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1. Final Report
Aircraft Design
ME 4770, C'08
Prof. D. Olinger
Dustin Bradway '08
Kyle Miller '09

2. 2
Contents
Introduction................................................................................................................................. 5
Specifications Table.................................................................................................................... 5
Dimensions and Detailed Specifications .................................................................................... 6
Background................................................................................................................................. 7
Mission Profile............................................................................................................................ 8
Initial Weight (W0) Estimation................................................................................................... 9
Trade Studies ............................................................................................................................ 10
Detailed Drawings .................................................................................................................... 12
Interior....................................................................................................................................... 13
Airfoil Selection........................................................................................................................ 14
Airfoil Performance .................................................................................................................. 15
Wing Area............................................................................................................................. 16
Aspect Ratio.......................................................................................................................... 16
Wingspan .............................................................................................................................. 16
Wing Sweep.......................................................................................................................... 16
Wing Taper Ratio and Root Chord....................................................................................... 17
Mean Chord Length .............................................................................................................. 17
Stall Behavior........................................................................................................................ 17
W/S Calculations ...................................................................................................................... 18
Cruise.................................................................................................................................... 19
Loiter..................................................................................................................................... 19
Landing/Stall......................................................................................................................... 19
Takeoff.................................................................................................................................. 19
Refined Weight (W0) Estimation.............................................................................................. 20
"Newton's Equations of Takeoff" ............................................................................................. 21
T/W Ratio and Fixed Engine Design........................................................................................ 22
Updated Wing Characteristics .................................................................................................. 23
Wing Area............................................................................................................................. 23
Wingspan .............................................................................................................................. 23
Wing Taper Ratio and Root Chord....................................................................................... 23
Mean Chord Length .............................................................................................................. 24
Tail Geometry........................................................................................................................... 24
Horizontal Tail Geometry..................................................................................................... 24
Horizontal Tail Area ............................................................................................................. 24
Horizontal Tailspan............................................................................................................... 25
Horizontal Tail Root Chord and Tip Chord.......................................................................... 25
Vertical Tail Geometry ......................................................................................................... 25
Vertical Tail Area ................................................................................................................. 25
Vertical Tail Height .............................................................................................................. 25
Wing and Tail Geometry Summary.......................................................................................... 26
Winglets.................................................................................................................................... 26
Structural Analysis.................................................................................................................... 27
Landing Gear ............................................................................................................................ 29

3. 3
Fuel Tanks................................................................................................................................. 30
Thrust-Drag Analysis................................................................................................................ 30
Stability Analysis...................................................................................................................... 32
Maneuvers................................................................................................................................. 34
Climb..................................................................................................................................... 34
Turn....................................................................................................................................... 35
Logo and Name......................................................................................................................... 35
Conclusion and Summary......................................................................................................... 36
Appendix A: Historical Comparison Data................................................................................ 37
Appendix B1: Initial Weight Estimate Iteration ....................................................................... 38
Appendix B2: Initial Weight Estimate Iteration ....................................................................... 39
Appendix C: Initial Weight Trade Studies................................................................................ 40
Appendix D: Artwork ............................................................................................................... 41
Appendix E: Airfoil Geometry Data......................................................................................... 47
Appendix F: XFOIL Analysis................................................................................................... 48
Appendix H: Calculated Drag................................................................................................... 50
Appendix I: Center of Gravity.................................................................................................. 51
Appendix J: Final Presentation Slides ...................................................................................... 52
Note: "Creativity" is denoted throughout by an asterisk (*) in the margin of the text.

4. 4
Table of Figures
Figure 1: Range Map ...................................................................................................................... 5
Figure 2: Historical design trends................................................................................................... 8
Figure 3: Mission profile ................................................................................................................ 9
Figure 4: Trade studies.................................................................................................................. 11
Figure 5: ProEngineer CAD model............................................................................................... 12
Figure 6: Concept art .................................................................................................................... 13
Figure 7: Interior layout................................................................................................................ 13
Figure 8: Cross-section of our NACA 64008a airfoil................................................................... 14
Figure 9: Airfoil performance plots .............................................................................................. 15
Figure 10: Wing sweep trends ...................................................................................................... 17
Figure 11: Leading edge flow separation at stall.......................................................................... 18
Figure 12: Moment coefficient about quarter-chord point ........................................................... 18
Figure 13: Fuel allocation............................................................................................................. 21
Figure 14: Takeoff capability........................................................................................................ 22
Figure 15: Installed PW308B engines .......................................................................................... 23
Figure 16: T-tail ............................................................................................................................ 24
Figure 17: Winglets....................................................................................................................... 27
Figure 18: Box spar....................................................................................................................... 27
Figure 19: Plot of shear force (N) and bending moment (N-m) throughout the wing.................. 28
Figure 20: Shear force (left) and bending moment (right) distributions in the wing.................... 28
Figure 21: Wing deflection........................................................................................................... 29
Figure 22: Fuel tank location and size.......................................................................................... 30
Figure 23: Coefficients of drag..................................................................................................... 30
Figure 24: Thrust-drag plot........................................................................................................... 32
Figure 25: Stability diagram ......................................................................................................... 34
Figure 26: Turn rate ...................................................................................................................... 35
Figure 27: Logo............................................................................................................................. 36

5. 5
Introduction
We have designed a two-pilot business jet, capable of transporting eight to ten passengers
and their baggage on routes on the order of Halifax-to-London and Los Angeles-to-Honolulu
(see range map below, showing capabilities from three major world cities). Our aircraft has a
range of 2,500 nautical miles, a cruise speed of 560 mph, and a cruise altitude of 45,000 feet. We
have planned for special situations and contingencies by building in a generous loiter time (one
hour), designing for use on short runways (< 5,000 ft), and specifying a modular interior so that
our aircraft can be used for other purposes (e.g., high-altitude photography). Our aircraft
employs composite materials in an effort both to keep weight down and to help examine the
feasibility of designing future aircraft in a similar manner.
Our aircraft will be designed for high reliability – to executives and companies, wasted
time is wasted money – and 24-hour readiness and operability. Twin engines allow for high
safety margins, and the aircraft will be able to fly and climb safely on a single engine. Our jet
features a generous fuel supply stored in the wings, air conditioning, soundproofing, and a full
glass cockpit. Following are the specifications for our new business jet.
Specifications Table
Aircraft type Small business jet
Aircraft purpose Intercontinental passenger travel
Crew number Two pilots
Estimated payload 2,000 kg (6-8 passengers and luggage)
Range 2,500 nautical miles
Propulsion system type Turbofan (2 x Pratt & Whitney 308B)
Cruise speed and altitude 485 kts at 45,000 feet (Mach 0.726)
Mission Takeoff – cruise – loiter – land
Loiter time 1 hour
Maneuverability Basic (climb, descend, turn)
Takeoff distance and speed ~ 4,000 feet at 135 kts
Stall speed 120 kts
Figure 1: Range Map
*

6. 6
Dimensions and Detailed Specifications
Specification Preliminary Interim Final Historical
General
Crew 2 2 2 ~ 2
Wcrew 200 kg 200 kg 200 kg -
Passengers 8 8 8 6 to 12
Payload 2,000 kg 2,000 kg 2,000 kg ~ 1,080 kg
Range 2,500 nm 2,500 nm 2,500 nm 1,500 to 3,000 nm
SFCcruise 0.5 0.5 0.5 0.5
SFCloiter 0.4 0.4 0.4 0.4
Loiter time 1 hour 1 hour 1 hour -
Maximum speed - - 524 kts (Mach 0.91) -
Cruise speed 485 kts 485 kts 485 kts (Mach 0.84) ~480 to 500 kts
Stall speed - 120 kts 120 kts 120 to 140 kts
Cruise altitude 45,000 ft 45,000 ft 45,000 ft 39,000 to 43,000 ft
W0 13,730 kg 14,543 kg 14,543 kg ~ 16,000 kg
T/W .0917 .0967 .0967 0.625 to 0.1
L/D 10.9 10.34 10.34 -
Oswald efficiency, e - 0.8258 0.8258 -
Wing loading, W/S - 82.09 lb/ft2
82.09 lb/ft2
60 to 95 lb/ft2
Wing
Airfoil type NACA 64008a NACA 64008a NACA 64008a -
Aspect ratio 7.461 7.461 7.461 7.25 to 9.10
Wingspan 19.96 m 15.98 m 15.98 m 12.2-17 m
Area, Swing 53.416 m2
34.26 m2
34.26 m2
24 to 48 m2
Wing taper, λ 0.45 0.45 0.45 0.4 to 0.5
Dihedral 5˚ 5˚ 5˚ 3-7˚
Wing incidence angle 1˚ 1˚ 1˚ -
Wing sweep 20˚ 20˚ 20˚ 14˚ to 31˚
Airfoil thickness 0.179m 0.179 m 0.179 m -
Mean chord, c 2.804 m 2.25 m 2.25 m -
Chordroot 3.7 m 2.95 m 2.95 m -
Chordtip 1.66 m 1.33 m 1.33 m -
Winglets Yes Yes Yes -
Tails
Tail type T-tail T-tail T-tail -
Horizontal
Tail sweep 25˚ 25˚ 25˚ 15˚ to 30˚
Tail taper 0.85 0.85 0.85 -
Aspect ratio 4 4 4 -
Tail span - 4.64 m 4.64 m -
Shtail - 5.37 m2
5.37 m2
-
Mean chord, c - - 1.16 m -
Vertical
Tail sweep 30˚ 30˚ 30˚ 35˚ to 55˚
Tail taper 0.7 0.7 0.7 -
Aspect ratio 1.0 1.0 1.0 -
Tail span - 2.71 m 2.71 m -
Shtail - 7.37 m2
7.37 m2
-
Mean chord, c - - 2.71 m -
Stability Analysis
Fuselage length - 17.69 m 17.69 m ~ 20 m
Length overall - - 18.61 m ~ 18 to 20 m
Height - - 6.60 m ~5.5 to 7.5 m
Fuselage diameter - - 2.35 m -
XNeutral Position - - 10.14 m -
XMost Forward - - 9.09 m -
XCenter of Gravity - - 9.58 m -

7. Background
We considered four aircraft as a basis for our own design. Examined in detail were the
following aircraft, with a summary of their major specifications:
Gulfstream G200 Learjet 60XR Cessna Sovereign Cessna Citation X
Range 3400 nm 2365 nm 2664 nm 3250 nm
Cruise speed 459 kt 466 kt 431 kt 595 kt
Cruise altitude 39,000 ft 41,000 ft 41,000 ft 41,000 ft
Max. ramp weight 35,600 lb 23,750 lb 30,550 lb 36,400 lb
Max. payload 4,050 lb 1,820 lb 2,500 lb 1,200 lb
Engines
2 Pratt &
Whitney Canada
306A
(6040lb each)
2 Pratt &
Whitney 305A
(4600lb each)
2 Pratt &
Whitney Canada
PW306C
(5690lb each)
2 Allison AE
3007C
(6400lb each)
Length 62.25 ft 58.7 ft 61.1 ft 72.1 ft
Wingspan 58 ft 43.8 ft 63.3 ft 63.9 ft
Wing area 369 ft2
264.5 ft2
510 ft2
527 ft2
Height 21.5 ft 14.5 ft 19.1 ft 18.9 ft
Each is an 8-12 passenger aircraft with two pilots. The overall design of the reference
aircraft also matched what we envisioned for our plane: low wing, high tail out of the way of
wing turbulence, two aft fuselage-mounted engines, and so on.
Averages of the above values have been computed and are used in some early
calculations later in this report. In addition to the four reference aircraft, we also examined a
larger pool of business jets of various sizes, from 5-passenger very light jets (VLJs) to large
planes capable of transporting 25 people distances exceeding 3700 nautical miles. Our goal with
this study was to determine whether trends can be discerned in aircraft development and
characteristics, to help us envision the "business jet of the future" and to ensure that our work
produces a realistic aircraft.
Some of our successful attempts at finding correlations can be seen below. Our aircraft's
final specifications are also presciently indicated in red, demonstrating that ours is a design that
successfully follows contemporary and historical design trends. It should be noted that our
adherence to trends is a result of following the tried-and-tested design process, not from an
attempt to "not deviate from the line."
*

8. 8
Figure 2: Historical design trends
Interestingly, when we attempted studies comparing changes over time, we found no
correlations – wingspan, passenger capacity, engine power, and other factors have not changed
markedly or predictably since the first business jets were produced in the 1960s.
The historical data used for these studies can be found in Appendix A: Historical
Comparison Data.
Mission Profile
Our aircraft's mission profile is very simple, consisting only of takeoff, climb, cruise
(with altitude variations as required by ATC), descent and landing. Our aircraft is also capable of
loiter for up to one hour (not pictured).
0
20
40
60
80
100
0 50 100 150
Length (ft, X) vs. Wingspan (ft, Y)
0
20
40
60
80
100
0 20000 40000 60000
Empty Weight (lb, X) vs. Wingspan (ft, Y)
0
10000
20000
30000
40000
50000
60000
0 10 20 30
Passengers (X) vs. Empty Weight (lb, Y)
0
2000
4000
6000
8000
10000
12000
14000
16000
0 50 100 150
Length (X) vs. Engine Power (Y)
*

9. 9
Figure 3: Mission profile
Initial Weight (W0) Estimation
The four reference aircraft matching our design requirements were examined in detail to
help prepare our estimate for W0; more information on these aircraft and other research can be
found in the Background section previous. We averaged maximum ramp weights for the four
aircraft, found it to be 14,322 kilograms, and used that as our starting point for estimating W0.
The following values were used in our initial weight estimate, most derived by rounding
off average values from the four reference planes:
 Range: 2,500 nm (allowing Halifax-to-London, NYC-to-LAX, and LAX-to-Honolulu)
 Cruise speed: 485 kts (250 m/s)
 Ccruise: 0.5 / hr (textbook)
 Cloiter: 0.4 / hr (textbook)
 L/D: 10.9 (textbook)
 Wcrew: 200 kg (two pilots and gear)
 Wpayload: 2000 kg (eight 220-lb passengers and 1,200 kg of cargo, baggage, gear, etc.)
 Loiter endurance: 1 hour (to allow flexibility in CEO schedules)
 Wingspan: 17.1 m (reference aircraft)
 Swetted: 510.6 m2
(preliminary sketches; see below)
Swetted was computed with the aid of our preliminary sketches (tails, wings, winglets) and
by calculating the surface areas of tubes (the main fuselage and two engines), a cone (fuselage
tail), and a hemisphere (the fuselage nose):
Swetted = Stails + Swings + Sbody tube + Snose + Stail cone + Swinglets + Sengines
0
10000
20000
30000
40000
0
250
750
1250
1750
2250
2500
Altitude
(feet)
Distance (nautical miles)
*

10. 10
To obtain an initial estimate of the weights in the different mission stages, historical trends and
equations were used as follows:
W1 = 0.97 W0
W2 = 0.985 W1 W5 = 0.995 W4
(where W1 = weight after take-off, W2 = weight after climb, W3 = weight after cruise, W4 =
weight after loiter, W5 = weight after landing).
Wf / W0 was then calculated with a 6% safety margin:
( )
We / W0 could then be computed using the formula provided in the textbook in Table 3.1:
And finally our initial weight estimate was updated:
Three iterations were necessary to bring our estimate to within a successive iterative error
of less than 0.5%. Our updated estimate for W0, using conventional construction materials, is
16,080 kilograms. For details, see Appendix B1: Initial Weight Estimate Iteration.
We have decided, however, to use a composite material construction, the details of which
are discussed in the Trade Studies section to follow. As a result, our final W0 estimate is actually
13,730 kilograms.
Trade Studies
We conducted several studies to determine the weight tradeoffs that would be required if
we changed our aircraft's specified range, payload capacity, and loiter time. Values were
computed with the use of our iteration algorithm, allowing for the quick and easy creation of
several data points in each study. The results are as follows, and details can be found in
Appendix C: Initial Weight Trade Studies.
*

11. 11
Figure 4: Trade studies
We reduced our specified payload early in the design process as a result of conducting
trade studies; our initial specification of 2,500 kg caused our aircraft to exceed our reference
aircraft average by too large a margin, and we now feel 2,000 kg is perfectly adequate given the
passenger capacity of our jet (and whatever luggage those passengers could possibly need to
bring). We are satisfied by where we sit on the range and loiter curves, feeling no need to
increase either specification; however, we note that loiter could be increased beyond its already
above-average value without incurring too much of a penalty.
We have also decided to use composite materials in our aircraft construction, in an effort
to keep rising weights down and to add an element of "futurism" to our design. As an example,
the Learjet 85, introduced in October of 2007, will feature an all-composite structure designed by
Grob Aerospace. We are of the opinion that composite aircraft, with declining material prices
and ever-advancing manufacturing processes, will become more common in the future.
To account for a composite structure, the We / W0 ratio is adjusted to 95% of its original
value with the following formula:
0
5000
10000
15000
20000
25000
30000
35000
40000
45000
0 1000 2000 3000 4000 5000
W0 vs. increasing range
0
5000
10000
15000
20000
25000
0 0.5 1 1.5 2 2.5
W0 vs. increasing loiter
0
5000
10000
15000
20000
25000
30000
0 1000 2000 3000 4000 5000
W0 vs. increasing payload
*

12. 12
W0 is then calculated as before. Our early estimate for an initial weight using composite
materials is 13,730 kilograms. This is 85% of the weight of a conventional aluminum structure.
Because of this, we have decided to use composite construction in our business jet.
Detailed Drawings
Full sets of detailed drawings are located in Appendix D: Artwork. A CAD model of our
aircraft was created with PTC's ProEngineer software. The dimensions used in the model are, of
course, the same as the final specifications we provide in this report. Here is one view of our
model:
Figure 5: ProEngineer CAD model
In the spirit of the design process (and simulating the role of an Art Department), we
have also produced "concept art" for our aircraft, to better illustrate our proposed aircraft's shape
and details. These graphics are entirely of our own creation, having been steadily modified
throughout the design process to represent our evolving design. For all the drawings one could
possibly desire, including a visual depiction of the evolution of our aircraft, see Appendix D.
Depicting details can be difficult in ProEngineer, but is quite simple with illustration software
(here, Adobe Flash). It should be noted, too, that these drawings are made exactly to scale.
*

13. 13
Figure 6: Concept art
Interior
Our aircraft's modular interior can be configured in a variety of seating arrangements, two
of which are depicted here. The red areas indicate exits. The rear of the cabin contains luggage,
loaded externally. An optional head is located directly behind the pilots.
Six seats:
Eight seats:
Figure 7: Interior layout
*

14. 14
Airfoil Selection
Our background research indicates that smaller-to-midsize business jets use a variety of
airfoil shapes for their wing cross sections, including the IAI Sigma 2, Cessna 7500 and the
NACA 64008a shown below. The NACA 64008a airfoil was chosen for our plane because of its
favorable characteristics when used in our type of aircraft – a thin airfoil is important when
flying at high subsonic Mach numbers, because it increases critical Mach and allows for lower
drag at higher flight speeds. A list of non-dimensional geometry data for a NACA 64008a airfoil
can be found in Appendix E: Airfoil Geometry Data.
Figure 8: Cross-section of our NACA 64008a airfoil
Typical flying conditions for our aircraft will be a cruise altitude of 45,000 ft (13,720 m)
and a velocity of 560 mph (250 m/s). To obtain realistic data for the NACA 64008a airfoil and
aircraft wing, we conducted analysis at those cruise conditions. The atmospheric conditions at
cruise altitude are shown here:
Properties at Cruise Altitude
Property SI Units English Units
Temperature 216.6 K 390°R
Pressure 15,327 N/m2
3095 psf
Density 0.24646 kg/m3
0.4623 slug/ft3
Dynamic Viscosity 1.42x10-5
kg/m-s .2969 sl/ft-s
Gamma 1.4
Gas Constant 287 J/Kg-K 1717.23 ftlbf/slug°R

15. 15
Airfoil Performance
The performance of the NACA 64008a airfoil was evaluated with XFOIL; details are
provided in Appendix F: XFOIL Analysis. Plots of the lift coefficient (Cl) and drag coefficient
(Cd) versus angle of attack are shown below.
Figure 9: Airfoil performance plots
These graphs are excellent references for deciding and confirming a wing's fixed angle of
attack. Historical data specifies that the typical angle of attack for a commercial aircraft is 1˚.
Referring to the table in Appendix F: XFOIL Analysis, our airfoil provides a lift coefficient of
0.1623 at α = 1˚. We verified that this value would be sufficient to overcome the weight of the
aircraft with the following equation:
,
where W0 = initial weight and . We calculated the minimum required lift coefficient
to be 0.04467, indicating that our jet will have no problem becoming airborne from the runway.
-0.8
-0.6
-0.4
-0.2
0
0.2
0.4
0.6
0.8
-4.5 -2.5 -0.5 1.5 3.5 5.5 7.5
Coefficient
of
lift
C
l
Angle of Attack
0
0.01
0.02
0.03
0.04
0.05
0.06
0.07
0.08
0.09
0.1
-4.5 -2.5 -0.5 1.5 3.5 5.5 7.5
Coefficient
of
Drag
C
d
Angle of Attack
*

16. 16
Wing Area
To calculate the initial size of the wing, the following equation was used:
,
where Swing = area of wing, W0 =initial weight of aircraft, and = wing loading.
The number we used for wing loading was the average of our reference planes: 70.475
lb/ft2
(344.087 kg/m2
). A (non-final) value of 53.42 m2
resulted for our wing area.
Aspect Ratio
To calculate the aspect ratio of the wing the following equation was used:
,
where AR = aspect ratio, b = wingspan, and c = chord length.
Since the wingspan, b, and chord length, c, were not yet known, historical data was
needed for initial aspect ratio estimation. This can be found in the course textbook, in table 4.1.
A provided formula,
AR = aMmax
C
allows the calculation of aspect ratio. Values of 7 and -0.02 were used for a and c respectively,
and a Mach number of 0.7267 was used (this Mach number was obtained by dividing the
aircraft's cruise velocity by the speed of sound at 45,000 feet).
The resulting final aspect ratio is 7.46, very close to those of our reference aircraft. This
is a value close to that of historical reference aircraft, and while it may be slightly high on an
absolute scale, our aircraft is nonetheless able to easily support the resulting increased wing root
structural forces (see Structural Analysis section later).
Wingspan
To calculate the wingspan, we used the following formula:
√
Our values for aspect ratio and wing yielded a (non-final) wingspan, b, of 19.96 m. This
value is consistent with and close to the wingspans of the reference planes.
Wing Sweep
Our aircraft will have a wing sweep of 20 degrees, based on historical trends. This value
was obtained from the textbook. The following is the plot used to estimate this value:

17. 17
Figure 10: Wing sweep trends
Wing Taper Ratio and Root Chord
A taper ratio, 0.45, chosen based on historical data, allows the root chord to be calculated
with the following equation:
where S is the wing reference area, b is the wingspan, and λ is the wing taper.
The resulting (non-final) root chord was 3.69 m. With this value, the wing's tip chord can
be calculated by manipulating the taper ratio formula and solving for Ctip:
The resulting (non-final) chord for the wingtip was 1.66 m.
Mean Chord Length
To calculate a mean chord length for our wing, we used the following equation:
̅
Our (non-final) mean wing chord length, c, was 2.804 m.
Stall Behavior
The NACA 64008a has a maximum thickness of 8% of its chord. This value is less than
14% and falls in the moderate airfoil thickness category. Because of its thickness, this airfoil
stalls at the leading edge, as shown below.

18. 18
Figure 11: Leading edge flow separation at stall
This airfoil's stall characteristics require a professional pilot familiar with leading-edge
stall behavior; in particular, during stall, the moment about the quarter-chord point changes
drastically, as shown here, and amateur pilots would likely be unable to maintain control of the
aircraft:
Figure 12: Moment coefficient about quarter-chord point
W/S Calculations
To size the aircraft wing accurately and safely, estimates must be made for the ratio of
W/S (weight to wing area). This value depends on the flight condition, and can vary substantially
during the flight. As a result, four values are calculated and compared, and the lowest (i.e., the
lowest wing loading) is selected for safety. Wing loading is important to determining an aircraft's
takeoff, stall and landing speeds, its cruise speed and, of course, its wing size.

19. 19
Cruise
⁄ √
Cruise values for ρ and V were used. CD0 was estimated at 0.015 for a streamlined jet,
and e0, the Oswald efficiency factor, was estimated using equation 12.50 in the textbook:
A value of 60.975 lb/ft2
resulted.
Loiter
⁄ √
The same values for ρ and V were used as in the cruise calculation. A value of 105.61
lb/ft2
resulted.
Landing/Stall
⁄
New values for ρ and V were used, assuming a generous landing altitude of 4,000 ft and a
stall speed of 120 kts, a good approximation of the stall speeds of our reference aircraft and
general trends. CLmax was approximated with Figure 5.3 from the textbook, assuming double-
slotted flaps, and was found to be around 2.5.
A value of 109.18 lb/ft2
resulted.
Takeoff
⁄
We have used the "alternative" (non-iterative) method to determine W/S for takeoff; the
iteration method produced contradictory values that would not allow us to complete the process.
In the above equation, SG is the takeoff distance (4,000 ft), g is acceleration due to gravity, and
the thrust-to-weight ratio was calculated (from Table 5.3) to be 0.23865 for our aircraft.
A value of 133.64 lb/ft2
resulted.
Each value was then corrected back to the "takeoff condition," i.e., in terms of the
aircraft's initial weight. The corrected values are presented below for ease of comparison.

20. 20
Takeoff Cruise Loiter Landing
133.64 lb/ft2
82.09 lb/ft2
145.22 lb/ft2
150.50 lb/ft2
Our takeoff W/S is very close to the historical trends given in the textbook in Table 5.5 – for jet
transports, 120 lb/ft2
is the norm. Our lowest wing loading value is the W/S for the cruise
condition. Using it to calculate our aircraft's wing area, we obtained the following:
( ⁄ )
This value is in perfect alignment with those of our reference aircraft; the Gulfstream G200 has a
wing area of 369 ft2
, and the others are spread to both sides of our value.
Refined Weight (W0) Estimation
Having pinned down some of our aircraft's specifications, we repeated our algorithm to
estimate its W0, this time with a few changes and updates. To summarize:
- The starting value for W0 was set to the final value from the first estimate.
- L/D was calculated based on power sizing estimates, as the inverse of T/Wcruise.
- Wingspan was updated to our new value, 15.98m.
- An additional weight segment was added, Wdescent, where
- We/W0 was improved with the formula in Table 6.1
- Composite construction was still used, resulting in the usual weight reduction.
The final value for W0 is now 14543.4 kg, a weight increase of 5.92% over our first estimate.
This difference is small enough that it does not warrant redoing calculations elsewhere in the
design process. Details from this iterative process can be seen in Appendix B2, following the
initial estimates.
The following chart gives an idea of how fuel use is distributed over the course of an
average mission for our aircraft:
*

21. 21
Figure 13: Fuel allocation
"Newton's Equations of Takeoff"
Simple equations of motion can be used to compute basic takeoff properties for our
aircraft. Both engines together produce 61,385 N of takeoff thrust (see the Engine Design section
that follows for details). For a takeoff weight of 15453 kg,
The resulting time to reach takeoff speed is
This translates to a takeoff roll of:
Thus, our aircraft can theoretically become airborne in 2,300 feet, although of course
friction and drag have been ignored here. Prevailing winds, runway conditions and other factors
would alter this somewhat, too.
The following graphic depicts our aircraft's theoretical minimum takeoff roll compared to
the lengths of the shortest runways at some of the world's major (and not so major) airports.
Taxi &
Takeoff
5%
Climb
5%
Cruise
76%
Loiter
11%
Descent
2%
Landing
1%
*
*

22. 22
Figure 14: Takeoff capability
It is apparent, then, that our aircraft will have no problem serving any destination
required by our customer base. While we have not analyzed braking and deceleration, it is a safe
assumption that, if our aircraft can be landed at an airport, it will be able to take off from that
airport, too.
T/W Ratio and Fixed Engine Design
With the completion of the initial weight estimate, the T/W ratio was calculated using the
following equation:
( )
Our cruise thrust-to-weight ratio is 0.0967. Multiplication of this fraction by the initial weight
yields the minimum thrust needed for our aircraft to fly at cruise conditions. Our resultant
minimum thrust is 1,400 kg, or 4,100 lbf.
Basing our work on data from similar business jets, we have selected a fixed engine
design that uses two Pratt & Whitney 308B high bypass turbofans. Each engine produces
comfortably more thrust than our minimum specification for cruise, an important consideration
for engine-out maneuverability and other emergency conditions. The following table contains the
specifications for the 308B model (per engine):
P&W 308B
Take-off thrust (per engine) 8242 lb 3738 kg
Cruise thrust (per engine) 7400 lb 3356 kg
Dry weight 1043 lb 473 kg
Length 6.3 ft 1.92 m
Width 3 ft 0.914 m
0 2000 4000 6000 8000 10000 12000
Logan
Worcester
T.F. Green
Newark
Honolulu
Frankfurt
Charles de Gaulle
LAX
Denver
Heathrow
Hong Kong
Runway length (feet)
Shortest runways

23. 23
Below, a depiction of our engines positioned inside the nacelles:
Figure 15: Installed PW308B engines
Updated Wing Characteristics
The slight increase in our initial weight estimate and the change in the wing loading
characteristics resulted in changes in the values for some of our wing geometries. Following are
the recalculated values.
Wing Area
A new, final value of 34.26 m2
resulted for our wing area.
Wingspan
√
Our values for aspect ratio and wing area yield a new, final wingspan, b, of 15.98 m. This
value is still consistent with wingspans of our reference planes.
Wing Taper Ratio and Root Chord
Our taper ratio remained the same, at 0.45. It allow calculation of the wing root chord:

24. 24
The resulting, final root chord is 2.95 m. With this value, the final wingtip chord can be
calculated: 1.33 m.
Mean Chord Length
Our new, final mean aerodynamic chord length, c, is 2.25 m.
Tail Geometry
Our aircraft has a T-tail configuration, chosen for several reasons. Despite the T-tail's
typical disadvantage of adding to aircraft weight (due to required extra structural strengthening),
its advantages of the T-tail outweigh the disadvantages. One such advantage is that the T-tail
puts the horizontal tail clear of wing wake and engine exhaust. Another is its aesthetically-
pleasing design. Overall, however, the T-tail results in higher efficiency and a smaller tail than
would be possible if it were of a different design. Depicted below is our T-tail design:
Figure 16: T-tail
Horizontal Tail Geometry
The sweep of the horizontal tail's leading edge has been set to 25 degrees. This value was
obtained from historical data; a trend in past aircraft has been to sweep the horizontal tail 5˚
further than the wings. This increase in sweep angle ensures that the tail stalls later than the
wing, important to maintaining control and maneuverability in adverse conditions. An increase in
tail sweep angle also increases its critical Mach number relative to the wing; this prevents the
loss of elevator effectiveness in case of shock formation.
The taper ratio of the horizontal tail has been set to 0.85, based on historical data for T-
tails.
Horizontal Tail Area
To calculate the initial size of the horizontal tail, the following equation was used:

25. 25
̅̅̅̅
,
where Swing = area of wing, ̅
̅
̅
̅ =mean chord length of wing, =tail volume coefficient and
LHT = length between wing and horizontal tail.
Our value for the tail volume coefficient was taken from historical data; this value is
0.085. The length between the wing and horizontal tail was estimated to be 6.34 m. A value of
5.37 m2
resulted for our horizontal tail area.
Horizontal Tailspan
To calculate the horizontal tailspan, we used the following formula:
√
Our values for aspect ratio and wing yield a tailspan, b, of 4.64 m. This value is
consistent with and close to the tailspans of reference aircraft.
Horizontal Tail Root Chord and Tip Chord
The root chord can be calculated as before, and results in a horizontal tail root chord of
1.25 m. The corresponding tailtip chord is 1.07 m.
Vertical Tail Geometry
The sweep of the vertical tail has been set to 30 degrees, again obtained using historical
data that indicates that vertical tails are swept 5-10 degrees further than the horizontal tail. The
increase in the sweep angle once again also increases the tail's critical Mach number relative to
the wing, preventing loss of critical yaw control during turbulence. The taper ratio of the vertical
tail has been set to 0.7, based on historical data for T-tails; T-tail vertical surface taper ratios are
in the range of 0.5 to 1.0, to provide adequate chord for the attachment of the horizontal tail and
associated control linkages.
Vertical Tail Area
Using a vertical tail volume coefficient taken from historical data (0.95), and an
estimated length from the wing to vertical tail of 6.34 m, we have assigned a value of 7.37 m2
for
the vertical tail area.
Vertical Tail Height
To calculate the height of our vertical tail, we used the following formula:
√
Our values for aspect ratio and wing yield a height, b/2, of 2.71 m.

26. 26
Vertical Tail Chords
The horizontal tail base chord is 3.19 m. The corresponding top chord is 2.23 m.
Wing and Tail Geometry Summary
Property SI Units English Units
Wingspan, b 15.98 m 53.46 ft
Mean chord length, c 2.25 m 7.37 ft
Wing area, S 34.26 m2
368.80 ft2
Aspect ratio, AR 7.461
Wing sweep angle 20˚
Taper ratio, λ 0.45
Root chord length, Croot 2.95 m 9.70 ft
Tip chord length, CTip 1.33 m 4.36 ft
Airfoil thickness 0.08
Tails
Vertical
Height 2.71m 8.91ft
Wing area, S 7.37m2
79.43ft2
Root chord length, Croot 3.19m 10.49ft
Tip chord length, CTip 2.23m 7.34ft
Vertical taper ratio, λ 0.7
Tail A.R., vertical 1.0
Vertical tail sweep 30˚
Horizontal
Span, b 4.64m 15.21ft
Wing area, S 5.37m2
57.85ft2
Root chord length, Croot 1.25m 4.91ft
Tip chord length, CTip 1.07m 3.49ft
Vertical taper ratio, λ .85
Tail A.R. 4
Horizontal tail sweep 25˚
Winglets
Our aircraft will employ winglets, as is common in modern business jets. While we will
not do detailed analysis on the benefits winglets impart, a few effects can be quantified in a basic
manner.
Winglets improve cruise speed, somewhere on the order of 5%. Maximum speed is not
increased by much, however. Rate of climb increases (see discussion in Maneuvers section) by
around 6%. Stall speed remains unaffected, and handling is improved.
*

27. 27
B
H
h
b
Winglets improve aerodynamic efficiency by reducing drag (they help to dissipate
wingtip vortices, a contributor to induced drag at wingtips, and increase Oswald efficiency by
around 10%). Only zero-lift drag increases marginally, because of a small increase in wetted area
(we account for this when required in this report).
The handling benefits offered by winglets are numerous, too: rudder yaw control
improves, heading overshoot is reduced, and stall speeds are lower.
Our winglets are depicted below:
Figure 17: Winglets
Structural Analysis
Research into the nature of composite (carbon fiber) materials in aviation yielded the
discovery that composite wings support loads in a markedly different manner than conventional
aluminum construction. The skin of a carbon fiber aircraft is much
more capable of supporting loads than an aluminum skin would be,
and so the interior structure is very different.
Without the tools at our disposal to completely alter the way
we conduct structural analysis, we have elected to use a box spar,
which, to a limited degree, simulates the skin effect and also the
tube spars used in some carbon fiber aircraft. Our composite spar is
depicted here.
The area of the spar
is BH – bh. Its moment of
inertia is:
[ ]
For structural analysis, we used an elliptic
loading, which is characterized by a distribution as
follows:
√ ( )
With a total aircraft weight of 14,543 kg, the load supported by each wing at cruise is
(14543 / 2) * 9.81 = 71,262 newtons. A safety factor of 3.5 is used henceforth.
Properties of carbon fiber
λ 275 GPa
ρ 1.75 g/cm3
Tensile strength 3.5 GPa
Compressive strength 1.25 GPa
Poisson's ratio 0.69
Shear modulus 15.15 GPa
Shear strength 55.15 MPa
Young's modulus 234.4 GPa
Front: Side: Top:
*
Figure 18: Box spar

28. 28
Integration of the distributed elliptic load results in the total reaction force that the
airframe must exert on the wing at the root. This value is 31,725 newtons.
Integration of the load multiplied by the distance at which it acts results in the total
reaction moment exerted on the wing at the root. This value is 675,977 newton-meters.
Integration of the distributed load with respect to the wingtip results in an equation for
the shear force acting at any point on the wing. The equation is as follows:
√
Integration of the load multiplied by the distance at which it acts, with respect to the
wingtip, results in an equation for the moment acting at any point on the wing:
√ √
Plots and graphical depictions of these functions can be seen below. Original data is
available in Appendix G: Structural Analysis Data.
Figure 19: Plot of shear force (N) and bending moment (N-m) throughout the wing
Figure 20: Shear force (left) and bending moment (right) distributions in the wing
*

29. 29
Our spar is symmetrical about its center, so the cross-sectional centroid is easily located.
The spar's moment of inertia is calculated by subtracting that of the small rectangle from that of
the large rectangle.
Compressive and tensile stresses are easily determined. One simply multiplies the
moment at the wing root by a vertical distance within the beam, y, and divides by the moment of
inertia. The result can be used to determine the stresses at the points of greatest tension and
compression (along the vertical axis, not longitudinally). Along the top surface of the wing, the
stress is | | . Along the bottom, the stress is | | .
The values are the same because the spar is symmetrical about the longitudinal axis. Needless to
say, these values are well within the limits of our material, even accounting for very large safety
factors.
Finally, deflection analysis can be conducted. Integrating the moment equation twice
yields an expression for deflection, too complicated to include here. The wing deflection looks
like this:
Figure 21: Wing deflection
The maximum deflection occurs at the wingtip, and has a value of 6.603 mm. This is
certainly reasonable. In non-dimensional form:
or 0.08%
Landing Gear
As with virtually every other business jet in existence, our aircraft will employ tricycle
landing gear. This permits ease of entry, excellent ground handling, good pilot visibility, and the
ability to land "crabbed" in adverse crosswind conditions. The nosewheel will have two tires, to
allow some degree of steering control should one be punctured.
Using Table 11.1 in the textbook as a reference, we have sized our aircraft tires as
follows:
Diameter Width
Rear tires 67 cm 19 cm
Front tires 60 cm 17 cm
Tires: 6 total Pressure: 120 psi
*

30. 30
Fuel Tanks
Based on a Wf/W0 of 0.3037 from our refined W0 estimate, our aircraft's fuel weight is
4416.8 kilograms. With a density of Jet A-1 fuel of approximately 800 kg/m3
, our fuel tanks
must have a volume of 5.521 m3
. Making some broad assumptions (like a uniform wing
thickness of 0.179 m, based on our airfoil characteristics), the tanks would need to have an area
(when seen from above) of 30.84 m2
. This is under our aircraft's wing area, which is around 34
m2
. As a result, our aircraft should be able to comply with FAA regulations for passenger-
carrying aircraft by storing all of its fuel in the wings.
The following is an approximate sketch of the fuel tanks' area and location. Of course,
real aircraft fuel tanks would be compartmentalized to permit pumping of fuel as ballast and to
maintain proper wing loading distributions.
Figure 22: Fuel tank location and size
Thrust-Drag Analysis
Drag is a function of two components that act at different speeds. Induced drag increases
in proportion to the square of the aircraft's velocity, while parasitic drag increases in proportion
to the inverse of the square of the aircraft's velocity. That is, induced drag goes to infinity as the
aircraft travels faster, while parasitic drag goes to zero:
Figure 23: Coefficients of drag
0
0.005
0.01
0.015
0.02
0.025
0.03
0 100 200 300
C
Do
Velocity (m/s)
0
0.0002
0.0004
0.0006
0.0008
0.001
0.0012
0.0014
0 100 200 300
C
D
Induced
Velocity (m/s)
*

31. 31
Induced drag is drag due to the lifting forces of the aircraft wing. According to one
explanation, lift is caused by a pressure differential between the top and bottom of the wing. At
the tips of the wing, however, air can move freely from the high pressure bottom to the low
pressure top of the wing. This movement induces drag through the creation of vortices (which
generate no lift). Two-dimensional induced drag coefficients can be calculated with the
following equation:
However, a better lift coefficient can be calculated for a 3D wing as follows:
√ ( )
,
where , AR = aspect ratio, M = Mach number, . The complete list of
calculations for induced drag are listed in Appendix H
Parasitic drag, as the name implies, is caused by friction between the air and components
that make up the aircraft's external structure. It can be approximated to any desired or required
degree using the "component build-up" method, which finds individual parasitic drags for each
part of the aircraft and sums them.
For our aircraft, we found drags for the following components: fuselage, tails (vertical
and horizontal), wings, winglets, and engines. We used the most recent version of our aircraft
drawings to ensure the most accurate dimensions for each component.
We used the following parameters, representing our aircraft's cruise conditions, in
determining its parasitic drag (V∞ and Mach number are variables).
Altitude Temp. ρ Vsound
45,000 ft 216 K 0.2371 kg/m3
295.06 m/s
The following are sample values taken at our cruise velocity of 485 kts:
Component Length Reynolds # Cf FF Swetted
Fuselage 18 m 74,703,141 0.00208 1.2 169.9 m2
Vertical tail 2.4 m 9,960,418 0.00281 1.449 18.0 m2
Horizontal tail 1.35 m 5,602,735 0.00309 1.530 18.0 m2
Wing 2.25 m 9,337,892 0.00284 1.458 54.0 m2
Winglets 0.6 m 2,490,104 0.00355 1.505 0.7 m2
Engines 2.7 m 11,205,471 0.00276 1.162 21.2 m2
To produce our final thrust-drag plot, we performed the above calculations for velocities from 10
m/s to 290 m/s in increments of 10 m/s.
Formulas used are as follows:
*

32. 32
Form factor calculations depend on the component being considered. Wings, tails and
winglets have a form factor determined by airfoil thickness and sweep, while the fuselage and
engine nacelles have a form factor determined by fineness ratios.
To calculate the total drag that of the plane the following equation is used.
A list of calculated values for total drag can be found in Appendix H. Our aircraft's cruise
speed can be confirmed with the following thrust-drag plot, indicating that cruise thrust and drag
intersect slightly above our specified cruise velocity of 250 m/s.
Figure 24: Thrust-drag plot
It should be noted that, as a result of thrust-drag analysis, we realized that our original
engine specification, the Pratt & Whitney 306A, was not powerful enough. The 306A was
therefore discarded in favor of the better-suited 308B.
Using maximum thrust instead of cruise thrust results in our aircraft's maximum speed:
524 kts (Mach 0.91). Our cruise speed is therefore 92.5% of our maximum speed.
Stability Analysis
The stability of an aircraft is a very important design consideration when conceptualizing
an aircraft. Pilot type becomes irrelevant if the plane cannot be made to fly in a stable,
predictable manner (whether by design or by complicated computer tricks). Complete stability
analysis typically considers pitch, roll and yaw moments and displacements within six degrees of
freedom. Due to time constraints, however, we will focus only on static pitch stability.
The first step of the stability analysis is to calculate the aircraft's neutral point, Xnp. This
location is significant because, if the center of gravity were to travel further aft, the aircraft
becomes unstable. As such, knowledge of its location is important to loading and weight
0
2000
4000
6000
8000
10000
12000
0 50 100 150 200 250 300
Thrust
&
Drag
Velocity (m/s)
Drag
Thrust

33. 33
distribution, factors that are determined by pilots and computer weight management systems
before and during flight. The value of Xnp can be calculated as follows:
̅̅̅̅̅
̅̅̅̅̅̅̅ ̅̅̅̅̅̅̅̅
A value of 10.15 m from the aircraft nose results for our plane's neutral point.
The next step is to calculate the aircraft's most forward point, Xmf. The most forward
point, like the neutral point, represents a plane beyond which the center of gravity cannot pass if
stability and proper control are to be maintained. A center of gravity forward of the most forward
point results in control sluggishness and potentially dangerous flight conditions. The following
equation was used to calculate the most forward point:
̅̅̅̅̅
̅̅̅̅̅̅̅ ̅̅̅̅̅̅̅̅
Our aircraft's most forward point is located 9.09 m behind the nose.
The final step is to calculate the aircraft's center of gravity. As long as the center of
gravity lies between the neutral and most forward points, the aircraft will be stable and will
respond properly to control inputs. Our aircraft's center of gravity was determined, like parasitic
drag, using a component method. A simple statics equation was used:
̅
∑
∑
Centers of gravity for the following components were used in the calculation:
Component Weight xCG
Wings 2467 kg 9.2 m
Horizontal tail 389 kg 16.7 m
Vertical tail 388 kg 15.8 m
Engines (both) 728 kg 13.5 m
Nose gear 100 kg 3.1 m
Main gear 407 kg 10.5 m
Avionics 225 kg 1.2 m
Fuel tanks (full) 4416 kg 10 m
Baggage 425 kg 13.8 m
Fuselage 3690 kg 7.8 m
Passengers 800 kg 10 m
Pilots 200 kg 3.6 m
Bathroom 300 kg 5.7 m
*

34. 34
The total weight for the above components is 14535 kg, very close to our W0 estimate of 14543
kg. This yields a CG location of 9.58 m. Details are in Appendix I: Center of Gravity.
The simplest assessment of stability can be conducted:
Does ̅̅̅̅̅ ̅̅̅̅̅ ̅̅̅̅̅?
Our values satisfy this relation, and our aircraft is therefore stable in its pitch axis. A graphical
representation of the above relation is presented here (to avoid cluttering up our drawings
elsewhere):
Figure 25: Stability diagram
The distance between xnp and xmf is 1.06 m, or 37.8% of our aircraft's mean aerodynamic chord.
The static margin,
̅
works out to 0.253, which is greater than zero, again indicating a stable design.
Maneuvers
Climb
We used the following equations to determine our aircraft's climb characteristics:
Climb angle: [ ]
Climb rate: [ ]
Our aircraft's resulting climb angle is 3.54˚. This corresponds to a climb rate of 15.42 m/s
or 3030 fpm. At this climb rate, our aircraft reaches its cruising altitude of 45,000 ft in
approximately 15 minutes, assuming a takeoff at sea level.
Research performed by NASA indicates that winglets, as used on our aircraft, can
dramatically improve rate of climb. Below 5,000 ft, winglets can raise ROC by 6%, and above
*

35. 35
that altitude the improvement increases to roughly 15% ("Flight Evaluation of the Effect of
Winglets…", Holmes et al., 1980). Theoretically, then, our aircraft could be capable of a climb
rate of up to 3485 fpm (above 5,000 ft), and thus could reach cruise altitude in the vicinity of 13
minutes, shaving two minutes off the non-winglet time. The corresponding improved climb
angle would be roughly 4˚.
Turn
We used the following equations to determine our aircraft’s turn characteristics:
Turn Rate: ̇ √
Turn Radius:
Our aircraft’s calculated turn rate is 0.1316 rad/s or 7.54 deg/s. This results in a turn
radius of 1899 m (6230 ft or 1.18 miles). At this turn rate, our aircraft can complete an 180o
turn
in roughly 24 seconds. This is a good value for our aircraft, providing a low response time if it
were necessary to turn the aircraft around and make an emergency landing. The load factor, n,
was set at 3.5 as in the structural analysis section.
An idea of the range of the turn radius values for our jet can be obtained from the
following graph.
Figure 26: Turn rate
Logo and Name
More for fun than anything else, we also named our aircraft company and model, as well
as creating simple logos for the same:
0
0.5
1
1.5
2
2.5
3
3.5
0 50 100 150 200 250 300 350
Turn
Rate
(rad/s)
Velocity (m/s)
*
*

36. 36
Figure 27: Logo
Vulcan was, of course, the Roman god of fire, metallurgy and technology. He produced
thunderbolts for Jupiter, the king of the Roman gods. Were Vulcan Aircraft to produce other,
perhaps larger jet models in the future, the naming theme lends itself to other logical (and
agreeable) names like "Jupiter" and "Lightning."
Conclusion and Summary
We present the Vulcan Thunderbolt, a new business jet capable of transporting six to
eight passengers on trans-Atlantic and inter-continental routes in complete comfort and luxury.
Featuring composite construction, a large fuel supply, and high safety margins, our aircraft is
well on its way through the design process.
No major pitfalls were encountered in our design process. Virtually every dimension,
weight and specification is in alignment with historical trends, although our aircraft does "push
the envelope" in a few areas like its composite construction. It checks out, structurally and in
maneuvers, and offers short-field capability and a rapid rate of climb.
As such, not much revision of existing work is necessary. Looking forward, however,
complete stability (in all three axes) would need to be conducted, the design would need to be
laid out and refined (detailed lofting), and scale-model testing and CFD analysis would need to
be conducted. Electrical, hydraulic, mechanical and aeronautical engineering expertise would be
required for these phases of design.
Because our aircraft is a business jet, comfort is of high priority to our customers. Our
interior has been laid out in a basic manner, but individual orders would have specific requests
for customization. Our large payload capacity would permit any number of interior appointments
to be installed.
Once design work ends, engineers perform cost analysis. Passing that, our aircraft would
be prototyped and would undergo rigorous testing by the FAA before receiving its flight
certification. If we were to go ahead with the design process immediately, we wouldn't expect
our aircraft to fly before sometime in 2010 or 2011. Reserve your Vulcan Thunderbolt today!

37. 37
Appendix A: Historical Comparison Data
Pax. Length (ft) Wingspan Height Empty weight Power Year
Hawker 800 8 51.1 54.3 18 15670 4660 1963
Hawker 1000 8 51.1 54.3 18 15670 4660 1962
Cessna S550 Citation 8 47.25 52.25 15 8060 2500 1978
Cessna Citation X 8 72.3 63.6 19 21700 6764 1996
Cessna Citation Excel 8 63.6 63.2 20.4 17700 5686 1996
Learjet 45 9 57.5 47.8 14 13695 3500 1995
Gulfstream G100 ? 55.6 54.6 18.1 14400 4250 1986
Hawker 400 7 48.4 43.5 13.9 10550 2965 1996
Sabreliner 6 44 44.5 16 9257 3000 1962
Learjet 35 8 48.6 39.5 12.25 10119 3500 1973
Learjet 28 8 47.5 43.8 12.23 ? 2944 1977
Learjet 40 6 55.5 47.8 14.1 ? 3500 2002
Learjet 55 7 55.1 42.75 14.6 12860 3750 1977
Learjet 23 6 35.6 43.25 12.25 6151 3000 1963
Gulfstream IV 15 88.3 77.9 24.4 35500 13850 1985
Gulfstream III 25 88.3 77.9 24.5 38000 11400 1979
Gulfstream G200 8 62.25 58 21.4 19200 6040 2000
Embraer Phenom 300 6 52 53.1 16.3 ? 3200 2008
Embraer Legacy 600 13 85.4 68.9 22.1 30000 8810 2000
Bombardier Global Express XR 12 99.4 94 24.9 49750 14750 1993
Bombardier Challenger 500 14 68.4 64.3 20.6 20485 9140 1978
Dassault Falcon 10 6 45.4 42.9 15.1 10760 3230 1970
McDonnell 119 10 66.5 57.6 23.6 41000 2980 1955
Eclipse 500 5 33 37.2 11 3550 900 2006
The following formulas, applying to business jets in general, can be derived:
 Wingspan = 25.16e0.012 (length)
 Wingspan = –7*10-9
(empty weight)2
+ 0.001 (empty weight) + 32.88
 Empty weight = 25896 ln (number of passengers) – 36903
 Engine power = 196.4 (length) – 6264
*

38. 38
Appendix B1: Initial Weight Estimate Iteration
Initial parameters
W0 Range Speed SFC – cruise SFC – loiter L/D
14322 kg 2500 nm 250 m/s 0.5 / hour 0.4 / hour 10.9
Wcrew Wpayload Loiter endurance Wingspan
200 kg 2000 kg 1 hour 17.1 m
First iteration
W0 – Initial W1 – Taxi, takeoff W2 – Climb W3 – Cruise
14322 kg 13892 kg 13683 kg 10807 kg
W4 – Loiter W5 – Landing W5 / W0 Wf / W0
10418 kg 10365 kg 0.7237 0.2928
We / W0 W0 – new
0.5744 16569 kg
Second iteration
We / W0 W0 – new
0.5694 15968 kg
Third iteration
We / W0 W0 – new
0.5707 16116 kg
Fourth iteration
We / W0 W0 – new
0.5704 16078 kg

39. 39
Appendix B2: Initial Weight Estimate Iteration
Initial parameters
W0 Range Speed SFC – cruise SFC – loiter L/D
16078 kg 2500 nm 250 m/s 0.5 / hour 0.4 / hour 10.34
Wcrew Wpayload Loiter endurance Wingspan
200 kg 2000 kg 1 hour 15.98 m
First iteration
W0 – Initial W1 – Taxi, takeoff W2 – Climb W3 – Cruise
16078 kg 15756 kg 14881 kg 12074 kg
W4 – Loiter W5 – Descent W6 – Landing W6 / W0 Wf / W0
11616 kg 11529 kg 11471 kg 0.71345 0.3037
We / W0 W0 – new
0.5539 15453 kg
Second iteration
We / W0 W0 – new
0.5539 15453 kg
Only two iterations were necessary to converge the result.
*

40. 40
Appendix C: Initial Weight Trade Studies
Range (nm) W0 (kg)
1500 12100
2000 14200
2500 16920
3000 20560
3500 25520
4000 32150
4500 38280
Weight = 0.002 (range)2
– 3.308 (range) + 12582
Loiter (hr) W0 (kg)
0.5 15819
0.75 16380
1 16980
1.25 17600
1.5 18280
1.75 18980
2 19740
Weight = 2609 (loiter) + 14421
Payload (kg) W0 (kg)
1000 9130
1500 12060
2000 14860
2500 17570
3000 20210
3500 22800
4000 25330
Weight = 5.387 (payload) + 3953
*

41. 41
Appendix D: Artwork
Dimensioned drawings:
*

42. 42

43. 43

44. 44
Side Concept art:
Front Concept art:
Top Concept art:

45. 45
Design evolution:
Preliminary: Interim: Final:
Color schemes:
Yellow Blue
"WPI" Green

46. 46
CAD drawing:

47. 47
Appendix E: Airfoil Geometry Data
Non-dimensional, normalized coordinates for a NACA 64008a airfoil:
X Yu X Yl
1 0.00018 0 0
0.95 0.00438 0.005 -0.00646
0.9 0.00858 0.0075 -0.00778
0.85 0.01278 0.0125 -0.00983
0.8 0.01698 0.025 -0.01353
0.75 0.02117 0.05 -0.01863
0.7 0.02521 0.075 -0.02245
0.65 0.02897 0.1 -0.02559
0.6 0.03234 0.15 -0.03047
0.55 0.03524 0.2 -0.03414
0.5 0.03757 0.25 -0.03681
0.45 0.03921 0.3 -0.03866
0.4 0.03998 0.35 -0.03972
0.35 0.03972 0.4 -0.03998
0.3 0.03866 0.45 -0.03921
0.25 0.03681 0.5 -0.03757
0.2 0.03414 0.55 -0.03524
0.15 0.03047 0.6 -0.03234
0.1 0.02559 0.65 -0.02897
0.075 0.02245 0.7 -0.02521
0.05 0.01863 0.75 -0.02117
0.025 0.01353 0.8 -0.01698
0.0125 0.00983 0.85 -0.01278
0.0075 0.00778 0.9 -0.00858
0.005 0.00646 0.95 -0.00438
0 0 1 -0.00018

48. 48
Appendix F: XFOIL Analysis
α CL CD CDp CM
-4 -0.5911 0.01868 0.01569 -0.0214
-3.9 -0.5788 0.01789 0.0148 -0.0206
-3.8 -0.5693 0.0167 0.01345 -0.0197
-3.7 -0.5585 0.01566 0.01225 -0.0187
-3.6 -0.5464 0.0147 0.01114 -0.0178
-3.5 -0.5334 0.01381 0.01013 -0.0169
-3.4 -0.5198 0.01301 0.0092 -0.016
-3.3 -0.5061 0.01237 0.00839 -0.0151
-3.2 -0.4914 0.01173 0.00763 -0.0143
-3.1 -0.4767 0.01118 0.00695 -0.0134
-3 -0.4623 0.01075 0.00635 -0.0126
-2.9 -0.4482 0.01038 0.00581 -0.0117
-2.8 -0.4413 0.01114 0.00557 -0.0104
-2.7 -0.4265 0.01073 0.00508 -0.0095
-2.4 -0.3825 0.00971 0.0039 -0.0071
-2.3 -0.3678 0.00944 0.00358 -0.0063
-2.2 -0.3529 0.0092 0.00328 -0.0056
-2.1 -0.3379 0.00899 0.00302 -0.0049
-2 -0.3226 0.0088 0.00278 -0.0043
-1.9 -0.3072 0.00865 0.00258 -0.0038
-1.8 -0.2915 0.00851 0.0024 -0.0033
-1.7 -0.2756 0.0084 0.00224 -0.0029
-1.6 -0.2596 0.00831 0.00211 -0.0025
-1.5 -0.2435 0.00823 0.002 -0.0022
-1.4 -0.2273 0.00817 0.0019 -0.0019
-1.3 -0.211 0.00812 0.00182 -0.0017
-1.2 -0.1948 0.00808 0.00175 -0.0015
-1.1 -0.1785 0.00805 0.0017 -0.0013
-1 -0.1622 0.00802 0.00165 -0.0011
-0.9 -0.146 0.008 0.00161 -0.0009
-0.8 -0.1297 0.00798 0.00157 -0.0008
-0.7 -0.1134 0.00796 0.00154 -0.0007
-0.6 -0.0972 0.00794 0.00152 -0.0006
-0.5 -0.081 0.00793 0.0015 -0.0005
-0.4 -0.0648 0.00791 0.00149 -0.0004
-0.3 -0.0486 0.0079 0.00147 -0.0003
-0.2 -0.0324 0.0079 0.00147 -0.0002
-0.1 -0.0162 0.00789 0.00146 -0.0001
0 0 0.00789 0.00146 0
0.1 0.0162 0.00789 0.00146 0.0001
0.2 0.0324 0.0079 0.00147 0.0002
0.3 0.0486 0.0079 0.00147 0.0003
0.4 0.0648 0.00791 0.00148 0.0004
0.5 0.081 0.00793 0.0015 0.0005
0.6 0.0972 0.00794 0.00152 0.0006
0.7 0.1134 0.00796 0.00154 0.0007
0.8 0.1297 0.00798 0.00157 0.0008
0.9 0.146 0.008 0.00161 0.0009
1 0.1623 0.00802 0.00165 0.0011
1.1 0.1785 0.00805 0.0017 0.0013
1.2 0.1948 0.00808 0.00175 0.0015
1.3 0.2111 0.00812 0.00182 0.0017
1.4 0.2274 0.00817 0.0019 0.0019
1.5 0.2436 0.00823 0.002 0.0022
1.6 0.2598 0.00831 0.00211 0.0025
1.7 0.2758 0.0084 0.00224 0.0029
1.8 0.2917 0.00851 0.0024 0.0033
1.9 0.3075 0.00865 0.00258 0.0037
2 0.323 0.00881 0.00279 0.0043
2.1 0.3383 0.00899 0.00302 0.0049
2.2 0.3534 0.0092 0.00328 0.0055
2.3 0.3683 0.00944 0.00358 0.0062
2.4 0.3831 0.00972 0.00391 0.007
2.5 0.3978 0.01002 0.00426 0.0077
2.6 0.4126 0.01037 0.00466 0.0085
2.7 0.4274 0.01074 0.00509 0.0094
2.9 0.4494 0.0104 0.00583 0.0115
3 0.4636 0.01076 0.00637 0.0123
3.1 0.4781 0.01121 0.00698 0.0132
3.2 0.4929 0.01176 0.00766 0.014
3.3 0.5076 0.0124 0.00842 0.0148
3.4 0.5214 0.01305 0.00925 0.0157
3.5 0.5351 0.01385 0.01017 0.0166
3.6 0.5481 0.01475 0.0112 0.0175
3.7 0.5602 0.01572 0.01232 0.0184
3.8 0.571 0.01677 0.01353 0.0193
3.9 0.5804 0.01798 0.0149 0.0202
4 0.5935 0.01868 0.0157 0.021
4.1 0.6096 0.01915 0.01622 0.0217
4.2 0.6211 0.02016 0.01736 0.0226
4.3 0.6299 0.02148 0.01883 0.0234
4.4 0.637 0.02294 0.02042 0.024
4.5 0.6428 0.02451 0.0221 0.0246
4.6 0.6477 0.02616 0.02385 0.0251
4.7 0.6515 0.02785 0.02565 0.0254
4.8 0.654 0.02955 0.02745 0.0256
4.9 0.6553 0.03126 0.02927 0.0256
5 0.6557 0.03301 0.03111 0.0255
5.1 0.6555 0.03479 0.03297 0.0253
5.2 0.6547 0.03663 0.03487 0.0248
5.3 0.6522 0.03839 0.03668 0.0244
5.4 0.6488 0.0402 0.03853 0.0238
5.5 0.6454 0.04217 0.04056 0.0225
5.6 0.6421 0.04426 0.04271 0.0207
5.7 0.6383 0.04636 0.04486 0.0189
5.8 0.6349 0.04862 0.04716 0.0167
5.9 0.6324 0.05105 0.04963 0.0141
6 0.632 0.05349 0.0521 0.0115
6.1 0.6324 0.05617 0.05479 0.0084
6.2 0.6324 0.05881 0.05745 0.0056
6.3 0.6322 0.06118 0.05986 0.0034
6.4 0.6314 0.06336 0.06209 0.0018
6.5 0.6307 0.06548 0.06424 0.0002
6.6 0.6299 0.06752 0.06632 -0.001
6.7 0.6287 0.06938 0.06829 -0.0019
6.8 0.6263 0.07104 0.07007 -0.0022
6.9 0.6274 0.07315 0.07224 -0.0039
7 0.6278 0.07519 0.07428 -0.0052
7.1 0.6284 0.07722 0.0763 -0.0064
7.2 0.6291 0.07922 0.0783 -0.0076
7.3 0.63 0.0812 0.08028 -0.0087
7.4 0.6309 0.08314 0.08222 -0.0098
7.5 0.6319 0.08504 0.08411 -0.0108
7.6 0.6332 0.0869 0.08598 -0.0117
7.7 0.6348 0.08873 0.0878 -0.0127
7.8 0.6365 0.0905 0.08957 -0.0135
7.9 0.638 0.09221 0.09128 -0.0143
8 0.639 0.09391 0.09298 -0.0149
*

49. Appendix G: Structural Analysis
z V M ω
0 -20309 103360.5 0
0.25 -19500.5 96949.47 1.37E-05
0.5 -18692.7 90657.66 5.35E-05
0.75 -17886.6 84501.9 0.000118
1 -17082.8 78497.73 0.000205
1.25 -16282.2 72659.4 0.000313
1.5 -15485.7 67000 0.000441
1.75 -14693.9 61531.43 0.000587
2 -13907.9 56264.5 0.00075
2.25 -13128.3 51208.95 0.000927
2.5 -12356.2 46373.44 0.001119
2.75 -11592.4 41765.66 0.001323
3 -10837.9 37392.24 0.001539
3.25 -10093.6 33258.87 0.001764
3.5 -9360.61 29370.2 0.001999
3.75 -8639.91 25729.95 0.002241
4 -7932.65 22340.81 0.00249
4.25 -7240.01 19204.49 0.002746
4.5 -6563.25 16321.69 0.003006
4.75 -5903.73 13692.05 0.003271
5 -5262.89 11314.15 0.00354
5.25 -4642.34 9185.437 0.003812
5.5 -4043.83 7302.174 0.004086
5.75 -3469.31 5659.363 0.004362
6 -2921 4250.638 0.00464
6.25 -2401.45 3068.129 0.004919
6.5 -1913.67 2102.279 0.005199
6.75 -1461.27 1341.585 0.005479
7 -1048.85 772.2251 0.00576
7.25 -682.492 377.486 0.006041
7.5 -370.997 136.7789 0.006322
7.75 -129.303 23.67364 0.006603

50. 50
Appendix H: Calculated Drag
V (m/s) Mach Cl afla Cd ind C_d_o Cd Drag Thrust
10 0.03389 0.082457 0.000351 0.0244 0.024743 16.28712 7476.00
20 0.067781 0.082464 0.000351 0.0224 0.022765 59.94061 7476.00
30 0.101671 0.082484 0.000351 0.0214 0.021761 128.9195 7476.00
40 0.135561 0.082523 0.000352 0.0208 0.021105 222.2743 7476.00
50 0.169452 0.082588 0.000352 0.0203 0.020621 339.3479 7476.00
60 0.203342 0.082685 0.000353 0.0199 0.020239 479.6047 7476.00
70 0.237232 0.082821 0.000354 0.0196 0.019924 642.6405 7476.00
80 0.271123 0.083004 0.000356 0.0193 0.019654 827.9743 7476.00
90 0.305013 0.08324 0.000358 0.0191 0.019416 1035.217 7476.00
100 0.338903 0.083537 0.000361 0.0188 0.019203 1263.996 7476.00
110 0.372793 0.083906 0.000364 0.0186 0.019009 1513.999 7476.00
120 0.406684 0.084355 0.000368 0.0185 0.018829 1784.713 7476.00
130 0.440574 0.084896 0.000372 0.0183 0.018661 2075.95 7476.00
140 0.474464 0.085541 0.000378 0.0181 0.018503 2387.185 7476.00
150 0.508355 0.086305 0.000385 0.018 0.018354 2718.292 7476.00
160 0.542245 0.087207 0.000393 0.0178 0.018211 3068.729 7476.00
170 0.576135 0.088266 0.000402 0.0177 0.018073 3438.17 7476.00
180 0.610026 0.08951 0.000414 0.0175 0.017943 3826.71 7476.00
190 0.643916 0.090971 0.000428 0.0174 0.017817 4233.679 7476.00
200 0.677806 0.09269 0.000444 0.0173 0.017696 4659.278 7476.00
210 0.711697 0.09472 0.000463 0.0171 0.01758 5103.371 7476.00
220 0.745587 0.097132 0.000487 0.017 0.017471 5566.222 7476.00
230 0.779477 0.100023 0.000517 0.0169 0.017369 6048.031 7476.00
240 0.813368 0.103529 0.000554 0.0167 0.017275 6549.689 7476.00
250 0.847258 0.10785 0.000601 0.0166 0.017192 7072.798 7476.00
260 0.881148 0.113295 0.000663 0.0165 0.017124 7619.774 7476.00
270 0.915038 0.120377 0.000749 0.0163 0.017082 8196.78 7476.00
280 0.948929 0.130023 0.000873 0.0162 0.017078 8813.531 7476.00
290 0.982819 0.14418 0.001074 0.0161 0.017151 9494.465 7476.00
295 0.999764 0.15428 0.00123 0.016 0.017243 9877.23 7476.00

51. 51
Appendix I: Center of Gravity
Component Weight (kg) xCG (m) Moment (N*kg)
Wings 2467 9.2 22696.4
Horizontal tail 389 16.7 6496.3
Vertical tail 388 15.8 6130.4
Engines (both) 728 13.5 9828
Nose gear 100 3.1 310
Main gear 407 10.5 4273.5
Avionics 225 1.2 270
Fuel tanks 4416 10 44160
Baggage 425 13.8 5865
Fuselage 3690 7.8 28782
People 800 10 8000
Pilots 200 3.6 720
Bathroom 300 5.7 1710
Total 14535 139241.6

52. 52
Appendix J: Final Presentation Slides
Business Jet
DUSTIN BRADWAY & KYLE MILLER
SPECIFICATIONS
• Two-crew small business jet (pax: 6 to 8)
• 485 kts cruise @ 45,000 feet
– 2 x PW308A
• 2,500 nm range, 1 hr loiter
• 2,000 kg payload including passengers
• Composite materials
• Quiet, efficient, comfortable
• Airborne at 135 kts, stalls at 120 kts
DRAWINGS DESIGN PARAMETER SUMMARY
• Payload: 2,200 kg (2 crew, 8 passengers and baggage)
• Stall: 120 kts; Cruise: 485 kts (M 0.84); VNE: 524 kts (M 0.91)
• W0: 14,543 kg; We: 8,559 kg
• L/D: 10.34; T/W: 0.0967
• Airfoil: NACA 64008a; AR: 7.46; Wingspan: 15.98 m
– ˚ dihedral; 8% thickness; ˚ incidence angle, % taper, ˚ sweep
• T-tail; tailspan: 4.64 m; tail AR: 4.0
• Wing loading: 82.09 lb/ft2
• Dimensions: 18.61 m long; 6.60 m tall; 2.35 m fuselage diameter
– Fuselage fineness ratio: 6.67
• Fuel tank volume: 5.52 m3 (4417 kg of Jet A-1)
STRUCTURAL ANALYSIS SUMMARY
• Box spar, carbon fiber
• Moment and shear diagrams
• Deflection
AERODYNAMIC ANALYSIS SUMMARY
Cruise Thrust from engine > Cruise Thrust Calculated
0
2000
4000
6000
8000
10000
12000
0 50 100 150 200 250 300
Thrust
&
Drag
Velocity (m/s)
Drag
Thrust

53. 53
CREATIVE DESIGN
• Concept art
• XFOIL
-0.8
-0.6
-0.4
-0.2
0
0.2
0.4
0.6
0.8
-4.5 0.5 5.5
Coefficient
of
lift
C
l
Angle of Attack
0
0.02
0.04
0.06
0.08
0.1
-4.5 0.5 5.5
Coefficient
of
Drag
C
d
Angle of Attack
CREATIVE DESIGN
• Historical data correlations
• Trade studies
0
10
20
30
40
50
60
70
80
90
100
0 20 40 60 80 100 120
Length (X) vs. Wingspan (Y)
0
2000
4000
6000
8000
10000
12000
14000
16000
0 20 40 60 80 100 120
Length (X) vs. Engine Power (Y)
FUTURE SOLUTIONS
• Complete Stability Analysis
• CFD code for fluid forces on Plane
CONCLUSIONS
• Plane is within Historical Data
• Plane is lighter than similar planes
– Composite Material
• Creative New Design