This document provides a final project report for the design of the Dragonfly Mk. I aircraft by TERA (Tyler, Elliot, Robert Aerospace). It summarizes the comparison to existing aircraft, the design requirements and specifications, and provides justifications for design elements like the wing loading, thrust-to-weight ratio, and T-tail configuration. The goal was to design a small personal jet capable of carrying 10 passengers for nearly 8,800 miles between Philadelphia and Bangkok, Thailand with minimized fuel consumption.

1. MEM 425
Aircraft Performance and Design
-- Summer 2014 --
FINAL PROJECT REPORT
Tyler, Elliot, Robert Aerospace (TERA)
Dragonfly Mk. I
Tyler Darrah (twd34@drexel.edu)
Elliot Farquhar (ef93@drexel.edu)
Robert Stricek (ras376@drexel.edu)
Submission Date: 9/4/2014
Section Number: 001
Team: I
Professor: Ajmal Yousuff

2. Comparison to Existing Aircraft
There are various different aircraft available with particular mission goals, specifications,
luxuries, etc. tailored to design requirements requested by a stakeholder. Our mission objective
had been to design an aircraft to take a family of 10 from Philadelphia International Airport
(PHL) to Suvarnabhumi Airport (BKK) in Bangkok Thailand as shown in Figure 1 below,
consuming a minimum amount of fuel for a non-stop flight. The total distance of the trip had
been approximated to be 8,721 miles [1], an extreme distance in the realm of passenger aircraft.
This kind of mission most often involves multiple stops, largely attributed to refueling purposes
and is therefore an inconvenience to the customer. These delayed trips range between 20 and 26
hours, specifically from PHL to BKK [2]. There exist planes that can handle such distances,
otherwise known as ultra-long-haul trips, like the Boeing 747 400ER in the 747 family. This
aircraft has range of 7,670 nautical miles, 8,826.48 miles, and cruises at 0.855 Mach, nearly the
speed of sound [3]. Though the range satisfies our customer needs, the problem with using this
aircraft for our mission is that all ultra-long-haul aircraft are extremely large and carry great
quantities of passengers. For example, the Boeing 747 400ER can carry at a minimum 416
passengers, and a maximum of 524 passengers [3]. The reason behind this is that such great
distances cost aircraft operators significant amounts of money, mostly attributed to fuel and
flight duration. Cost is always a large concern, but the more time an aircraft is in the air for a
single mission renders it unable to be used for other missions, or multiple flights on a time
interval basis. The solution to these previous two problems is to put as many passengers as safely
possible on the plane to make up differences in operational costs. In summary, there is no small
personal jet class aircraft that can carry a maximum of 10 passengers (2 pilots and 1 crew) for
nearly 8,800 miles, let alone minimize fuel costs. Our aircraft has been designed to adhere to the
aforementioned mission objectives as well as other, more specific, customer needs and
specifications, while dually minimizing fuel consumption. Not only has TERA developed a new
aircraft, but we are bringing to life a new classification of aircraft as well, dubbed ultra-long-
range personal jet.

3. Figure 1. Distance, PHL to BKK airport [1]
The aircraft had been designed with various stakeholder needs and specifications in
consideration, organized in Table 1 below. These specifications, along with fuel consumption
minimization, were the main driving force for various initial estimates including weight, wing-
loading, thrust to weight, take-off distance, and other elements.
Requirements Specifications
Passengers 10 Family Members, 2pilots, 1 attendant
Cruise Velocity 0.75 Mach
Cruise Altitude 40,000 ft.
Stall Velocity 90 miles/hr.
Range Philadelphia to Bangkok, Thailand, non-stop
Loiter 30 minutes at 30,000 ft. @ 0.6 Mach or less
Rate of Climb 1,500 ft/ min (1,000 ft/ min minimum)
Power Plant more than 1; high-bypass turbofan (existent)
Structure Composite Material
Max Load Limit Factor 4
Take-Off Distance 6000 ft. w/ 50 ft. clearance
Landing Distance 4000 ft. w/ 50 ft. clearance
Table 1. Stakeholder needs

4. Performance Justifications
To start the design process, the team chose a plane which exhibited a lift to drag ratio of 20 or greater.
Using Figure X, below, it was determined that the B-52 bomber fit the desired constraint the best with a
lift to drag ratio of 21. The B-52 used the root airfoil NACA63219.3 and the tip airfoil NACA 65A209.5,
according to the University of Illinois. [x] These specific airfoils, however, cannot be found since they are
modifications of standardized configurations. Thus, the TERA team decided to use the NACA63210 and
the NACA65210 airfoils which should exhibit similar lift to drag characteristics.
Figure X: L/D versus wetted aspect ratio [1]
Using coefficient of lift versus angle of attack plots for these airfoils, seen in Figures X and X
below, the max coefficient of lift was determined to be 1.6.
Figure X: Cl versus angle of attack [2]

5. Though the specific fuel consumption of our engine(s) were not known at the start of this
project, a rough estimate was taken using the statistical data, provided by Raymer, in Table 3.3 (Raymer,
19). For a high-bypass turbofan the cruise sfc value was .5 (lbs/s) and for loiter it was .4 (lbs/s). Assuming
that the flight path comprised of 5 stages, shown in Figure X below, the following weight ratios between
stages were used.
Figure X: Flight path
Table 1: Weight ratios [3]
Unlike Anderson, the TERA group decided to calculate the weight ratio of the loiter phase, stage
3 to 4, since there was a significant amount of time the plane was to be in this transition. For the
cruising phase of the flight, stage 2 to 3, the L/D ratio was not equal to 21 but instead was calculated to
be .866 percent of this maximum value since jets cruise more efficiently at higher velocities.
In order to calculate the empty weight, some research was performed in order to choose an
appropriate empty weight to takeoff weight ratio. According to Figure 3.1 in Raymer, shown below, the
empty to takeoff weight ratio should be about 5.4. Since one of the key design parameters is to use the
least amount of fuel possible, other alternatives were sought out to lower this ratio. Currently, there is
only one plane on the market which utilizes composite material to a large extent and that would be the
Boeing 787. The fact that fifty percent of the airframe is comprised of these materials, according to
Stages Ratio
W1/W0 0.97
W2/W1 0.985
W5/W4 0.995

6. Boeing, is the reason why the airframe weight dropped 20 percent from previous models (Boeing, 18).
The problem with composite materials, however, is that they are brittle in comparison to metals. This is
due to a variety of factors which include, but are not limited to, delamination, manufacturing errors due
to the precision required, and micro cracks. Based on our preliminary investigations our team chose to
reduce the empty weight of the aircraft by only 16.7, still using metals for key components. With this
said the empty to takeoff weight ratio came out to be .45.
Figure X: Empty weight trends [1]
The next step was then to calculate the wing loading ratio. For stage 5, landing, the landing
velocity multiplier (j) and the coefficient of friction (μr) values were considered to be 1.15 and .4 which
are standard commercial airlines. The amount of time before the brakes are applied after the plane has
touched down (N) was chosen to be 3 seconds in order to account for human reaction time. For the
cruising condition, stages 2 to 3, the team decided to make the parasitic drag (Cdo) to equal .12 since this
is approximately the same as the B-52 and because the wings and body will be relatively thin and
smooth. Though the Oswald efficiency factor for most jet aircraft are around .8, the dragonfly Mk I has
an efficiency factor of .6, according to McCormick, due to its low wing design.
The Dragonfly Mk. I was designed with only one engine since the thrust needed could be
provided by only one engine and a second engine would just increase the weight of the aircraft. The
reliability of modern jet engines has also improved to the point where long distance flight can be
guaranteed by the used of one engine.

7. Performance
Wing Loading
Wing loading is the weight of an airplane per unit wing area, i.e. lb/ft2
. This parameter can be
calculated various ways depending on the segment of the mission the airplane is in. One is limited by the
takeoff or stall velocity where:
Eq. (1)
This produces the value of 46.592 lb/ft2
. Another method of calculating the wing loading is using
conditions that an airplane undergoes when landing which is a function of air density at sea level,
maximum coefficient of lift, stall velocity, etc. This gives a wing loading value of 115.6891 lb/ft2
.
The final condition from which wing loading can be calculated from is at cruise. The equation
that dictates this is:
√
Eq. (2)
This gives the wing loading as 38.9767 lb/ft2
. The valued used for the Dragonfly Mk. I was the wing
loading value at cruise since it was the lowest value and therefore is the limiting factor.
Thrust to Weight Ratio
Like wing loading, the thrust to weight ratio needs to be chosen from different values depending
on the segment of the mission the airplane is in. The thrust to weigh ratio calculated from takeoff which
is a function of the stall velocity and air density at sea level. This gives a thrust to weight ratio of 0.0578.
The thrust to weight ratio is different when climbing and is dependent on the rate of climb as
shown in the equation below:
Eq. (3)
This produces the thrust to weight ratio of 0.3143. Another value can be inferred from data collected
from previous examples of airplanes with the use of tabulated constants:
, a=0.267, c=0.363 Eq. (4)
This thrust matching value comes out to be 0.2405. A tabulated value for the trust to weight ratio of
previous transportation jet airplanes is 0.25.
The final source for thrust to weight ratio comes from sustained turn requirements:
Eq. (5)

8. The thrust to weight required for the sustained turn is then 0.3016. The values for the thrust to
weight ratios of the sustained turn and cruise were checked in order to confirm their validity. The values
calculated above were greater than the check values meaning they were valid for consideration. The
thrust to weight ratio that is used for the Dragonfly Mk. I is the largest value calculated above which is
the thrust to weight ratio required by the rate of climb.
Rate of Climb
The maximum rate of climb at sea level for our air craft is calculated as a function of the
maximum lift to drag ratio, the thrust to weight ratio at climb, and varies the density of air. The
maximum climb rate at sea level is 97.7624 ft/s. The relationship between altitude and maximum rate
of climb can be seen in Figure [X] below.
Figure [X]. Maximum Rate of Climb vs. Altitude
Ceilings
Based on the rate of climb and air density, both the absolute and service ceiling could be
calculated where the absolute ceiling is where the airplane can maintain horizontal flight, which was
calculated to be 48,474.83 ft. The service ceiling was calculated to be 47,341.92 ft, which is where the
airplane’s rate of climb is 100ft/min [X].

9. Design Justifications
The design process began by selecting a fuselage design. In order to mimic similar business jets a
double-bubble fuselage shape was chosen to have the same dimensions of the Bombardier Global 6000.
This shape can be seen in, Figure X, below.
Figure X: Double bubble fuselage shape of a 717 [5]
After determining that the total amount of fuel needed for the trip would fit in the wings, a
decision had to be made on where to put the extra 113 cubic feet. If this amount of fuel were put in the
back, it would severely shift the center of gravity toward the back and make the plane very unstable
during climb. On the other hand, the fuel could not be put underneath the passengers in order to ensure
their safety in case of an emergency. With approximately 4,200 lbs of fuel calculated to be stored in
each wing, the addition of 2,500 lbs external tanks is highly impractical. Considering the yield strength of
Aluminum 2024-T4, a commonly used alloy in the aerospace industry, to be 5,760,000 psf there would
not be enough fuel to cause failure. The moment on the shoulders, however, is of more concern. As the
wing stands, it would impart 3,347,232 psf shear stress on the shoulder. This force is just 972,768 psf
less than the shear strength of Aluminum 2024-T4. [6] Though adding the 2,500 lbs would not increase
this shear stress enough to cause aluminum failure, there is only a factor of safety of about 1.3. Also,
when lifting off the ground this stress would increase due to the initial lift off, acceleration in the vertical
direction. Lastly, by adding these external fuel tanks onto the wings our designated drag coefficients
would have to increase in order to accommodate these disturbances in laminar flow. Thus it was
decided that the excess fuel would have to be placed somewhere inside the fuselage. Instead of placing

10. the fuel beneath passengers and crew, the team decided that it would be best to place it on either side
of the fuselage, providing ample room for the aisle. These compartments would also have to be toward
the front of the plane in order to keep the center of gravity from shifting too far backward.
Next, the design shifted to engine placement. Since only one engine was required to produce
9,500 lbs of thrust, the only logical choice was to position the engine in the rear of the plane to avoid
turbulence across either the wings or tail. To keep the drag coefficients at a minimum, the engine was
placed inside the aircraft and fed air by two, horizontally placed, scoops. The reason why the scoops
were placed horizontally was to prevent turbulent flow over the vertical tail stabilizer.
With the scoops positioned on either side of the fuselage, the horizontal tail stabilizers could not
be placed in a conventional tail configuration. In order to prevent the effect of downwash, the cruciform
tail configuration was also disregarded which left a T-tail design. An example of how wing downwash
affects the horizontal tail stabilizers can be seen, in Figure X, below.
Figure X: Downwash [7]
The T-tail design has other benefits, such as a higher longevity due to the reduction of vibrations
imparted by the engine. Since the wings do not play a role on the airflow across the tail, the horizontal
tail area can be made smaller. Lastly, since the horizontal tail has what is known as an end-plate affect,
the vertical tail area can also be reduced. The disadvantages of said design are an overall heavier tail
structure since the main vertical spar has to take the bending moment imparted by the horizontal tail
and a condition known as deep stall. Essentially as the angle of attack increases the wake and downwash
of the fuselage and wings begins to create turbulent flow over the horizontal tail. In said state the

11. aircraft may suddenly pitch to higher angles of attack. At some point, however, the tail will exit the
turbulent flow and give the pilot control over the aircraft once again. [7] This condition can be seen in
Figure X, below. The problem lies with trying to pitch the aircraft into a lower angle of attack since it will
re-enter the turbulent stream from the fuselage and wings. This, however, can simply be mitigated by
appropriate warning methods. Once such warning method is to properly designate the tapper ratios of
both the wings and the tail, this will create an initial stall condition toward the shoulder joints. The user
should feel this and still have control over the ailerons and elevators which are located toward the ends
of the wings. The smaller the ratio the further the separation of flow occurs from the shoulder. For
standard airplanes this ratio ranges between .4 and .6 and thus chose the greatest value of .6.
Figure X: Deep stall condition [7]
As for the physical dimensions of the T-tail they were determined using ratios purposed by a
Stanford professor. For example, the vertical and horizontal tail volumes were chosen to be .0425 and .5
based on the following Figures.
Figure X: Statistical approach to vertical tail volume [8]

12. Figure X: Statistical approach to horizontal tail volume [8]
According to this source, it was also reasonable to choose aspect ratios for the vertical and
horizontal tail to be 4 and 1.65. The taper ratio of the vertical tail was determined to be .85 in order to
insure enough area for the vertical section to attach onto. [8]
The main justification of our team’s decision to use a low wing design is due to the effect it has
on the stability of the aircraft. With the wings lower there would be less down wash on the T-tail system
in both normal operating angles and during deep stall. As stated before, with less air disruption the tail
can be made smaller. In terms of longitudinal stability, the wing drag line is positioned lower than the
center of gravity of the aircraft and thus creates a nose down pitching moment. This means that after a
sudden increase in angle of attack the plane will return back to its original state. This configuration also
increases static lateral stability due to the fuselages contribution to the wing dihedral effect. However, if
the wing is angle too far up the amount of lift decreases and the effect of cross winds become greater.
For this project a small angle of 5 was chosen in accordance with other aircraft of this kind. Other
benefits include shorter landing gear, better take off performance due to the ground effect, better
maneuverability, and less induced drag. Some disadvantages are less lift production, turbulent flow
separation located at the connection between the body and the wing, higher stall speed, less laterally
dynamically stable, and a longer landing distance. Even with these disadvantages the low wing design
was chosen since the large Cl/Cd ratio and simple fillets would counteract these effects. [9]

13. The flaps on the wing were designed using the standard that the width of the flap be 30% of the
wings chord length at any given outward location.
With these issues taken care of the passenger, baggage, crew, and amenities placements were
left. Essentially these positions were based on their effect to the center of gravity. With more than 18
hours of flight, a bathroom and a small kitchen seemed mandatory. These were placed toward the front
of the plane due to their significant weight. The passengers were placed toward the back not only
because their weight would not affect the center of gravity as much as other components but also
because our team wanted to make sure they were as far away from the auxiliary fuel tanks.
Once the wing configuration and weight distribution were decided upon the landing gear could
be chosen. The landing gear configuration was decided to be a tricycle configuration since the center of
gravity of the airplane would be in front of the wings which were of a low wing design, further
supporting the decision to make the configuration a tricycle.
Dragonfly Mark I Design
The first step, in the design process, was to determine the aerodynamic chord of the wings. This
was achieved by using Equations X and X below. For this particular aircraft the chord location and chord
length are 18.3169 ft and 9.8 ft. These equations can also be used for both the horizontal tail and
vertical tail.
̅ ( ) (
( )
)
̅ ( ) ( )
Though Anderson does not provide a detailed method of determining the fuel which can be
stored inside the airfoil, a method was devised using the thickness of the airfoils along the chord
lengths. With fuel cell thicknesses of .7ft and .6 ft. only a certain amount of the wing could hold said fuel

14. cells while maintaining a minimum clearance of 6 inches between the fuel cell and the outer airfoil. The
resulting integration can be seen with the other finalized drawings.
The center of gravity is essentially a sum of all the moments about the nose divided by the total
weight of the objects which create a moment affect. In order to simplify the equation each objects
center of gravity was determined in order to treat them as a point mass in relation to the front of the
aircraft. The distances and weights of the considered objects can be found, in Table X, below. The center
of gravity is then positioned 50.1750 ft from the front of the aircraft as dictated by Equation X.
Table X: Center of gravity relations
( ) ( )
Component Distance X (ft) Weight W (lbs)
Engine 80.8115 2301.6
Flight Attendant 24.222 180
Bathroom 28.972 130
Kitchen Area 28.972 400
Passengers 56.157 2000
Auxillary Fuel 36.797 4767.5
Crew Baggage 36.797 60
Electrical Systems/Fuel Pumps/Miscelaneous 43.297 200
Pilots 20.472 360
Passenger Baggage 70.342 500

15. Since this center of gravity calculation does not include the wings, it is safe to assume that this
first estimate will change. To place the wing, the mean aerodynamic center is lined up with the first
estimate of the center of gravity. The aerodynamic center and the center of gravity of the wing were
calculated to be 2.45 ft and 3.92 ft from the leading edge of the wing along the mean aerodynamic
chord. Finally, using Equation X, the new center of gravity location of 50.8889 ft from the front was
calculated by also considering the wing as a point mass.
( ) ( )
( ) ( )
( )
Dividing the distances to the tail mean aerodynamic centers, from the front of Anderson’s
aircraft, by the distances to the tail center of gravity locations, ratio were determined to find the
aerodynamic center locations of our aircraft. This can be seen below.
( )
( )
The aerodynamic center distances from the planes center of gravity then become 35.9837 ft and
33.61 ft. Using Equations X and X, the plane view areas were determined. Other parameters such as
wing span, root chord length, and tip chord length are calculated the same as the main wing.

16. √( )
In order to properly place the landing gear, a final adjustment must be made to position the
wings in relation to the center of gravity of the aircraft. Then the positions of the rear wheels are a
matter of the new center of aerodynamic center location of the wings. This is done by using the
following equations.
( )
( ( ) )
A ratio again was calculated between the distance of the front wheel and the length of the
fuselage of Anderson’s example aircraft to determine the position of the front wheel. Thus, for our
plane, the front wheel should be positioned 7.8185 ft from the nose as shown below.
( )
Using the force diagram in Anderson’s book, the load distribution among the tires was
determined to equal 1883.2 lbs on the front wheel and 13,341.5 lbs on each back wheel. Andersons
diagram is given below.

17. Figure X: Wheel load distribution [3]
Determining the wheel diameter is the last thing which needs to be done. This is achieved by
using the A and B values, provided by Anderson, to be used in Equations X and X. For our aircraft the
wheel diameters and widths came out to be 3.5081 ft and 1.1673 ft for the rear wheels and 1.7488 ft
and .6265 ft for the front wheel.
The last aspect of the design involves determining a better weight estimate. This is done by
using the equations listed, in Table X, below. It is important to note that all area which are used are the
exposed areas only, thus any part of the wings which are considered part of the body are not
considered. Roughly estimating the airplane as a combination of a cone, a cylinder, and an eclipse, the
general wetted surface area was calculated as shown below.
Table X: Weight estimations

18. ( √ )
( ) ( )
The new takeoff weight then becomes the sum of the weight of the crew, the payload, the
previously estimated fuel, and the new empty weight which was calculated above. From here the
processes was iterated multiple times, plugging in each new consecutive takeoff weight into the landing
gear and all else empty equations to determine an empty weight. The Fuel weight then becomes the
multiplication of each consecutive new takeoff weight by the originally determine Wf/Wo ratio of .4451.
The final values of We, Wf, and, Wo, as well the values for each iteration, can be found in Table X below.
Table X: Convergence of weight characteristics
Component Equation Weight (lbs)
Wing Weight 2.5*(Wing Area) 864
Horz. Tail Weight 2*(Horz. Tail Area) 113.5348
Vert. Tail Weight 2*(Horz. Tail Area) 163.0125
Fuselage Weight 1.4*(Wetted Area) 6246.4
Landing Gear Weight .057*(Wo) 1685.3
Installed Engine Weight 1.4*(Engine Weight) 2301.6
All Else Empty .1*(Wo) 2956.6
Empty Weight ∑ of components 14330

19. After putting all the information together, the Dragonfly Mk I CAD designs were finalized and
are shown below.
Figure X: Cross section right view of main fuselage given room lengths and rudder details
Iteration We (lbs) Wf (lbs) Wo (lbs)
1 14330 13161 30592
2 14491 13618 31209
3 14588 13893 31581
4 14647 14058 31805
5 14682 14158 31940
6 14703 14218 32021
7 14716 14254 32070
8 14724 14724 32100
9 14728 14289 32117
10 14731 14297 32128
11 14733 14302 32134
12 14734 14305 32138
13 14734 14306 32141
14 14735 14307 32142
15 14735 14308 32143
16 14735 14308 32143

20. Figure X: Cross section right view of main fuselage given seating, engine, and fuel comp. dimensions
Figure X: Right view of exterior given engine scope and rudder Cg dimensions
Figure X: Cross section front view given fuselage dimensions

21. Figure X: Cross section front view given auxiliary fuel cell dimensions
Figure X: Top view of exterior given wing and horz. tail span, plane length, and main wing chord lengths

22. Figure X: Top view of exterior given scoop layout, aerodynamic center, and chord lengths of horz. tail
Figure X: Top view of exterior horz. tail given sweep angles and elevator dimensions

23. Figure X: Cross section top view of wing given fuel cell dimensions
Figure X: Cross section top view of wing given spar locations
Figure X: Cross section top view of wing given aileron dimensions

24. Appendix
Calculations
Reference
[1] - https://www.google.com/?gws_rd=ssl#q=PHL+to+BKK+distance&safe=off
[2] -
https://www.google.com/flights/#search;f=PHL,ZFV;t=BKK;q=phl+to+bkk+nonstop;d=2014-
09-19;r=2014-09-23 14-09-23
[3] - http://www.boeing.com/boeing/commercial/747family/pf/pf_400er_prod.page
[X] - Dr. I. Halliwell, “An Improved Engine for a High Altitude Long Endurance Unmanned Air
Vehicle”, Joint AIAA Foundation / ASME, 2012, pp. 1-35.
Robs sources
[1] Raymer
[2] http://www.dept.aoe.vt.edu/~mason/Mason_f/B52S05.pdf
[3] Text Book
[4] http://www.boeing.com/commercial/aeromagazine/articles/qtr_4_06/AERO_Q406_article4.pdf
[5] http://adg.stanford.edu/aa241/fuselayout/images/717coachx.gif
[6] http://asm.matweb.com/search/SpecificMaterial.asp?bassnum=MA6061t6
[7] http://faculty.dwc.edu/sadraey/Chapter%206.%20Tail%20Design.pdf
[8] http://adg.stanford.edu/aa241/stability/taildesign.html
[9] http://faculty.dwc.edu/sadraey/Chapter%205.%20Wing%20Design.pdf