapons bay.
41

9 Weight & Balance
Weight and balance has proven to be a challenge in designing the Vendetta for both subsonic and supersonic
flight conditions. After having sized the aircraft using a weight fraction method, and after having developed an initial
configuration, the next step was 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
Weight (lb)
Accuracy
Component Roskam Nicolai Raymer Average Roskam Nicolai Raymer
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 9,280 9,280 9,280 9,280 0% 0% 0%
Fuel 58,974 58,974 58,974 58,974 0% 0% 0%
TOGW 123,433 128,215 128,215 127,778 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. A more accurate and detailed component weight buildup
was developed by considering all three methods and taking the average shared between them. One author’s estimation
was discarded if it did not agree to within ±30% of the average of the other authors’ estimations. The remaining weights
were averaged in order to develop a weight buildup for the entire aircraft. The class II weight buildup for the Vendetta
after the elimination process is shown in Table 9.II
42

Table 9.II - Final Component Weight Buildup
Weight (lb)
Component Roskam Nicolai Raymer Average
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 11,098 11,352 11,662 11,675
Systems 18,649 14,506 14,350 20,574
Payload 9,280 9,280 9,280 9,280
Fuel 58,974 58,974 58,974 58,974
TOGW 125,051
Inertias were calculated using guidelines outlined by the Society of Allied Weight Engineers (SAWE). Each
component 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 to calculate 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. The inertias are shown in
Table 9.IV and the non-dimensional radii of gyration coefficients as
Figure 9.1- Principle Axes
compared to the SAWE predicted coefficients are shown in Table 9.III.
Table 9.III 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 nonfixed
equipment, fuel, and payload.
43

Vendetta’s configuration was
created by placing components about
a predetermined CG location. The
components were arranged to create
the smallest airframe possible while
leaving room for the fuel and
payload. The heaviest fixed
equipment items were placed first
Table 9.III - Inertia Estimation
Inertias (slug ft 2 )
Ix Iy Iz Ixy Ixz Iyz
69,547 1,165,870 1,228,330 0 -6,478 0
Table 9.IV – SAWE Inertia Validation
Non-dimensional Radii of Gyration
Rx Ry Rz
Vendetta 0.16 0.35 0.46
SAWE 0.31 0.33 0.47
Accuracy 49% 5% 1%
followed by the smaller and lighter systems. Once the aircraft balanced empty, the payload was placed. This caused a
slight rearrangement of items until the aircraft balanced both empty and with the payload. This process was repeated for
the fuel loadout.
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 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 require additional fuel volume in an already congested aircraft. To minimize trim drag without the use
of a trim tank, the aircraft would have to fly with an unstable static margin, subsonically, and with a stable static margin,
supersonically until enough fuel could be consumed to trim the aircraft.
A center-of-gravity monitor makes use of fuel burn control in order to keep the aircraft as close as possible to a
neutrally stable flight condition. Furthermore at both TOGW and empty weight the aircraft is balanced such that it
provides for a 5% unstable static margin. With an aerodynamic shift of 12%, the aircraft transition to a 7% stable static
margin as it accelerates to supersonic flight. The center-of-gravity monitor then controls the fuel burn in such that the
Vendettas center-of-gravity follows the aerodynamic center and thus maintains neutral stability.
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 and payload throughout the mission. The code required four
inputs including; the locations and weights of Vendetta’s 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
44

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 throughout the 5-hour
mission. Using this data, the center-of-gravity path can then be plotted against 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
45

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-ofgravity
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.
46

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.
Table 10.I - Historical Aircraft Tail Volume Coefficients
Tail Volume Coefficients
Aircraft
V H
Boeing SST (2707-300) 0.36 0.049
Concorde n/a 0.080
GD F-111A 1.28 0.064
Rockwell B-1B 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
aircrafts’ 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 used to see how the increased weight of a
bigger horizontal affects the longitudinal static margin. It became apparent 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. At approximately 108 ft 2 of horizontal area the Vendetta has a neutrally stable static margin at
V V
47

Mach 0.3. A horizontal that is bigger than 108 ft 2 yields a stable aircraft but will pay the price in trim drag if the aircraft
is too stable. This size will certainly increase due to other constraints.
A stable static margin is necessary in flight without the use of a digital flight control system. The RFP mandates
an unaugmented static margin between -30% and 10% 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 drag that a neutrally stable or marginally stable (1-3%) aircraft is desired in cruise where the
aircraft does not need to maneuver much.
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 difference in neutral point and center-of-gravity greater.
The 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 more
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.
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 waste of space
and would complicate ground procedures where refueling would have to leave the tank partially empty. A canard could
be used to destabilize the aircraft by moving the neutral point forward and closer to the CG but it would make the
Vendetta less controllable 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 trim the aircraft (Section 7), with the same
being true when decelerating. Use of a digital flight control system (DFCS) which is provided as Government Furnished
Equipment (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
48

management system could then be used to enhance cruise performance by pumping fuel in a way which results in neutral
or marginal static stability. 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 10.1.
Figure 10.1 - Horizontal Area Required for Static Stability with Cant Angle
It can be seen from the plot that as cant angle increases, 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 eases the job of control system design. 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. 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 10.2.
49

Figure 10.2 - 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
.
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 10.3 for 5 GHz
monostatic radar sweeping a full 360° azimuth.
m r
50

40
30
20
10
0
-10
-20
-30
-40
-50
20° Canted Vertical
30° Canted Vertical
Figure 10.3 - Radar Cross Section Impact of 20° vs. 30° Vertical Cant Angle
Figure 10.3 clearly shows that there is an impact on the RCS for changing the cant angle. The RFP required -13
dBm 2 return is shown in red for those azimuth angles it is fulfilled. 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,
for this coupling term and various cants.
Cm
δ r
is nonexistent. Table 10.II shows the values
Table 10.II - Pitching Moment Coupling with
Rudder Deflection for Various Vertical Cant Angles
Vertical Cant Angle C
(165 ft 2 m r
27% m.a.c. Rudder) δ
0° 0.0000
10° 0.0004
20° 0.0009
30° 0.0021
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.
51

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. Use of flight simulation and dynamic analysis tools are
utilized for these concerns.
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, a large yawing moment will be created if the
135,000 ft-lbs
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 at a takeoff speed of 1.2 times
the stall speed at sea level. In this configuration the Vendetta can
45,000 lbs off
center
maintain a 953 fpm climb at military power and 3,435 fpm at
maximum afterburning thrust. This performance is overkill, but is
Figure 10.4 - OEI Forces and Moments
driven by the RFP requirement for zero foot per second specific excess power at a load factor of two.
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 lb
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 V-tail was shown to be ill-advised. 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
52

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 RCS-friendly layouts as well as increasing complexity and cost.
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.
Figure 10.5 - Vendetta Empennage Configuration
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 10.6 shows that as the aircraft
exceeds the critical Mach number, the center
Mach trim
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
mg
Figure 10.6 - Mach Tuck Illustrated
12% m.a.c.
was calculated with the Air Force’s Data Compendium (DATCOM) methods. The Vendetta cannot maintain trimmed
flight with the CG any further forward than 9% stable configuration. The aircraft would not have enough control power.
The use of a fuel monitor and a DFCS will be used to control the Vendetta throughout the flight envelope.
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. This is required 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 advanced
fighters being designed today utilize such systems. 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
53

ehaviors and to adhere to MIL-8785C. It was calculated that the Vendetta’s fuselage forebody 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.
-0.6
Lift Coefficient (C L )
-0.4
UNSTABLE
NEUTRAL
-0.2
-0.25 -0.2 -0.15 -0.1 -0.05 0 0.05 0.1 0.15 0.2 0.25
0
0.2
M = 0.2
M = 1.6
0.4
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 (Mach 0.2)
regime, it will want to continue to rotate and break away. In the supercruise, the aircraft behaves much more linearly.
The subsonic characteristics are of some concern, but even simple feedback schemes in the DFCS 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 or fluttering 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 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 conformity with the military specifications for
handling quality is shown in Table 10.IV.
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Table 10.IV - Longitudinal and Lateral Dynamic Mode Conformity with MIL-8785C
Damping ratio (ζ) Natural Frequency (ω n )
Mode 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 has 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 remain there in the supercruise.
From inertia computations illustrated in the weights and balance section (Section 7), it became apparent that the
Vendetta has a very small inertia that would need to be overcome to roll. This is due to the wings being the only
significant structure located off the centerline. This makes for very favorable roll damping and allows for the flaperon
and aileron configurations to be driven by the sizes required for high lift augmentation as presented in the aerodynamics
section. The final sizes and parameters for the empennage and roll control are presented in Table 10.V.
Table 10.V – Empennage Surfaces
Surface Area (ft 2 ) Control Surface
Horizontal
Stabilator
270.0 Full-Flying
Vertical
Stabilizer
165.0
Rudder
@ 27% m.a.c.
10.1 Simulation
Validation of a large supersonic aircraft like Vendetta is difficult due to limitations in experimental tools.
Subsonic wind tunnel models would be limited to testing takeoff and landing aerodynamics and would be inaccurate due
to Reynolds number discrepancies. Because of this, flight simulation was utilized to test the design of the aircraft. The
Cal Poly Flight Simulator was used to evaluate handling qualities, ground handling, up-and-away tasks, and low speed
performance. The flight simulator consists of a flight cab and instrument panel as shown in Figure 10.8 and Figure 10.9.
55

Figure 10.8 - Pheagle Simulator
Figure 10.9 - Flight Cab and Instruments
Desktop computers running a Windows operating system and two analog computers control the instrumentation,
force- feedback, and control inputs. The simulation architecture is built using Simulink, though most of the
computationally intensive components such as the six-degrees-of-freedom (6DOF) model are written in C++ as S-
Functions. The equations of motion used in the 6DOF are based on NASA Dryden equations of motion.
A non-linear aerodynamics model was created for Vendetta and implemented in the simulator using a table
lookup system. This system allows aerodynamic force and moment coefficients to be looked up using a series of user
defined tables. The force coefficients for each of Vendetta’s flying surfaces were defined as functions of Mach number,
relative airflow angle, and control surface deflections. Moment coefficients were calculated based on the forces and
moment arms of each surface. The longitudinal moment arms varied with CG and neutral point locations. Drag build up
data was used to accurately model the variation of zero-lift drag coefficient with Mach number and altitude (due to
Reynolds number variation). Additional fuselage force and moment contributions as well as linear dynamic stability
derivatives and downwash at the horizontal tail were calculated using DATCOM and incorporated into the model. A
total of 10 control surfaces were modeled in the simulation: left and right elevator, rudder, aileron, leading edge flaps,
and trailing edge flaps. Center-of-gravity location and landing gear extension were also modeled using control inputs.
The simulation model was built from an aerodynamic point of view to avoid building predefined stability and
control performance into the simulation. For example, rather than defining a stick-fixed neutral point location for the
configuration, the aerodynamic forces and moments that define the neutral point were modeled. The resulting simulation
is only limited by the accuracy of the aerodynamic data. Because no experimental methods could be used to obtain data,
the data is most likely inaccurate in extreme conditions such as high angles-of-attack or sideslip angles or under highly
dynamic flight conditions.
56

Additional components were integrated into the simulation model or modified from existing components to meet
Vendetta’s exact specifications. The engine deck included in this report was integrated into the flight simulator by
implementing code to lookup, uncorrect, and output the thrust and fuel flow values for the current flying condition and
throttle setting. Fuel flow was integrated during the simulation to accurately model the consumption of fuel during a
flight and its effect on the weight and moment of inertias of the aircraft. The landing gear model calculates the external
forces produced by each landing gear leg based on its position and properties. Friction, braking, and steering are
modeled allowing the ground handling qualities of Vendetta to be simulated and evaluated. Additional systems such as a
thrust reverser model, crash detector, and nonlinear actuators were utilized in the simulator. The simulator uses 3DLinx,
an OpenGL based graphics package as shown in Figure 10.10. It provides pilot feedback and situational awareness by
modeling of terrain, runways, and other aircraft in addition to a heads-up-display (HUD) (Figure 10.11).
Figure 10.10 - Graphics and Environment
Figure 10.11 - Heads up Display
The results of the flight simulation indicate that the unaugmented Vendetta is a difficult aircraft to fly. The
aerodynamic model shows that the aircraft is statically unstable in subsonic conditions, however due to the high
moments of inertia, the time to double is large enough that it can be controlled by an experienced pilot. The addition of
simple pitch and yaw rate feedback greatly improved the handling qualities and reduced the workload on the pilot while
the control surfaces remained unsaturated. Clearly, sophisticated outer loop controls including an altitude hold, heading
hold, and a waypoint navigator would be required to complete the design mission. This result confirms the need to
include a DFCS on Vendetta. The results of simulated takeoffs and landings indicate that Vendetta can easily meet the
required RFP takeoff and landing runway lengths. The thrust reversers provide enough stopping power to bring the
aircraft to a stop without the use of wheel brakes on the NATO 8,000 ft runway modeled in the simulator at 3,000 ft
above sea-level. Takeoff is best achieved with only partial trailing edge flaps (15°), because the higher takeoff speed
57

allows Vendetta to remain on the front-side of the power curve. The additional angle-of-attack provided by the leading
edge flaps provides a margin for error during takeoff and landing, and is useful during slow speed turns.
Ground handling tasks performed with the second revision made it apparent that the loading on the nose gear was
too small. Because of this, the nose gear on the final Vendetta configuration was moved back 8.5 ft to take 8% of the
weight. This enhanced the ground handling qualities substantially.
Initial sizing of the vertical stabilizer for static stability yielded a rather small area. After flying this
configuration, it became very apparent that the lateral stability was inadequate. The vertical area was increased a 35 ft 2 .
This greatly increased lateral stability. As shown in Table 10.IV, Vendetta’s high frequency Dutch roll mode still
required attention. The addition of a rate-feedback yaw damper in the form of a washout compensator added damping
and made the Vendetta receive higher pilot ratings from the test pilots who flew the simulator.
58

11 Performance
11.1 Specific Excess Power Requirements
Compliance with RFP specific excess power requirements is best shown using specific excess power envelopes
including those required as measures of merit. Figure 11.1 shows the 1-g military specific excess power envelope. The
RFP requirement of 0 ft/s at Mach 1.6 and 50,000 ft is met with 33.7 ft/s specific excess power. The 1-g maximum
(afterburner) specific excess power envelope, in Figure 11.2, shows that the RFP requirement of 200 ft/s is met with a
value of 212.5 ft/s. This envelope also shows that the maximum Mach number at 36,000 ft measure of merit is 2.18.
The 2-g maximum specific excess power envelope, in Figure 11.3, shows that the RFP requirement of 0 ft/s is met with a
value of 8.0 ft/s. This requirement is design driver for the thrust produced by the propulsion system. The 5-g maximum
specific excess power envelope and maximum sustained load factor envelope required as measures of merit are shown in
Figure 11.4 and Figure 11.5 respectively.
70,000
60,000
RFP Requirement 0
50,000
P s = 0 ft/s
Altitude (ft)
40,000
30,000
Stall Limit
50
100
150
150
200
20,000
10,000
Flaps
200
300
400
q Limit
0
0 0.5 1 1.5 2 2.5
Mach
Figure 11.1 - 1-g Military Specific Excess Power Envelope at Maneuver Weight
59

70,000
60,000
50,000
RFP Requirement
200 ft/s
Stall Limit
P s = 0 ft/s
100
200
Altitude (ft)
40,000
30,000
20,000
10,000
Flaps
300
400
500
600
700
700
600
q Limit
36,000 ft
Mach 2.18
0
0 0.5 1 1.5 2 2.5
Mach
Figure 11.2 - 1-g Maximum Specific Excess Power Envelope at Maneuver Weight
70,000
Altitude (ft)
60,000
50,000
40,000
30,000
20,000
10,000
RFP Requirement 0 ft/s
Stall Limit
100
200
300
400
500
600
700
P s = 0 ft/s
600
q Limit
0
0 0.5 1 1.5 2 2.5
Mach
Figure 11.3 - 2-g Maximum Specific Excess Power Envelope at Maneuver Weight
60

70,000
60,000
50,000
Altitude (ft)
40,000
30,000
20,000
10,000
Stall Limit
500
400
200
300
P s = 0 ft/s
100
200
q Limit
0
0 0.5 1 1.5 2 2.5
Mach
Figure 11.4 - 5-g Maximum Specific Excess Power Envelope at Maneuver Weight
70,000
n = 1
60,000
Altitude (ft)
50,000
40,000
30,000
Stall Limits
2
3
4
5
20,000
10,000
7
6
q Limit
0
0 0.5 1 1.5 2 2.5
Mach
Figure 11.5 - Maximum Sustained Load Factor Envelope at Maneuver Weight
61

11.2 Turn Rate Requirement
The maximum instantaneous turn rate requirement of 8.0 deg/s at 15,000 ft and Mach 0.9 is shown in the
maneuverability diagram in Figure 11.6. The maneuverability diagram shows that the required turn rate can be sustained
with military power. The maximum sustainable turn rate using afterburner is 11.8 deg/s. The maneuverability diagram
at sea-level required as a measure of merit is shown in Figure 11.7.
30
n 2
3
4 5 6 7
r = 2,000 ft
25
4,000 ft
Turn Rate (deg/s)
20
15
10
AB P s = 0
Mil. P s = 0
RFP Requirement
8 deg/s
Stall Limit
6,000 ft
8,000 ft
10,000 ft
5
q Limit
0
0 0.5 1 1.5 2
Mach
Figure 11.6 - Maneuverability Diagram at 15,000 ft and Maneuver Weight
30
n 2
3
4 5 6 7
r = 2,000 ft
25
4,000 ft
Turn Rate (deg/s)
20
15
10
AB P s = 0
Mil. P s = 0
Stall Limit
6,000 ft
8,000 ft
10,000 ft
5
q Limit
0
0 0.5 1 1.5 2
Mach
Figure 11.7 - Maneuverability Diagram at Sea-Level and Maneuver Weight
62

11.3 Mission Requirements
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 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), 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 initial cruise condition is then found by drawing a flight path to the initial cruise conditions
that follows the maximum climb rate to fuel flow ratio. The resulting flight path and fuel consumption envelope are
shown in Figure 11.8.
55,000
50,000
45,000
40,000
Stall Limit
Initial Cruise Condition
dh/dW F = 0 ft/lb
10
20
Altitude (ft)
35,000
30,000
25,000
20,000
30
40
50
20
30
15,000
10,000
q Limit
5,000
0
0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2
Mach
Figure 11.8 - 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 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 49,000 ft for the initial cruise and increased by 3,000 ft for each successive cruise or dash segment resulting in a
final cruise altitude of 58,000 ft. The two dash segments occur at 52,000 ft and 55,000 ft both meeting the RFP
requirement to dash above 50,000 ft. The optimum loiter speed for maximum endurance was determined to be Mach
63

0.35 or 390 ft/s by finding the minimum total drag on the aircraft at sea-level and loiter weight. The resulting mission is
listed in Table 11.I including the fuel consumption by mission segment and the corresponding RFP mission segments.
By completing the design mission, the requirements for a supercruise Mach number of 1.6 and mission radius of
1,750 nm 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 was also used to optimize
certain aspects of the mission such as cruise altitudes. Table 11.II lists the results of the mission simulation and Figure
11.9 shows a breakdown of fuel consumption by mission segment
Table 11.I - Design Mission
Detailed Mission Segment
Fuel
RFP Mission
Segment
Warm-up 2 min at idle thrust 89 lb 1a 89 lb
Takeoff – Accelerate to takeoff speed 270 ft/s
170 lb
Accelerate to Mach 0.80 at maximum military thrust
609 lb
1b 779 lb
Climb to 17,500 ft and accelerate to Mach 0.88
807 lb
Climb to 32,000 ft and accelerate to Mach 1.59
3,715 lb
Climb to 37,500 ft and accelerate to Mach 1.66
626 lb 2 6,128 lb
Climb to 47,000 ft at Mach 1.66
948 lb
Climb to 49,000 ft and decelerate to Mach 1.6
32 lb
Cruise 1,000 nm at 49,000 ft and Mach 1.6 16,139 lb 3 16,139 lb
Climb to 52,000 ft at Mach 1.6 774 lb 4 774 lb
Dash 750 nm at 52,000 ft at Mach 1.6 10,613 lb 5 10,613 lb
Descend to 50,000 ft at Mach 1.6
39 lb
Turn 180º at n = 1.25
867 lb
Drop 4 × 2,000 lb JDAMs
0 lb
6 1,425 lb
Climb to 55,000 ft at Mach 1.6
519 lb
Dash 750 nm at 55,000 ft and Mach 1.6 8,814 lb 7 8,814 lb
Climb to 58,000 ft at Mach 1.6
401 lb
Cruise 1,000 nm at 58,000 ft and Mach 1.6
10,268 lb
9 10,669 lb
Descend to Sea-Level and decelerate to loiter speed 391 ft/s 885 lb 10 0 lb *
Loiter 30 min at Sea-level and 390 ft/s 2,453 lb 11 2,453 lb
Decelerate to landing speed 270 ft/s 174 lb – –
Land – Decelerate to zero speed 27 lb – –
Unload non-fixed equipment (2 × AMRAAMs and crew 1,280lb) 0 lb – –
* The RFP specifies no fuel used in descent 58,968 lb 57,882 lb
Table 11.II - Mission Results
Total fuel consumption
58,968 lb
Mission radius
1,750 nm
Total distance traveled over mission 4,100 nm
Total mission duration
5 hr. 6 min.
Takeoff weight
125,051 lb
Empty weight
56,797 lb
Fuel weight (total fuel onboard) 58,974 lb
Maneuver weight
95,624 lb
Landing weight
58,077 lb
Average cruise lift to drag ratio 6.55
Cruise Back
17%
Reserve
4%
Dash Back
15%
Misc.
7%
Dash Out
18%
Accelerate
& Climb
11%
Cruise Out
28%
Figure 11.9 - Fuel Consumption over Mission
64

11.4 Takeoff & Landing
The RFP requires that the aircraft be able to takeoff and land on an icy standard NATO runway 8,000 ft 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 the thrust reverser during landing. The takeoff and landing profiles used in the simulation are shown in
Figure 11.10 and Figure 11.11.
Altitude (ft)
60
50
40
30
20
10
0
V stall = 146 knots
Max. Tire Speed = 210 knots
µ roll = 0.025
Roll
V TO = 160 knots
V 50 = 196 knots
Climb
V 50 > 175 knots
Pull-up
n = 1.15
Rotate 3 sec.
0 500 1,000 1,500 2,000 2,500 3,000 3,500 4,000 4,500 5,000
Distance (ft)
Figure 11.10 - Takeoff Profile
Altitude (ft)
60
50
40
30
20
10
0
7,000
Approach
Flare
n = 1.15 V TD = 160 knots
Roll 3 sec.
6,000
V 50 = 177 knots
V 50 > 175 knots
5,000
4,000 3,000
Distance (ft)
Figure 11.11 - Landing Profile
V stall = 146 knots
Max. Tire Speed = 210 knots
Brake
2,000
µ brake = 0.3 Dry
µ brake = 0.1 Ice
- 25% Mil. Thrust
1,000
0
MIL-C5011A defines field length to be the distance required to takeoff and clear a 50 ft obstacle or the distance
to land from a 50 ft obstacle. Takeoff and touchdown speed are defined as 1.1 times the aircraft’s stall speed, and the
speed over the 50 ft obstacle must be greater or equal to 1.2 times the stall speed for both takeoff and landing. The
takeoff gross weight for the design mission of 125,051 lb 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 1.1 times the 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
obstacle was cleared before the climb angle was reached, so the climb segment of the profile was ignored. Also, to
65

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 listed in Table 11.III and Table 11.IV show
that the RFP requirements for takeoff and landing on an icy 8,000 ft runway are met.
Table 11.III - Takeoff Results
Table 11.IV - Landing Results
Weight 125,100 lb Weight 125,100 lb
Maximum lift coefficient 1.16 Maximum lift coefficient 1.16
Stall speed 246 ft/s Stall speed 246 ft/s
Takeoff speed 271 ft/s Takeoff speed 271 ft/s
50 ft obstacle speed ≥ 290 ft/s 395 ft/s 50 ft obstacle speed ≥ 290 ft/s 395 ft/s
Rolling friction coefficient 0.025 Dry braking friction coefficient 0.3
Runway length 4,000 ft Icy braking friction coefficient 0.1
Field length over 50 ft obstacle 5,460 ft Thrust reverser effectiveness 25% Mil.
Dry runway length
4,030 ft
Dry field length over 50 ft obstacle 5,450 ft
Icy runway length
5,890 ft
Icy field length over 50 ft obstacle 7,310 ft
66

12 Payload
Weapon internal layout was a design
driver for the Vendetta. For small CG excursion
due to weapons deployment all stores were
initially positioned as close to the CG 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 in length after room
Figure 12.1 - L to R configurations 1, 2, 3
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 revision increased the maximum cross sectional
area by 5ft 2 and shortened the length of the aircraft to 95ft.
The next iteration of the design utilized the existing 180in MPRL out of the B-1B and shown in Figure 12.2. This
caused the final configuration to grow to 103ft in length and maximum cross sectional area of 88ft 2 utilizing the proven
rotary launcher would decrease development costs and time. It also allows Vendetta to perform several alternate
missions outlined in the next section.
In an effort to ascertain the feasibility of RFP
delineated weapons as supersonic deployment candidates,
each weapon system was analyzed. Foldout 5 shows that only
one of the weapons has been wind tunnel tested for supersonic
Figure 12.2 - 180 inch MPRL
deployment. Retrofitting the weapon systems with a ballute and sabot, shown in Figure 12.3, would aide in supersonic
stability. The use of a weapons bay supersonic flow deflector, an acoustical
resonance damping system, and a flow modification system may be needed to aid
in weapons deployment.
Figure 12.3 - Ballute and Sabot
67

Standard ten degree fall clearance is maintained for all weapons. The weapons bay doors were designed to rotate
into the bomb bay and not into the free stream. This is illustrated in Foldout 2 – FS 688.9. 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.
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
ejector racks illustrated in Figure 12.4. The rack has electrically fired impulse
cartridges, a gas operated mechanism, and is designed to forcibly eject
Figure 12.4 - 30in Ejector Rack
conventional or nuclear weapons in the 4000 lb weight class. The LAU-142A
ejector is used with the AIM-120C shown in Figure 12.5.
Weapons guidance is accomplished with
the RFP GFE ICNIA providing GPS/INS
guidance data (all weapons), an AN/APG-77
RADAR system (for the AIM-120), as well as
an on-board second generation Tessa Infrared
Search and Track System (IRSTS) system (for
the GBU-27). The IRSTS can only be utilized in
alternate subsonic mission due to line of sight
Figure 12.5 - LAU-142A Ejection Sequence
and range limitations. More detailed information on weapons can be found in Folodout 5.
12.1 Alternate Missions
In addition to the design mission, the Vendetta can perform alternate missions. The MPRL, shown in Figure 12.6,
carried by the Vendetta allows it to carry a total of 8 × 2,000 lb bombs (Figure 12.6) compared to the 4 required for the
design mission (no AMRAAMs can be carried in this configuration.) The weapons bay designed for the MPRL is only
4 inches greater in diameter than a previous custom design that carried only the RFP loadout. The extra cross-sectional
area of to the MPRL results in an extra 1,300 lb of fuel consumption over the design mission; however, the added
weapons capability and the fact that the MPRL is proven equipment, justify its use. The performance of the Vendetta
over four alternate missions was calculated. Fully loaded missions and subsonic missions flown at Mach 0.85 and an
altitude of approximately 30,000 ft were considered. The results shown in Table 12.I indicate that only a small loss of
68

ange occurs due to the additional weight of 8 × 2,000 lb bomb loadout, and the range of the aircraft can be greatly
extended by flying subsonic (although it extends the mission duration to 11 hours.) A subsonic ferry mission was also
considered using the storage space in the MPRL for additional fuel capacity. If 16,000 lb of additional fuel are carried in
the weapons bay, the total ferry range of the Vendetta can be extended to 6,200 nm allowing it to be quickly and easily
transported anywhere in the world without the need for tanker aircraft or multiple refueling stops.
The use of the RFP unspecified AGM-158A (JASSM), which would require no modification of the MPRL, offers
an extension of combat mission radius by over 100nm. This low observable weapon is seen as the future of ALCM’s
(Figure 12.7)
Figure 12.6 - MPRL with 8 × 2,000 lb JDAMs
Table 12.I – Alternate Mission Results
Design Mission
Mission Radius
1,750 nm
Takeoff Weight
125,100 lb
Mission Time
5 hr. 6 min.
8 × 2,000 lb bombs – Supersonic
Mission Radius
1,590 nm
Takeoff Weight
133,100 lb
Mission Time
4 hr. 47 min.
4 × 2,000 lb bombs – Subsonic
Mission Radius
2,500 nm
Takeoff Weight
125,100 lb
Mission Time
11 hr. 13 min.
8 × 2,000 lb bombs – Subsonic
Mission Radius
2,400 nm
Takeoff Weight
133,100 lb
Mission Time
10 hr. 52 min.
Subsonic Ferry – 16,000 lb additional fuel
Total Range
6,200 nm
Takeoff Weight
133,100 lb
Mission Time
13 hr. 40 min.
Figure 12.7 – MPRL with 8 × AGM-158A (JASSM)
69

(4) Mk-84 LDGP + (2) AIM-120
(4) GBU-27 + (2) AIM-120
(4) 2000lb JDAM +(2) AIM-120
(4) AGM-154 JSOW + (2) AIM-120
(16) 250 lb Small Smart Bomb
AIM-120 C AMRAAM
Weapon Weight
1967 lb
Weapon Weight
2165 lb
Weapon Weight
2100 lb
Weapon Weight
1064 lb
Weapon Weight
250 lb
Weapon Weight
327 lb
Installed Configuration Weight
10222 lb
Installed Configuration Weight
11014 lb
Installed Configuration Weight
10754 lb
Installed Configuration Weight
6610 lb
Installed Configuration Weight
5500 lb
Installed Configuration Weight
5500 lb
Weapon Length
Weapon Diameter
Tail Span
Max Drop Height
Max Tested Drop Velocity
Guidance
Weapon Information:
12.6 ft
18 in
2 ft
Unlimited
M=1.3
Ballistic
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.
Weapon Length
Weapon Diameter
Tail Span
Max Drop Height
Max Tested Drop Velocity
Guidance
Weapon Information:
13.9 ft
14.6 in
2 ft
Unlimited
Unknown
Semi-Active
Laser
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.
Weapon Length
Weapon Diameter
Tail Span
Max Drop Height
Max Drop Velocity
Guidance
Weapon Information:
13.2 ft
18 in
2 ft
Unlimited
M=1.3 tested
GPS / INS
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 a I-1000 (1000lb)(452.5kg) 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 bomb
bay's.
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 10m (32.8ft) to 15m
(49.2ft) of their targets. Wind tunnel tests in 1996 are reported to
have cleared JDAM for release at up to M 1.3.
Weapon Length
Weapon Diameter
Tail Span
Max Drop Height
Max Drop Velocity
Guidance
Weapon Information:
14 ft
21 in
24 in
Unlimited
Subsonic
GPS / INS
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 cross-section 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.
Weapon Length
Weapon Diameter
Max Drop Height
Max Drop Velocity
Guidance
Weapon Information:
8.2 ft
6 in
Unlimited
Unknown
GPS / INS
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.
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 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.
Vendetta
Chris Droney
Nate Schnaible
Rev. 3
Weapon Length
Weapon Diameter
Fin Span
Max Drop Height
Max Drop Velocity
Guidance
Kolby Keiser
Chris Maglio
12 ft
7 in
INS
High Rollers
1 ft 6 in
Unlimited
Supersonic
Command
from Launch
Aircraft
Monopulse
Radar Seeker
Foldout 5
Weapon Systems
Chris Atkinson
Dan Salluce
5/23/02

13 Cockpit
Cockpit design began with the RFP requirement for a crew of two. A comparison of tandem versus side-by-side
seating arrangement, and its affect on cross sectional area, was conducted. A solid model was constructed with room
provided for instrumentation, controls, circuit breakers and military aft pilot vision requirements (MIL-STD-850B)
allowing the frontal cross-sectional area of each configuration to be determined. The results are shown in Figure 13.1.
Figure 13.1 - Cockpit Width Trade Study
The results show that the configuration has very little effect on frontal area of
the cockpit. For this reason, other factors were taken into account before a
final decision on pilot configuration was made. The use of the 180 in MPRL,
favored the side-by-side seating arrangement because the width of the
fuselage was already driven to be large. This arrangement allowed greater
pilot communication as well as the elimination of many redundant circuit
Figure 13.2 – Fuselage Comparison
breakers and instruments; however, preliminary stability and control analysis
revealed a need to narrow the forward fuselage due to the undesirable C mα characteristics of a wide nose section.
Therefore the decision was made to utilize a tandem seating configuration. This configuration offered a better
field of vision for the primary pilot and decreased the width of the forward fuselage as shown in Figure 13.2. Utilizing
this information the detailed 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 military vision
specifications outlined in Table 13.I.
71

Table 13.I - Military Vision Specifications
5.1.1 Forward Pilot Vision
azimuth (°) up (°) down (°)
0 10 11
20 20
30 25
90 40
135 20
5.1.2 Aft Pilot Position
0 5
11°
5°
Figure 13.3 - Virtual Cockpit Model
Further vision refinement produced a rectilinear vision plot shown in Figure 13.4. Canopy reinforcing structure
was placed such that the view angles between 25 and 40 degrees up were unobstructed thus allowing vision for in-flight
refueling. Every effort was made to increase downward vision to aid in ground handling, takeoff, and landing.
Takeoff and Landing
vision is inherently limited
in supersonic aircraft due
the required low forward
fuselage angles. In an effort
to reduce pilot workload,
multifunction
displays
(MFD) can be used to
enhance pilot vision. The
vendetta display layout is
Figure 13.4 - Rectilinear Vision Plot of Forward Cockpit Position
illustrated in Figure 13.5. MFD 1 incorporated into the glare shield and upper instrument panel will be used in takeoff
and landing to increase downward vision, meshing seamlessly with the actual cockpit over nose view. It would also
utilize infrared or night vision to enhance visibility during night and poor weather
operation, thus increasing the all weather capability of the Vendetta. MFD’s 2, 3,
and 4 display moving map imagery, flight critical data, and mission critical
information. The standard dash mounted HUD was replaced by a current helmet
mounted HUD system thus increasing the pilot’s situational awareness.
Figure 13.5 - Cockpit Displays
72

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. NASA
studies have shown that and human life functions become critically affected by the lower pressure and oxygen content of
the upper atmosphere. These studies outline how physiological effects such as
the bends and hypoxia are accelerated by the extremely low temperatures of
high altitudes. These effects will render a human unconscious in seconds, and
dead in minutes. Balancing these effects against the economics and long
preparation and turn around times associated with full pressure suits, the
decision was made to utilize a partial pressure suit. The advanced fighter
crew protection system shown in Figure 13.6 was specifically developed for
missions in this altitude range. It offers low unit cost in comparison to full
pressure suits in addition to low turn around time by alleviating the necessity
for a suiting procedure similar that used in the U-2. Pilot oxygen is provided
by the RFP GFE OBOGS.
Figure 13.6 - Advanced Fighter
Crew Protection System
The K-36D ejection seat was chosen because its ejection
envelope (Figure 13.7) encompasses that of the Vendetta. The seat
is manufactured by Boeing North American who subcontracts
Zvezda, the seats original designer, for further development. A
shaped charge cutting system was chosen over a canopy ejection
for high safety throughout the RFP mission.
Figure 13.7 - K-36D Performance Envelope
73

14 Systems
The systems of the Vendetta will closely follow the design architecture of the F-22; however, technology
advances by 2020 will render most of the electronics aboard the F-22 antiquated, 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) and a selfcontained
energy storage system (SES). The
RFP lists a GFE APU but the cost, volume,
and weight listed suggests that it is actually a
Ram Air Turbine (RAT). The RAT was
Table 14.I – APU Selection Table
Company Model
Startup
Ceiling
Dry
Weight Rating Power/Weight
ft lb Hp Hp/lb
Honeywell 131-9A 41000 350 460 1.31
Honeywell 36-300 35000 300 390 1.3
Honeywell 331-200 43000 500 579 1.16
Pratt and
Whitney PW901 25000 835 1535 1.84
Sundstrand APS2100 37000 270 504 1.87
Sundstrand APS3200 39000 305 603 1.98
eliminated due to the need for ground power (RFP specifies the aircraft must operate with minimal ground support) and
previous service experience showing the unreliability of RATs in supersonic aircraft. APU selection involved examining
a number of mid-sized gas turbine generators with output exceeding the estimated minimum 350 Hp required by the
Vendetta. A shortened list of APU candidates appears in Table 14.I.
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 and small size was seen as the main APU
selection criteria. With this in mind the Sundstrand APS
3200 was selected. By the RFP 2010 cutoff data a new
APU may become available; therefore the Sundstrand is
mainly used for sizing considerations rather than a
definite APU for the vendetta. Placement of the APU
can be seen in Figure 14.1 and Foldout 4.
Modern designs utilize a SES for power backup
Figure 14.1 - APU Placement
due to its high reliability, invulnerability to aircraft
attitude and airspeed. The SES is a design point to be focused on in the next level of design. Current hypergolically
74

fuelled systems offer high power to weigh ratio but fuels are hazardous and expensive. Next generation systems should
be available by 2010.
14.2 Vehicle Management System
The vehicle management system (VMS) manages following systems: control stick, throttle lever, rudder pedals,
air data probes, accelerometers, and actuators.
Four separate actuator systems were initially considered.
1) Electrohydrostatic
2) Electric
3) Pneumatic
4) Hydraulic
Pneumatic systems were eliminated due to low power-to-weight ratios, large comparative size, and low power
transmission efficiencies. Electrohydrostatic actuators offer many benefits such as a line replaceable units, optimized per
service dynamic impedance shaping, and optimized K factor; Electrohydrostatic actuators create low observability and
weight problems as a consequence of the need for electric power transmission and the loss of the heat sink capabilities of
hydraulic lines. An all electric actuator system has the same low observability problems as the electrohydrostatic system
as well as a comparatively large and heavy actuator.
Thus the decision was made to utilize an all hydraulic system with digital fly-by-wire control. The F119 engine
utilized in the F-22 has a PTO driving two 72 gpm pumps for a total of four pumps that supply two independent 4000 psi
systems. The high pressure system was chosen because of 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.
A schematic of the hydraulic system is shown in Foldout 4.
The Smiths Industries 270 volt, direct current (DC) electrical system was chosen identical to that used in the F-22.
It uses two PTO driven 65 kilowatt generators as shown in Foldout 4. 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 a Fail Op/Fail Op/Fail Safe or better level of redundancy. GPS/INS and other navigational systems should be
examined and a proper redundancy levels chosen.
14.3 Fuel System
Initial fuel system design began with the placement of fuel tanks symmetrically about the CG. The final
configuration provides 58,974 lb of fuel capacity. 80% of the fuselage tank volume and 70% of wing tank volume was
75

utilized for fuel capacity. The remaining fuel tank volume was left for self sealing linings, structure, and fuel expansion
(See Foldout 2). All tanks in the aircraft are pressurized with nitrogen gas from the onboard inert gas generating system
(OBIGGS). Pressurization is minimal due to structural constraints and the low vapor pressure of JP-8. Nitrogen reduces
the concentration of fuel vapor and thus the chance of an 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 a port on the starboard side of the forward fuselage
(Foldout 4). This fueling point shares common lines with the AAR retractable fueling boom port located on the upper
portion of the forward fuselage. Both ports offer fueling rates as high as 1,100 gpm, which is the maximum KC-10 fuel
probe refueling rate (Table 14.II). Fuel is directed to the forward fuselage tank and then power transferred to the
remaining three tanks. A gravity feed system can be utilized in flight in case of a failure of the power feed system. All
major thermal transfer within the aircraft is performed by the fuel system. The air cooled fuel cooler utilizes inlet duct
boundary layer and flow control diversion air to dissipate all kinetic heating experienced by the airframe as well as heat
generated by onboard systems. Dual heat exchangers are utilized for combat survivability. Fuel flow rate 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. Vendetta is fitted with a halon fire suppression system.
Table 14.II - Fuel System Sizing Requirements
Fuel System Sizing Requirement
Fuel Flow Rate
GPM
CG Shift Requirement Between #3 and #4 Fuel Tanks 100
KC-10 Probe Maximum Refueling Rate 1100
(2) F-119 Turbofan @ Max Thrust With Reheat 294
Air Cooled Fuel Cooler (ACFC) 400
76

14.4 Government Furnished Equipment
Table 14.IIIprovides a list of GFE and identifies weither or not the equipment was utilized. Some GFE was not
used because it was not applicable to the design. A brief description of the reasoning is also provided
Table 14.III - List of Government Furnished Equipment
Government Furnished Equipment Application Utilized
ICNIA INS/GPS guidance and weapons targeting YES
3 x MFD Crew situation awareness enhancement YES
Heads-Up Display Replaced by Helmet Mounted Display NO
Data Bus New System Architecture Replaces Data Bus with VMS NO
Vehicle Management System Manages Aircraft Systems YES
IRSTS (Tessa Derivative) Low Speed and Altitude Targeting for GBU-27 YES
Active Array Radar Use the AN/APG-77 LPI System Due to LO Requirement NO
LANTIRN Navigation System Not a Low Altitude Night Aircraft NO
LANTIRN Targeting System Not a Low Altitude Night Aircraft NO
HARM Targeting System Does Not Carry HARM Missiles Because of LO Design NO
Electrical System Engine Control YES
Auxiliary Power Unit (APU) Description in RFP is of a Ram Air Turbine NO
Ejection Seat Standard Seat Will Not Cover Flight Envelope. Use K-36D NO
OBOGS Oxygen Generation for Pilot YES
OBIGGS Nitrogen Generation for Fuel System YES
AIM-120 AMRAAM Advanced Medium Range Air to Air Missile YES
M61A1 20 mm Cannon Supercruise Mission Eliminates Viability of Cannon NO
14.5 Maintenance and Servicing Plan
Vendetta is optimized for lowest possible integrated combat turnaround times focusing on open maintenance
access for all major components and fast weapon, fuel, and crew refurbishment. Maintenance panels are located
regularly across the bottom and top sides of main fuselage sections for easy open access to all serviceable equipment.
RCS considerations are taken into account for placement and geometry of access paneling. Major component placement
allows field servicing without major overhaul equipment jigs and specialized oversized tooling.
Utilizing a hydrant pressurized fueling rate of 1100 gallons per minute fueling time is only 7.9 minutes. More
common fueling rates of 100 gallons per minute drive fueling time to 87 minutes. Combat weapon refurbishment takes
approximately 10 minutes per weapon or one hour for the RFP mission loading. Therefore with RFP weapons loadout,
turnaround times between 60 and 87 minutes, depending upon the available refueling rate, are possible. Due to the
choice of partial pressure suit, crew suiting and preparation is far below this time.
77

15 Manufacturing
Ease of manufacturing was a design criterion for the
Vendetta. Part commonality will reduce 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 control 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
Figure 15.1 - Routing Tunnel
would be interconnected through two separate routing tunnels in the fuselage of the aircraft. This would reduce
installation complexity and system redundancy and therefore reduce the amount of labor involved in the installation
process. The routing tunnel is shown in Figure 15.1.
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 and Foldout 1.
The entire propulsion system is capable of
being dropped out the bottom of the aircraft which
provides for easy installation and maintenance.
Figure 15.2 - Manufacturing Breaks
Computer-aided manufacturing will enable more
complex parts to be machined by computernumerically-controlled
machining tools. Large items such as bulkheads can be machined from a single piece of metal.
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 be required to meet the stringent structural load limits. These access panels will ease maintenance
and reduce maintenance hours required per flight hour.
78

The assembly line would allow for major components, such as the wing, fuselage, and empennage to be prefabricated
at other site locations and brought in to a central assembly line as shown below in Figure 15.3.
Figure 15.3 - Assembly Line
79

16 Cost Analysis
The final and most important issue in the purposed development of the Vendetta is the cost analysis. Before any
aircraft can win a contract it must be reasonably priced. 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 (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, there are more aircraft available to help pay the $6.5 billion cost associated with RDT&E. Furthermore, there
80

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 in Figure 16.3 - Lifecycle
Cost.
Figure 16.3 - Lifecycle Cost
81

17 Conclusion
The Vendetta presented by The High Rollers from California Polytechnic State University, San Luis Obispo is the
optimum solution to the AIAA 2001/2002 Undergraduate Team Aircraft Design Request for Proposal. The Vendetta is
designed to replace the stealthy F-117 Nighthawk and B-2 Spirit as well as the supersonic F-15E Strike Eagle and B-1
Lancer as a supersonic stealth interdictor. The configuration of Vendetta was created through the use of design
philosophies such as simple low observable shaping, weight and balance centered design, detailed mission and
performance simulation, and realistic systems integration. Vendetta utilizes current and future technologies to provide
the best possible performance with at minimum cost. The use of RadBase 2 software throughout the design process
allowed the creation of a stealth frontal aspect. The use of a solid modeling program throughout the design process
allowed for faster design iterations to be performed while evaluating the RCS. The use of the Cal Poly Flight Simulator
aided in the design process allowing for problem areas to be quickly identified and fixed. As shown in Table 17.I, it
meets or exceeds all the design requirements.
82

Table 17.I - RFP Compliance Checklist
RFP Requirement Met Page# RFP Requirement Met Page#
Crew Performance Requirements
• Two pilot cockpit design 71 • 0 ft/sec specific excess power at 1-g mil. Thrust (1.6 M/50,000 ft) 59
Maintenance • 200 ft/sec specific excess power at 1-g max. thrust (1.6 M/50,000 ft) 60
• Easy access to and removal of major systems 78,79 • 0 ft/sec specific excess power at 2-g max. Thrust (1.6 M/50,000 ft) 60
Structure • 8.0 deg/sec instantaneous turn rate (0.9 M/15,000 ft) 62
• +7, -3 vertical g’s (clean, 50% fuel) 38 Weapons Carriage
• 2,133 psf max. dynamic pressure (Mach 0.2, sea-level) 38 • (4) Mk-84 LDGB 67,FO5
• 1.5 factor of safety on all design ultimate loads 36 • (4) GBU-27 + (2) AIM-120 67,FO5
• 12,000 hour service life 80-81 • (4) 2,000 lb JDAM + (2) AIM-120 67,FO5
Fuel/Fuel Tanks • (4) AGM-154 JSOW + (2) AIM-120 67,FO5
• Design fuel is JP-8 (6.8 lb/gal) FO2,75 Measures of Merit
• All fuel tanks self-sealing and retained throughout mission 75 • Weight Summary (TOGW, W e , W f , W/S, T/W, W f /W) 7
Stability • Aircraft Geometry
• Closed loop static/dynamic stability meets MIL-F-8785B 47-57 ◦ Wing/control surface area 19
• Static margin within +10% and -30% limits 42-45 ◦ Fuselage size and volume FO1
• Digital control system for longitudinally unstable designs 47-57 ◦ Frontal cross sectional area distribution 24-25
Balanced Observables ◦ Wetted area 25
• Balanced radar, IR, visual, acoustical, and electromagnetic signatures 13, FO3 ◦ Inlet and diffuser 31-34
• Front aspect RCS less than 0.05 m 2 against 1-10 GHz radar 13, FO3 • Systems integration
• All stores carried internally 67,FO5 ◦ Landing gear 40-41
Operation ◦ Weapons carriage 67-69,FO5
• All weather operations and weapons delivery from NATO runways,
◦ Sensor and avionics locations FO1, FO2
65,72
shelters, facilities
◦ Cockpit 71-73
Cost • Mission duration, radius, fuel burn by mission segment 64
• Flyaway cost less than 150 million (200 unit buy) 80 • Takeoff and landing distance (standard and icy conditions) 65
• Minimize life cycle costs 80 • Performance at 50% internal fuel
Engine Deck ◦ Max. Mach at 36,000 ft 60
• Include an engine data package 26-35, 84 ◦ 1-g max. thrust specific excess power envelope 60
Mission Performance ◦ 2-g max. thrust specific excess power envelope 60
• Weapons Load – (2) AIM-120 + (4) 2,000 lb JDAM 67-69,FO5 ◦ 5-g max. thrust specific excess power envelope 61
• Takeoff fuel for warm-up and acceleration (sea level, 59°F) 64 ◦ Max. thrust sustained load factor envelope 61
• Climb from sea level to optimum supercruise altitude 64 ◦ Max. thrust maneuvering performance diagrams at sea-level 62
• Supercruise out 1,000 nm at M=1.6 and optimum altitude 64 ◦ Max. thrust maneuvering performance diagrams at 15,000 ft 62
• Climb and Dash out 750 nm above 50,000 ft at M=1.6 64 • Fly away and life cycle costs (cost trades 100 to 1,000 units) 80-81
• Drop (4) 2,000 lb JDAM’s, turn 180° at M=1.6 at 50,000 ft 64 • Design Drawings
• Dash back 750 nm above 50,000 ft at M=1.6 64 ◦ Detailed three view FO1
• Supercruise back 1000 nm at M=1.6 and optimum altitude 64 ◦ 3-D perspective FO1
• Descend to sea level (no distance or fuel credit) 64 ◦ Internal layouts FO2
• Reserve fuel for 30 min. at sea level and maximum endurance speed 64 ◦ Materials selection FO4
83

Appendix A – Engine Deck
Part Power Data Idle Maximum Military
Altitude Mach -5,691 -2,846 0 2,846 5,691 8,537 11,383 14,228 17,074 19,920 22,765 25,611 28,457 31,302 Thrust Fuel Flow Thrust Fuel Flow
0 0 -1,460 -327 806 1,939 3,072 4,216 5,386 6,590 7,854 9,216 10,674 12,203 13,936 16,006 1,312 1,329 30,000 15,008
0 0.2 -1,888 -520 847 2,214 3,581 4,838 6,062 7,358 8,736 10,244 11,835 13,457 15,353 16,698 1,792 1,707 28,925 15,634
0 0.32 -1,723 -323 1,078 2,478 3,879 5,168 6,442 7,803 9,267 10,867 12,502 14,249 16,360 18,473 1,529 1,830 27,334 15,554
0 0.45 -1,350 31 1,411 2,790 4,170 5,503 6,872 8,333 9,909 11,585 13,287 15,232 17,508 19,783 1,274 2,029 25,938 15,495
0 0.6 -567 694 2,013 3,215 4,475 5,812 7,200 8,762 10,825 12,666 14,681 17,026 19,371 21,716 354 2,165 22,563 14,528
0 0.75 216 1,357 2,616 3,640 4,781 6,121 7,528 9,192 11,742 13,747 16,074 18,819 21,234 23,650 -566 2,302 19,188 13,561
0 0.9 999 2,020 3,219 4,065 5,086 6,430 7,856 9,621 12,658 14,829 17,468 20,613 23,097 25,583 -1,485 2,438 15,813 12,593
0 1.1 2,042 2,903 4,022 4,632 5,493 6,842 8,293 10,194 13,880 16,270 19,327 23,005 25,581 28,161 -2,712 2,620 11,313 11,303
0 1.3 3,086 3,787 4,826 5,198 5,900 7,254 8,731 10,767 15,101 17,711 21,185 25,396 28,066 30,738 -3,938 2,803 6,813 10,013
1,500 0 -1,454 -322 811 1,944 3,078 4,221 5,390 6,594 7,858 9,220 10,678 12,206 13,939 16,009 1,285 1,323 29,990 15,003
1,500 0.2 -1,817 -455 907 2,269 3,631 4,883 6,102 7,393 8,765 10,268 11,855 13,470 15,362 16,702 868 1,323 28,919 15,638
1,500 0.32 -1,640 -246 1,148 2,542 3,936 5,219 6,486 7,840 9,297 10,891 12,519 14,259 16,366 18,472 357 1,323 27,323 15,553
1,500 0.45 -1,286 89 1,464 2,839 4,214 5,541 6,904 8,360 9,931 11,602 13,299 15,239 17,510 19,781 -293 1,323 25,926 15,492
1,500 0.6 -552 708 2,026 3,227 4,487 5,823 7,209 8,771 10,833 12,672 14,686 17,030 19,374 21,719 -1,526 1,323 22,553 14,526
1,500 0.75 -378 879 2,194 3,392 4,648 5,980 7,356 8,921 10,986 12,804 14,851 17,207 19,562 21,917 -1,890 1,323 23,460 15,409
1,500 0.9 -204 1,050 2,362 3,557 4,809 6,137 7,504 9,071 11,138 12,935 15,015 17,383 19,749 22,115 -2,253 1,323 19,181 16,292
1,500 1.1 28 1,278 2,586 3,777 5,023 6,347 7,700 9,272 11,342 13,110 15,235 17,618 20,000 22,380 -2,737 1,323 13,476 17,469
1,500 1.3 260 1,506 2,810 3,996 5,238 6,556 7,896 9,472 11,545 13,285 15,455 17,853 20,250 22,644 -3,221 1,323 7,771 18,646
5,000 0 -1,447 -314 820 1,954 3,087 4,231 5,400 6,605 7,870 9,232 10,690 12,219 13,952 16,022 1,310 1,342 29,969 15,002
5,000 0.2 -1,873 -505 861 2,229 3,596 4,853 6,078 7,374 8,752 10,260 11,852 13,473 15,369 16,715 1,789 1,721 28,895 15,633
5,000 0.32 -1,714 -313 1,089 2,491 3,893 5,183 6,458 7,819 9,285 10,886 12,522 14,271 16,384 18,497 1,526 1,840 27,299 15,550
5,000 0.45 -1,344 37 1,419 2,801 4,183 5,518 6,888 8,351 9,930 11,607 13,311 15,259 17,536 19,814 1,273 2,037 25,902 15,493
5,000 0.6 -562 697 2,022 3,217 4,477 5,812 7,199 8,760 10,822 12,662 14,675 17,019 19,363 21,708 351 2,173 22,494 14,471
5,000 0.75 216 1,416 2,680 3,817 5,016 6,293 7,612 9,121 11,129 12,891 14,882 17,181 19,481 21,780 -1,323 2,097 23,438 15,409
5,000 0.9 887 2,137 3,472 4,636 5,887 7,226 8,640 10,209 11,977 13,768 15,763 18,105 20,448 22,790 -3,127 2,006 22,303 15,419
5,000 1.1 1,782 3,098 4,528 5,729 7,048 8,470 10,012 11,660 13,109 14,938 16,938 19,337 21,736 24,137 -5,532 2,150 20,790 15,432
5,000 1.3 2,678 4,060 5,584 6,821 8,208 9,714 11,383 13,110 14,240 16,108 18,113 20,569 23,025 25,483 -7,938 2,294 19,277 15,445
5,000 1.5 3,573 5,021 6,640 7,914 9,369 10,958 12,754 14,561 15,371 17,278 19,288 21,801 24,314 26,830 -10,343 2,438 17,764 15,459
10,000 0.2 -2,046 -622 800 2,222 3,645 4,902 6,105 7,395 8,758 10,272 11,844 13,438 15,424 17,412 1,132 1,624 31,603 17,702
10,000 0.32 -1,702 -299 1,104 2,506 3,909 5,203 6,483 7,848 9,315 10,913 12,555 14,326 16,457 18,589 1,524 1,855 27,269 15,547
10,000 0.45 -1,330 51 1,433 2,814 4,195 5,530 6,892 8,340 9,918 11,608 13,325 15,289 17,577 19,865 1,271 2,050 25,870 15,488
10,000 0.6 -546 714 2,034 3,236 4,496 5,831 7,206 8,762 10,852 12,692 14,697 17,004 19,310 21,616 348 2,184 22,464 14,473
10,000 0.75 60 1,387 2,714 3,873 5,115 6,448 7,867 9,429 11,188 12,962 14,894 17,151 19,408 21,664 -1,327 2,106 23,361 15,350
10,000 0.9 1,046 2,118 3,190 4,366 5,644 7,017 8,487 10,083 11,810 13,654 15,802 17,950 20,098 22,246 -3,127 2,014 22,267 15,417
10,000 1.1 2,535 3,626 4,710 5,882 7,158 8,550 10,044 11,763 13,696 15,677 17,658 19,639 21,620 23,602 -7,450 1,817 18,845 14,927
10,000 1.3 2,766 3,831 4,883 6,001 7,239 8,616 10,103 11,865 13,919 16,030 18,141 20,253 22,133 23,898 -8,735 1,518 15,159 12,516
10,000 1.5 2,489 3,620 4,729 5,884 7,173 8,629 10,197 12,093 14,359 16,690 19,022 21,355 23,687 26,019 -8,413 1,323 10,943 9,938
20,000 0.32 -1,653 -252 1,148 2,548 3,948 5,241 6,521 7,886 9,348 10,931 12,572 14,362 16,512 18,661 1,521 1,897 27,195 15,537
20,000 0.45 -1,280 98 1,475 2,853 4,231 5,563 6,927 8,376 9,943 11,620 13,349 15,334 17,632 19,930 1,267 2,088 25,793 15,475
20,000 0.6 -511 748 2,071 3,268 4,527 5,863 7,240 8,800 10,883 12,719 14,744 17,075 19,405 21,735 346 2,219 22,399 14,468
20,000 0.75 180 1,422 2,748 3,905 5,146 6,477 7,887 9,446 11,214 13,001 14,945 17,204 19,463 21,722 -1,328 2,128 23,293 15,350
84

20,000 0.9 581 2,164 3,231 4,407 5,682 7,042 8,502 10,114 11,850 13,696 15,828 17,961 20,094 22,226 -3,130 2,039 22,153 15,359
20,000 1.1 2,644 3,960 5,277 6,593 7,758 9,033 10,426 12,025 13,826 15,730 17,635 19,540 21,444 23,348 -7,399 1,855 19,459 15,419
20,000 1.3 2,819 4,052 5,285 6,518 7,612 8,840 10,206 11,824 13,708 15,721 17,735 19,748 21,242 22,476 -8,456 1,622 18,502 14,710
20,000 1.5 2,406 3,485 4,552 5,666 6,890 8,271 9,811 11,649 13,815 16,138 18,460 20,783 23,106 25,429 -8,158 1,415 14,426 11,790
20,000 1.6 1,585 2,730 3,875 5,020 6,393 7,718 9,284 11,166 13,297 15,745 18,193 20,641 23,089 25,536 -6,624 1,209 14,060 11,047
20,000 1.8 878 2,133 3,389 4,644 6,036 7,415 8,984 10,967 13,170 15,615 18,060 20,504 22,949 25,394 -4,683 1,323 11,658 9,161
25,000 0.32 -1,540 -146 1,247 2,641 4,034 5,320 6,593 7,951 9,406 10,982 12,616 14,401 16,542 18,684 156 1,323 27,157 15,542
25,000 0.45 -1,190 183 1,556 2,928 4,301 5,629 6,987 8,431 9,993 11,664 13,388 15,368 17,661 19,954 -483 1,323 25,754 15,479
25,000 0.6 -473 785 2,106 3,302 4,560 5,895 7,270 8,829 10,910 12,746 14,770 17,098 19,427 21,756 -1,694 1,323 22,353 14,460
25,000 0.75 241 1,386 2,683 3,851 5,105 6,448 7,868 9,440 11,219 13,017 14,974 17,245 19,516 21,787 -2,988 1,323 23,247 15,344
25,000 0.9 960 2,173 3,242 4,421 5,698 7,061 8,523 10,137 11,876 13,724 15,858 17,993 20,097 22,264 -4,907 1,323 22,111 15,357
25,000 1.1 417 1,936 3,456 4,976 6,419 7,972 9,650 11,523 13,594 15,795 17,997 20,199 22,401 24,602 -3,994 1,323 19,358 15,357
25,000 1.3 1,097 2,485 3,873 5,260 6,580 8,028 9,630 11,472 13,566 15,826 18,086 20,346 22,607 24,867 -5,229 1,323 19,052 15,129
25,000 1.5 2,312 3,386 4,453 5,566 6,779 8,142 9,688 11,516 13,654 15,992 18,330 20,669 23,007 25,345 -8,190 1,323 16,036 12,845
25,000 1.6 1,958 3,014 4,071 5,128 6,377 7,616 9,089 10,876 12,911 15,253 17,596 19,938 22,281 24,623 -7,402 1,323 16,032 12,133
25,000 1.8 1,012 2,210 3,408 4,607 5,923 7,245 8,755 10,678 12,821 15,200 17,579 19,958 22,337 24,715 -4,953 1,323 13,327 10,036
25,000 2 72 1,448 2,746 4,082 5,463 6,864 8,408 10,465 12,712 15,125 17,702 20,280 22,858 25,436 -3,115 1,323 11,048 8,207
30,000 0.45 -1,224 156 1,536 2,915 4,295 5,630 6,995 8,447 10,015 11,694 13,425 15,411 17,712 20,012 1,265 2,148 25,708 15,485
30,000 0.6 -452 808 2,128 3,326 4,586 5,921 7,298 8,859 10,931 12,747 14,818 17,242 19,667 22,092 340 2,274 22,309 14,465
30,000 0.75 230 1,472 2,801 3,956 5,198 6,529 7,935 9,491 11,246 13,044 15,013 17,273 20,220 21,793 -1,333 2,184 23,193 15,341
30,000 0.9 627 2,207 3,271 4,440 5,711 7,071 8,527 10,127 11,870 13,728 15,868 17,445 19,021 20,597 -3,134 2,081 22,061 15,358
30,000 1.1 2,642 3,845 5,068 6,292 7,515 8,846 10,308 11,957 13,797 15,796 17,795 19,794 21,793 23,792 -7,402 1,887 19,303 15,357
30,000 1.3 2,763 3,918 5,072 6,227 7,381 8,662 10,112 11,788 13,705 15,824 17,943 20,062 22,180 24,298 -8,405 1,662 18,993 15,124
30,000 1.5 2,359 3,412 4,464 5,561 6,747 8,080 9,616 11,421 13,515 15,854 18,194 20,532 22,871 25,211 -7,863 1,525 17,744 14,053
30,000 1.6 2,312 3,284 4,256 5,228 6,358 7,513 8,897 10,593 12,538 14,780 17,020 19,262 21,503 23,743 6,345 6,614 18,078 13,311
30,000 1.8 1,065 2,213 3,362 4,510 5,759 7,030 8,488 10,360 12,450 14,771 17,091 19,411 21,732 24,052 5,165 5,524 15,171 11,027
30,000 2 -64 1,287 2,561 3,873 5,229 6,606 8,124 10,161 12,384 14,771 17,323 19,875 22,427 24,979 4,032 4,434 12,796 9,091
30,000 2.2 -1,764 -27 1,709 3,445 5,193 6,564 7,650 9,878 12,181 14,453 16,726 18,999 21,272 23,544 2,682 3,344 10,313 7,184
36,089 0.6 -211 988 2,187 3,386 4,643 5,977 7,355 8,921 10,987 12,807 14,855 17,213 19,569 21,927 336 2,327 22,250 14,464
36,089 0.75 443 1,608 2,775 3,941 5,192 6,532 7,948 9,515 11,281 13,072 15,065 17,400 19,734 22,069 -1,334 2,229 23,132 15,353
36,089 0.9 952 2,232 3,298 4,471 5,745 7,107 8,563 10,157 11,894 13,771 15,925 17,524 19,124 20,723 -3,131 2,119 21,983 15,353
36,089 1.1 2,620 3,790 4,958 6,127 7,389 8,756 10,264 11,946 13,812 15,871 17,930 19,989 22,047 24,107 -7,401 1,918 19,228 15,363
36,089 1.3 2,740 3,847 4,953 6,059 7,249 8,562 10,062 11,773 13,709 15,891 18,073 20,255 22,437 24,619 -8,407 1,685 18,867 15,070
36,089 1.5 2,365 3,419 4,479 5,585 6,765 8,086 9,638 11,440 13,508 15,875 18,241 20,608 22,974 25,341 -7,821 1,554 18,134 14,370
36,089 1.6 2,563 3,512 4,461 5,410 6,476 7,610 8,965 10,630 12,545 14,743 16,940 19,138 21,336 23,534 7,799 7,301 19,642 14,524
36,089 1.8 1,602 2,634 3,667 4,699 5,811 6,968 8,305 10,060 12,032 14,226 16,419 18,613 20,808 23,003 6,531 6,142 17,495 12,347
36,089 2 388 1,613 2,763 3,950 5,180 6,432 7,823 9,737 11,838 14,101 16,525 18,950 21,375 23,799 5,416 5,060 14,759 10,116
36,089 2.2 -1,785 -155 1,475 3,104 4,755 6,191 7,523 9,762 12,112 14,495 16,878 19,261 21,644 24,027 4,578 4,121 12,280 8,168
36,089 2.4 -2,612 -803 1,006 2,814 4,620 5,976 6,987 9,287 11,622 13,862 15,903 17,945 19,987 22,028 3,405 3,183 9,958 6,422
43,000 0.75 517 1,682 2,848 4,014 5,265 6,605 8,022 9,592 11,361 13,154 15,150 17,493 19,837 22,181 -1,345 2,297 23,115 15,425
43,000 0.9 1,022 2,300 3,365 4,537 5,810 7,172 8,630 10,224 11,963 13,844 16,004 18,166 20,326 22,488 -3,151 2,178 21,970 15,420
43,000 1.1 2,676 3,845 5,014 6,184 7,446 8,813 10,323 12,008 13,877 15,939 18,002 20,065 22,127 24,189 -7,432 1,961 19,209 15,417
43,000 1.3 2,789 3,896 5,002 6,107 7,298 8,612 10,114 11,828 13,769 15,954 18,137 20,322 22,506 24,691 -845 9,348 18,843 15,114
43,000 1.5 2,402 3,456 4,515 5,621 6,802 8,123 9,679 11,485 13,559 15,927 18,295 20,663 23,032 25,400 -7,859 1,578 18,092 14,387
43,000 1.6 2,576 3,539 4,502 5,465 6,496 7,645 9,009 10,679 12,599 14,788 16,975 19,164 21,352 23,540 7,770 7,319 19,608 14,542
43,000 1.8 1,624 2,663 3,702 4,740 5,835 6,998 8,338 10,096 12,071 14,259 16,449 18,637 20,826 23,016 6,505 6,155 17,463 12,363
85

43,000 2 393 1,622 2,776 3,967 5,202 6,456 7,850 9,773 11,879 14,146 16,572 18,678 21,425 23,852 5,398 5,073 14,738 10,137
43,000 2.2 -887 500 1,889 3,277 4,694 6,083 7,575 9,705 11,987 14,375 16,762 19,150 21,537 23,925 4,562 4,131 12,259 8,186
43,000 2.4 -1,743 -237 1,268 2,773 4,295 5,740 7,250 9,511 11,888 14,318 16,748 19,178 21,608 24,038 3,796 3,284 9,934 6,431
50,000 0.75 589 1,759 2,927 4,096 5,351 6,695 8,115 9,688 11,461 13,257 15,256 17,603 19,950 22,298 -1,355 2,370 23,098 15,519
50,000 0.9 1,132 2,406 3,484 4,660 5,929 7,278 8,736 10,329 12,065 13,937 16,108 18,280 20,451 22,623 -3,153 2,283 21,954 15,511
50,000 1.1 1,080 2,449 3,819 5,190 6,627 8,186 9,891 11,769 13,831 16,076 18,320 20,566 22,811 25,056 -7,433 241 19,192 15,495
50,000 1.3 1,688 2,940 4,192 5,444 6,738 8,195 9,836 11,687 13,764 16,071 18,377 20,684 22,990 25,298 -8,469 465 18,829 15,174
50,000 1.5 2,571 3,612 4,657 5,763 6,886 8,211 9,760 11,555 13,619 15,959 18,299 20,639 22,980 25,320 -7,854 1,749 18,046 14,402
50,000 1.6 2,161 3,187 4,212 5,238 6,284 7,496 8,916 10,639 12,614 14,841 17,069 19,295 21,523 23,750 7,740 7,137 19,575 14,566
50,000 1.8 1,339 2,422 3,506 4,588 5,704 6,911 8,292 10,091 12,108 14,329 16,550 18,771 20,993 23,215 6,488 6,029 17,437 12,384
50,000 2 401 1,633 2,791 3,986 5,223 6,480 7,876 9,806 11,918 14,188 16,615 18,726 21,470 23,898 5,383 5,087 14,714 10,154
50,000 2.2 5 1,151 2,295 3,440 4,623 5,967 7,622 9,644 11,859 14,252 16,647 19,041 21,435 23,829 4,553 4,137 12,239 8,199
50,000 2.4 -443 681 1,804 2,928 4,086 5,546 7,487 9,632 11,977 14,524 17,272 19,617 22,768 25,517 3,787 3,294 9,919 6,445
55,000 0.9 1,180 2,460 3,527 4,701 5,976 7,340 8,801 10,404 12,148 14,022 16,200 18,378 20,556 22,733 -3,155 2,337 21,942 15,594
55,000 1.1 2,804 3,976 5,148 6,320 7,584 8,956 10,471 12,164 14,042 16,102 18,163 20,223 22,282 24,343 -7,434 2,087 19,185 15,562
55,000 1.3 2,888 3,997 5,106 6,215 7,407 8,725 10,231 11,955 13,905 16,089 18,273 20,457 22,642 22,480 -8,463 1,808 18,821 15,225
55,000 1.5 2,477 3,532 4,594 5,701 6,882 8,206 9,766 11,579 13,662 16,032 18,401 20,770 23,139 25,509 -7,872 1,645 18,037 14,446
55,000 1.6 2,636 3,600 4,565 5,529 6,562 7,713 9,078 10,755 12,685 14,878 17,071 19,263 21,456 23,649 7,730 7,368 19,547 14,585
55,000 1.8 1,666 2,708 3,748 4,790 5,885 7,050 8,390 10,157 12,142 14,343 16,544 18,745 20,946 23,146 6,465 6,191 17,407 12,393
55,000 2 409 1,643 2,805 4,003 5,242 6,500 7,895 9,831 11,952 14,242 16,698 19,154 21,610 24,067 5,371 5,101 14,697 10,168
55,000 2.2 22 1,170 2,316 3,464 4,642 5,983 7,638 9,686 11,949 14,419 16,888 19,358 21,827 24,297 4,539 4,153 12,224 8,211
55,000 2.4 -471 665 1,801 2,937 4,097 5,560 7,514 9,715 12,160 14,850 17,785 20,720 23,655 26,590 3,781 3,302 9,906 6,454
60,000 1.1 1,234 2,589 3,942 5,296 6,768 8,333 10,036 11,924 13,998 16,253 18,508 20,763 23,019 25,274 -7,433 406 19,173 15,654
60,000 1.3 1,826 3,056 4,286 5,516 6,873 8,315 9,951 11,811 13,898 16,224 18,550 20,876 22,245 23,135 -8,469 626 18,806 15,300
60,000 1.5 2,544 3,600 4,660 5,766 6,948 8,271 9,830 11,642 13,725 16,093 18,462 20,831 23,199 25,568 -7,873 1,713 18,028 14,500
60,000 1.6 2,274 3,285 4,297 5,308 6,388 7,585 8,998 10,722 12,699 14,939 17,179 19,418 21,658 23,898 7,729 7,227 19,539 14,634
60,000 1.8 1,414 2,492 3,569 4,646 5,779 6,979 8,355 10,158 12,180 14,417 16,653 18,890 21,127 23,364 6,460 6,091 17,394 12,424
60,000 2 449 1,682 2,842 4,038 5,276 6,532 7,925 9,859 11,980 14,267 16,722 19,177 21,632 24,087 5,362 5,132 14,671 10,178
60,000 2.2 1 1,155 2,309 3,463 4,649 5,997 7,658 9,714 11,984 14,460 16,937 19,413 21,890 24,366 4,527 4,152 12,208 8,221
60,000 2.4 -521 627 1,776 2,924 4,096 5,572 7,539 9,751 12,666 14,911 17,858 20,806 23,753 26,701 3,771 3,290 9,895 6,464
65,000 1.3 3,005 4,098 5,190 6,283 7,545 8,855 10,363 12,100 14,066 16,277 18,488 20,698 22,908 25,119 -8,470 1,939 18,803 15,396
65,000 1.5 2,576 3,637 4,700 5,808 6,989 8,312 9,876 11,700 13,796 16,164 18,533 20,901 23,269 25,636 -7,874 1,738 18,022 14,568
65,000 1.6 2,738 3,701 4,664 5,627 6,656 7,803 9,169 10,854 12,798 15,001 17,203 19,405 21,608 23,810 7,723 7,456 19,533 14,695
65,000 1.8 1,740 2,779 3,819 4,857 5,950 7,111 8,451 10,223 12,229 14,467 16,705 18,942 21,180 23,418 6,453 6,250 17,387 12,468
65,000 2 449 1,685 2,850 4,050 5,293 6,553 7,951 9,894 12,048 14,406 16,969 19,531 22,094 24,657 5,352 5,143 14,670 10,214
65,000 2.2 8 1,171 2,334 3,497 4,683 6,032 7,690 9,718 11,948 14,375 16,803 19,231 21,659 24,088 4,515 4,181 12,190 8,235
65,000 2.4 -510 648 1,805 2,963 4,125 5,595 7,545 9,691 12,029 14,558 17,281 20,003 22,725 25,446 3,761 3,321 9,876 6,470
70,000 1.6 2,801 3,764 4,726 5,688 6,718 7,864 9,229 10,910 12,846 15,066 17,284 19,504 21,722 23,942 7,571 7,455 19,436 14,680
70,000 1.8 1,749 2,797 3,846 4,894 5,995 7,161 8,506 10,283 12,281 14,575 16,870 19,164 21,459 23,753 6,378 6,267 17,338 12,485
70,000 2 443 1,687 2,866 4,078 5,327 6,589 7,990 9,940 12,076 14,523 17,172 19,820 22,469 25,118 5,348 5,174 14,669 10,255
70,000 2.2 24 1,189 2,355 3,522 4,707 6,052 7,714 9,775 12,051 14,608 17,165 19,721 22,278 24,834 4,515 4,205 12,191 8,266
86

Appendix B – Design Tools
Company Software Description Utilization
Solidworks
Solidworks
Parametric, Feature Based Two and
Three Dimensional Modeling and
Computer Aided Drafting
Aircraft Solid Model Generation
Nemetschek
VectorWorks
Two Dimensional Computer Aided
Drafting
Three View and Inboard Drawing
Generation
Microsoft
Excel/Visual Basic
Spreadsheet Utilizing Visual Basic for
Complex Numerical Analysis
Table Generation, Numerically
Integrated Simulated Flight Missions
Microsoft Word English Written Language Editing
Main Report Body and Layout,
Automated Table of Figures and
Tables
Adobe Photoshop Image Editing
Figure, Picture, and Table Touch-up
and Editing
The Mathworks
Matlab/Simulink
Mathematical Package with Built In
Graphical Simulation Tools
Six Degree of Freedom Simulation
Utilizing Non-Linear Aerodynamic
Models
The Mathworks
Simulink/C++
Programing Language, C++ Compiler
via Simulink
Creation of Simulation Customized
Aerodynamic and Control Code
Global Majic
3DLinx/OpenGL
2D and 3D Graphics Application
Programming Interface, OpenGL
Programming via 3dLinx
Creation of Simulated Environment
Graphics for Pilot Feedback
PDAS
Digital Datcom
Calculates Static Stability, High Lift
and Control, and Dynamic Derivative
Characteristics
Calculation of Aerodynamic and
Stability Characteristics for Simulator
Desktop
Aerodynamics
Linair
Aerodynamic Characteristics of Multi-
Element, Nonplanar Lifting Surfaces
Aerodynamic Lift Analysis
PDAS
Panair
Subsonic/Supersonic Panel Method
Based on Linear Aerodynamic Theory
Aerodynamic Lift Analysis
Surface Optics RadBase 2
Discreet
3D Studio Max
Radar Cross Section (RCS) and
Amplitude and Phase Data for Both
Complex Targets and Cultural
Features
Character Animation, Next
Generation Game Development, and
Visual Effects Production
Radar Cross Section Analysis, Spyder
Plot Generation
Solid Model Mesh Optimization for
RCS Anaylsis
87

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