Blended Wing Body Aircraft

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iter8 Manta
AIAA Design Project Spring 2020
University of Southern California
Department of Aerospace and Mechanical Engineering

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Abstract – The iter8 Manta is a top-of-the-line Blended Wing Body (BWB) aircraft intended
to absolve increasing congestion in major international airports by carrying a high passenger
capacity for short range trips. The described market deficit requires a unique and unopposed
aircraft that is competent under visual- and instrumental-flight-rules-autopilot with a range
of 3500nmi (6482km) but which is optimized for 700nmi (1296.4km) trips carrying 350
passengers in economy class and 50 in business class. The mission design objectives also call
for a balanced field length of no more than 9,000ft (2743.2 m) on dry asphalt or concrete
pavement as found on a typical airport runway. Manta achieves these requirements as
determined via iterative structural and performance analysis based on a folding-wing BWB
design with supercritical airfoils composing its profile to mitigate drag and promote
efficiency. A double-decker mini-behemoth equipped with two rear-vertical-stabilizer-
mounted high bypass GEnx engines and revolutionary folding wings, Manta directly
addresses the mission requirements with high fuel efficiency, quick trip times, short to
medium range, dramatically low operating costs yielding passenger affordability, and
parkability in nearly any domestic gate with a folded span of just 45m. Manta has a service
ceiling of 17.5km but cruises economically at 12.5km at the Carson cruise velocity for
maximum range, 220
𝒎
𝒔
= M0.75, leading to a maximum range of 8,162km at cruising
altitude, a 4h30m flight. With a sleekcarbon-fiber composite fuselage nacelle, an aspect ratio
of 5.38, a maximum lift-to-drag ratio of 20.91, maximum takeoff weight of 1.227E6 N, and
an impressively low specific fuel consumption of 0.47 N/N-hr in takeoff and 0.4 N/N-hr in
cruise, Manta has the ability to carry up to 400 island hoppers or transatlantic travelers in
what is potentially one of the most sustainably, affordably, and comfortably optimized planes
on the 2029 aviation market.

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THE TEAM
Table 1. The team members and engineers of Iter8 are all listed below with their respective roles
and responsibilities.
Name Role Responsibilities
Natalie
Warren
Project Manager
Team logistics & organization, report literature and layout, trade
studies and market research, materials, instrumentation, and
avionics selection, assisting sub teams with various research and
computational tasks
Steven
Douglass
Project Manager
Organization and flow of project, set goals and deadlines,
organize/run meetings, assist all other groups as needed,
research materials, run cost analysis on each component,
prioritize, organize and review report holistically
Damien
Xie
Aerodynamicist
Conduct graphical and computational analysis on various
aircraft structures, airfoils selections and simulations, adjust
design parameters to achieve optimization, generate relevant
plots and diagrams
Kevin
Moran
Aerodynamicist
Research various airfoil designs and modern commercial wing
design principles that can be incorporated into the design
project. Perform trade studies on wing configurations during
takeoff, inflight, and landing operations, ensuring reasonable
paraments can be met. Generate figures, graphs, and tables
illustrating expected performance. Additionally, will be
researching tail design requirements and innovative fuselage
designs.

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Hayden
Shively
Propulsion
Research existing jet engines. Decide (1) which one to use and
(2) how many to use, based on required aerodynamic
performance parameters, CFR rules, and weight/balance.
Analyze where to put selected engines. Provide input on
instrumentation based on my experience as a pilot. Help with
Github and coding.
Nick
Ilibasic
Propulsion
Perform trade studies on preexisting turbine engines and decide
which powerplant to utilize in our design. Assure plane can be
approved for flight according to CFR 25 requirements.
Determine which instrumentation devices to incorporate in our
design.
Diana
Salcedo-
Pierce
Weight &
Configuration
Read and study chapters on weight analysis, weight estimation,
create aircraft CAD models, keep track of weight and CG for
various components
Seant
Minassian
Weight &
Configuration
Explicitly read through the posted material on Blackboard
regarding Weight Layout, create a CAD Model of the aircraft,
and use CAD tools to construct an FEA analysis on the model.
SCHEDULING
In order to ensure that deadlines were met and to give structure to each week of work, a GANTT
chart was created for organization. It was followed and was successful in keeping the team on task
throughout the project. Some modifications were made towards the end of the project due to
obstacles faced in the design process. A copy of the GANTT chart and roles key can be seen in
Figure 1.

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INTRODUCTION
The growth of commercial flight has shown light on issues regarding congestion at popular
airports, forcing passengers to connect and travel to satellite airports to reach their desired
destination. Travel between major airports that are closer together is still essential and requires a
lot of mid to short range aircraft to support the demand for these popular flights. The problem with
the current aircraft on the market is that they are too small to carry the amount of passengers that
are travelling between the desired airports and, the ones large enough to complete this mission, are
designated for long range travel and do not complete the shorter, more popular flights. This leads
to delays and an increased cost of the smaller flights for the consumer due to a large demand. The
request for proposal aims to solve these problems by designing an aircraft that is a high capacity,
short range transport aircraft designed to alleviate airport congestion, without the size and cost that
comes with long range capability. Specifically, the aircraft needs to be designed for a 6482 km
range that is able to carry up to 400 passengers in a dual class set up with a target entry into service
of 2029.
The general design objectives are to design an aircraft that is optimized for a 1296 km voyage,
must be as or more reliable than comparative aircraft, must have aircraft maintenance that is equal
to or better than comparable aircraft, and must minimize the production cost by choosing materials
and manufacturing methods appropriate for the annual production rate that is supported by the
Figure 1. The GANTT chart that led the team’s design process. Organized by weeks, the engineers in
charge for each task were responsible for presenting material to the rest of the team every suggested 1st
due date.

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team’s assessment of the potential market size. In addition to these requirements, there are
additional mandatory and tradeable requirements as seen in Figure 2.
TAKEOFF & LANDING INFLIGHT
CONFIGURATION &
PAYLOAD
PROPULSION
Capable of taking off and
landing from runways
(asphalt or concrete)
Capable of VFR (Visual
Flight Rules) and IFR
(Instrument Flight Rules)
flight with an autopilot
Passenger/pilot/attendant
weight of 90.7185 kg,
Baggage/occupant, 13.6078
kg
Engine/propulsion system
assumptions documented
Maximum takeoff length of
2743.2 m over 10.668 m
obstacle from a runway
with dry pavement (Sea
Level, Std. Atmospheric
Conditions) at Maximum
Takeoff Weight
Meets applicable
certification rules in
FAA 14 CFR Part 25 [3]
400 passengers in a dual
class configuration
Use of engine(s) that will
be in service by 2029
TO and Landing distances
should be equal and meet
CFR requirements
Cabin pressurized to
2438.4 m pressure
altitude at maximum
flight altitude
Crew: 2 pilots, 8 flight
attendants
Assumptions on specific
fuel
consumption/efficiency,
thrust/power and weight
should be given
Maximum Approach Speed
of 79.2244 m/s at the end of
the design range mission
50 passengers in Business
class with 0.9144 m pitch,
0.5334 m width
6482km design range
mission
350 passengers in Economy
with 0.8128 m pitch, 0.4572
m width
Flight to alternate airport
370.4 km from destination
airport
0.141584 𝑚3 per passenger
for baggage
A 30 minute hold at the
alternate airport
Galleys, Lavatories, and
Exits meet CFR
5% contingency fuel,
defined as 5% of non-
reserve block fuel
Figure 2. This figure displays mandatory requirements in red, tradeable requirements in blue,
and general information/requirements in green.
When looking at solutions that already are on the market, the range matches that of a Boeing 737
but requires a passenger capacity of a Boeing 777. With this in mind, one of the additional
requirements is that the designed aircraft must be able to fit into gates and terminals that the 737
can fit in. This increases the number of airports that the design will be able to travel to and ensure
that no special gate/jetway types are needed for the aircraft.
The current competitors are based on range of aircraft which include the Boeing 737and Airbus
A320. Their ranges are about 5600 km and 6150 km respectively. Although their ranges are similar

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to that of this design, the passenger capacities are far less than that required for this design, 215
and 180, respectively. There are planes that hold the required number of passengers such as the
Boeing 777 (314-450) and the Airbus A380 (440), but their ranges far exceed that of this mission
and the aircraft would be underutilized for its design.
An unlisted constraint is to design an original aircraft that is not similar to the solutions already on
the market. In order to validate the design as a feasible solution, it must not create any new
problems from market aircraft, but also be something new, more efficient, more cost effective,
eco-friendlier, reduce noise pollution, faster, and more passenger friendly. For the purpose and
success of this design, iter8 focused on creating a brand-new solution not yet seen in the
commercial industry that has increased efficiency for the range requirements.
Initial Ideas
Based on these parameters and constraints, a few unique concepts were generated. The first was a
delta wing concept (DELT) of a single fuselage, double aisle idea that featured a delta wing
towards the back of the aircraft. The second was a single fuselage, double decker (DDeck) that
resembles the form of a normal cylindrical fuselage and wing concept. Another possible solution
was utilizing telescoping wings (TELE) for different stages of flight on a single fuselage, double
aisle traditional cylindrical body idea with the engines placed on the tail. The final idea was a
blended wing body (BWB) design that was a flying wing design with engines on the outside of the
main body. The constraints listed above were used to design a decision matrix as seen in Table 2.
Table 2. Decision matrix rated with 1-5, the highest score being most desirable.
Concept Aerodynamics Originality Efficiency Manufacture
Ability
Ergonomic Total
DELT 4 3 3 3 3 16
DDeck 3 2 2 4 4 15
TELE 2 5 2 2 2 13
BWB 5 4 4 2 3 18
The BWB design totaled the highest overall score based on the possible parameters given. Given
its attributes as a flying wing, it ranked the highest in aerodynamics and also ranked highly in
efficiency for
Similar Solutions Study
The blended wing body design was explored further through the use of trade studies of three
different aircraft: the Northrop Grumman B-2 Spirit, the Airbus MAVERIC, and the Airbus Beluga
as seen in Figure 3.

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Figure 3. Displayed are the original 3 similar solutions to be investigated/compared. From left
to right: the Airbus MAVERIC, the Northrop Grumman B-2 Spirit, and the Airbus Beluga.
These three solutions were simply used for proof that this type of design could work. The
MAVERIC was chosen because it is a conceptual design that can be investigated to understand
setbacks/challenges and what the successes are thus far of the aircraft. The B-2 was investigated
in order to discover properties of blended wing body flight. However, its mission is much different
than the parameters of this design problem, such as, speed, size, capacity, and material. The study
was mainly used for successful flight of a full sized BWB and aspects like takeoff and landing.
The Beluga was used for its odd shape. In order to fit passengers into a BWB design, it would have
to be wide in diameter, but not in all places and would require some sort of asymmetric fuselage
shape. The Beluga has an eccentric shape compared to current market aircraft, so it was studied to
uncover the challenges of flying with an asymmetric fuselage present.
INITIAL AERODYNAMICS
Airfoils
The BWB design would require a section large enough to fit payload and passengers but would
still have to be able to taper towards the wing tips without creating extreme areas of stress and
separating the wing from the body. Consistent with a BWB design, the entire body would have to
be an airfoil. Through the use of supercritical airfoils, the design would be able to hold payload
while still being an aerodynamic shape. As seen in Figure 4, the section towards the front of the
airfoil is thicker than the rest, allowing more space for passengers and cargo.
Figure 4. Displays an example of a supercritical airfoil shape versus a conventional design.

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Supercritical airfoils were chosen over conventional airfoils not only for their ability to store more
payload towards the front, but because studies have shown that they create more efficient and
cheaper flight. The flat top on the airfoil slows down the acceleration of air over the wing and
delays the onset of the shock wave and also reduces aerodynamic drag associated with boundary
layer separation. This leads to cheaper and more efficient flight. The graph seen in Figure 5
demonstrates a supercritical airfoil’s drag coefficient based on Mach number versus the same for
a more conventional NACA airfoil .
Figure 5. Graph of coefficient of drag versus Mach number.
The NACA airfoil yields higher values of drag beginning at around 0.7 Mach. Most commercial
aircraft cruise at around 0.8 Mach, so supercritical airfoils become much more efficient for cruising
speeds, reducing the cost of flight and drag. Thus, this further validates the reason to use
supercritical airfoils in the design.
Once it was determined that supercritical airfoils would most effectively meet mission
requirements, a family of airfoils had to be decided on that yielded the best values of lift drag ratio
for various angles of attack. XLFR5 was used to compare airfoils as seen in Figure 6.
Figure 6. Lift drag ratio curves for supercritical airfoils.

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The analysis performed using XFLR5 showed that NASA second generation supercritical airfoils
consistently outperformed every family of supercritical airfoils at lower angle of attacks. Final
airfoil selection was based on incompressible performance at high Re numbers and sizing
requirements. The biggest restriction when choosing airfoils was the thickness to chord
requirements for the fuselage/cabin region. The airfoil must have been large enough to house the
entire cabin and support systems with minimal altercations. However, not all of the airfoils could
be the same thickness due to the thickness of the fuselage section being very large, and the
thickness of the ends of the wings being thin. For this reason, the Manta was split up into three
different sections of airfoils.
The Manta can be divided into three sectioned airfoils as seen in Figure 7.
Figure 7. The left image shows how the three airfoils are split up over the body. The red section
matches the top airfoil (SC2-0518 with t/c=.18 and 𝑆 𝑤𝑒𝑡=1041.5 𝑚2
), the green section is the
middle airfoil (SC2-1010 with t/c=.10 and 𝑆 𝑤𝑒𝑡=607.5 𝑚2
), and the yellow section is the bottom
airfoil (SC2-1006 with t/c=.06 and 𝑆 𝑤𝑒𝑡=35.1 𝑚2
). Where t/c is the thickness to chord ratio and
𝑆 𝑤𝑒𝑡 is the wetted surface area.
For the aircraft body, where the fuselage on a traditional aircraft would be, the NASA SC2-0518
airfoil has been chosen based on it being the thickest airfoil available in the database; it was also
the only one that was able to fit the cabin.
For the mid-wings, the NASA SC2-1010 is picked based on preliminary results from XFLR5
analysis: SC2-1010 has demonstrated its ability to ensure performance while still providing some
thickness at the center mid-section of the wing where fuel can be stored.
For the end-wings, the priority lands on providing reliability in terms of a structure that provides
high lift, low drag and high endurance. SC2-1006 was chosen because it was the thinnest airfoil
with good performance for weight reduction in the yellow region.
Aerodynamic Design
As preliminary airfoil research was being conducted, the general shape and geometry of the aircraft
was also being determined. Designing the body focused on two primary objectives, ensuring that
the central part of the wing can adequately hold the required payload and reducing the wetted area
of the aircraft. To streamline the design, the flow process illustrated in Figure 8 was used.
Furthermore, MATLAB was also used to rapidly generate sizing parameters, while Fusion 360
was used as a CAD platform due to its ability to easily import and size airfoils.

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Figure 8. Decision chart/design process used for preliminary body design.
Initial research revealed that the aspect ratio for a blended wing body (BWB) aircraft was typically
lower than most commercial planes [Ko]. This is attributed to the webbed region between the
fuselage and main wing that traditional aircraft do not have. The aspect ratio was worsened by
realistic wingspan limitations and airport infrastructure requirements. In effort to overcome this
issue, it was decided that the cabin would consist of two decks. This design decision, along with a
substantial leading-edge sweep and moderate trailing-edge sweep, allowed for a 43% increase in
aspect ratio. The significance of this optimization is discussed in greater detail in the proceeding
sections but can be qualitatively observed in Figure 9.
Figure 9. Initial fuselage design (Left) and final fuselage design (Right). The initial fuselage
design included a longer center chord length and larger planform area.
In addition to seeking improvements in drag reduction, another cornerstone of this design features
foldable wings for improved ground mobility and airport access. Inspired by Naval and Marine
Corps Aviation such as the E-2 Hawkeye and V-22 Osprey, and motivated by competition such as
the Boeing – 777X, the Manta will feature foldable wings to assist with ground operations.
According to a study done by Airbus, traffic at airport terminals is forecasted to increase alongside
growing markets and estimates that the largest airports will operate at more than 10,000 monthly
landings per runway [Global Market Forecast]. In effort to plan for impending congestion at airport
gates, the Manta will feature a folded span of only 45 m as seen in Figure 10.

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Figure 10. Dimensioned front view of wingspan folded and unfolded.
Although there will be an increase in operating cost associated with extra maintenance, this cost
will be offset by savings in airport space associated fees. According to the U.S. Government
Accountability Office, the expected cost of airport infrastructure projects to meet increasing
demands will most likely exceed $22 billion between 2019 and 2023 [US Government]. While
some of the cost will be funded by federal airport improvement grants, airliners may soon be
subjected to additional fees. Additionally, the decision to fold the wing so far inboard gives the
Manta a competitive edge at smaller domestic airports.
High Lift Devices
High lift devices ensure an extra layer of safety during takeoff and landing maneuvers and improve
low speed performance. By design, BWB aircraft have a much lower wing loading when compared
to conventional aircraft and would suggest improved performance during takeoff and landing. The
planform shape, however, makes BWB aircraft more susceptible to wind gust at larger angles of
attack and encourages the development of high-lift devices for BWB [Chen]. Additionally, without
the same horizontal control surfaces found on a conventional tail, the use of typical slotted flaps is
not recommended on BWB aircraft [Liebeck]. Thus, due to inherently different control
characteristics, unconventional methods of improving low-speed performance may have to be
investigated. Instead, a simple hinge flap offers a more suitable pitching moment while still
delivering improved performance at low speeds.

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Figure 11. 𝐶 𝐿 with various flap deflection angles. Airfoil analysis using simple flap was
performed using XFLR5
One new type of control surface known as a belly-flap may prove to significantly improve the low
speed performance of a BWB aircraft. Wind tunnel testing showed that an optimal belly-flap
demonstrated a 20% to 30% increase in lift-coefficient at low angles of attack without
compromising pitching moment [Staelens]. When compared to other alternatives such as thrust
vectoring, controllable belly-flaps may offer a far less complex and practical method of generating
higher lift. Additionally, high lift lead edge devices such slats or Krueger flaps may offer the same
benefits as on conventional aircraft.
In addition to the high lift device installed, air brakes will also be installed to ensure that the aircraft
can slow down to a safe landing speed.
CONFIGURATION
Cabin and Fuselage Design
Once the general shape of the aircraft was decided, the configuration inside must be determined.
The first issue was a desire to keep passengers in the thickest airfoil, most centralized to the body
of the aircraft. Most importantly, this would allow for pressurization throughout a uniform
cylindrical cabin size as opposed to distributing a cabin across the tapered BWB. In addition, this
would reduce the amount of turbulence, stall, and G-forces that passengers experience, as the
closer to the wing tip one gets, the stronger these forces become and lead to a decrease in passenger
experience. With that being said, a single fuselage could not be utilized because the aircraft would
become too long and skew the desired aspect ratio in order to accommodate 400 passengers. At
this point, it was determined that a multi fuselage design was required for success.
After research of similar ideas, it was determined that, in order to reduce stress, more than two
fuselages would be needed. In order to not make the aircraft too wide, it was decided that fuselages

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would be stacked on top of each other. This decision was verified in the research conducted by V.
Mukhopadhyay in which it is stated that “a modified vaulted shell partial multi-bubble type
fuselage which has better stress distribution, for same material and dimension” is the best
organization for fuselages in a BWB design. It can be seen in Figure 12 what a possible
configuration could look like.
Figure 12. Demonstration of a three level two fuselage design.
Based on the width and height of the middle section of the airplane, a fuselage design that
comprised of six cabins as seen in Figure 13 was determined to be most efficient.
Figure 13. Layout of a double decker fuselage with three cabins per deck.
A 5 person economy fuselage layout was based on the design parameters given in the handout and
assuming that a 0.508 m aisle space must be maintained. To max structural integrity, it was
assumed that each cabin section was a perfect circle. It can be shown that increasing the diameter
of the fuselage to seat a 6th passenger (Δd) would cause an overall increase in height of 2Δd, that
is ℎ 𝑓= ℎ𝑖 + 2Δd. A height difference can be seen in Figure 14.

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Figure 14. Height difference in cabin with increased diameter.
Since the airfoils making up the fuselage have a max t/c of .18, the new minimum center chord
length will be:
Where cfinal and cinit are the final and initial center chord lengths, and Δd is the required change
in cabin diameter. What Equation 1 is essentially saying is that changes in the diameter will result
in a change in the center chord to match a t/c of .18. Thus, introducing the first constraint in
dimensioning the cabin.
Similarly, removing passengers and reducing the diameter of fuselages has a similar negative
effect. For every 24 passengers removed, an additional row must be set to accommodate them.
Reducing the diameter to accommodate 4 economy passengers per cabin would possibly decrease
the number of business seats per cabin, thus increasing the center chord length even more.
The desire to reduce the center chord length comes in terms of improving aspect ratio. The area
center of a blended wing body can be approximated using a set of simple triangles as seen in
Figure 15. Where c represents the initial center chord length, c’ represents the extended cabin, and
b is the span. The area of the triangle can be represented as a function of the chord as

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Figure 15. A schematic of area of a BWB design
Assuming that b must be held constant due to airport infrastructure limitations, the aspect ratio of
the aircraft can be written as
Or
Which shows that the AR for the aircraft quickly diminishes as Δd increases. Also, it is also worth
mentioning that Δd is essentially quantized by the dimensions of the required pitch. It is also worth
mentioning that the primary goal of the blending wing body is minimizing drag. Thus, increasing
S also has the negative effect of increasing the wetted area, 𝑆 𝑤𝑒𝑡. Since parasite drag is directly
proportional to 𝑆 𝑤𝑒𝑡, any unnecessary exposed surfaces negatively impact performance. A full
aircraft view with fuselage design can be seen in Figure 16.
Figure 16. An isometric image of the aircraft with the fuselage design inside.

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Cabin Layout
The cabin layout features an identical double deck 14-row design with 3 rows being business class
as seen in Figure 17. This particular configuration allots for 72 business class customers (as seen
in blue) and 330 economy class (as seen in green) for a total of 402 passengers. According to an
airline’s needs/luxury level, these numbers can be adjusted, and specific cabin layouts can be
designed at an extra cost. The layout features 2 lavatories per individual cabin (as seen in yellow)
for a total of 12 lavatories on board. In addition, there are 2 extra seats per cabin (as seen in red)
for flight attendants if needed. Emergency exits are located at the red arrows on each deck. There
is also a door on the same row for the middle cabins to escape to the outer cabins should there be
an emergency. Since the wing lays on top of the aircraft, it will not interfere with these doors in an
emergency. These are in addition to entry doors and galley doors in the front and rear of the aircraft.
The middle cabin on the upper deck continues all the way to the cockpit in order to keep it
pressurized as well.
REAR
FRONT
Figure 17. Basic cabin layout for one deck of the Manta.

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Doors
In efforts to fit into the desired gates already mentioned at smaller airports, the double deck Manta
will need to accommodate the gate types available at these airports. There are two entrance doors,
one on each deck, signified by the green arrow in Figure 17. They are slightly skewed so the
boarding process is not disturbed by jetway/stairs interfering with each other. Not both doors are
necessary for entry, as there are stairs at the front of the plane that connect the two decks. However,
the recommended loading process is to attach the jetway to the upper deck door and have a ground
loading stair system (or second jetway if available) in place to load the bottom deck. In this case,
the upper deck would have reserved seats for handicapped persons so that boarding through the
use of stairs would not be an issue. The galley doors, where service and food come, are signified
by blue arrows. In case of an emergency, all doors can be used as emergency exits.
Cockpit
The next significant step after cabin design was addressing the avionics to be integrated into the
Manta. In order to meet the requirements laid out by the CFR, a plethora of various instrumentation
devices are necessary. To quickly list them all, these instruments include an altimeter, an airspeed
indicator, a variometer, a magnetic compass, an artificial horizon, a horizontal situation indicator,
a turn indicator, a free air temperature indicator, a machmeter, an aural speed warning device, a
VOR, and an NDB. Since integrating all these instrumentation devices into our aircraft is
extremely complicated, it was decided to use a previous solution. The flight deck configuration
and avionics layout was completely replicated from that of the Boeing 737. The reasoning behind
this was that the Boeing 737 is the single most common commercial airliner so it would provide
maximum familiarity for the pilots. The flight deck meets all CFR equipment and safety
requirements. The price for all the equipment pertaining to the flight deck and the instrumentation
totals out to $1.5 million. This is a fairly steep price for the avionics, but it provides the most
reliability and minimizes the chances of pilot confusion. The pilots would board through the upper
deck loading door and walk into the middle cabin section that leads to the flight deck. Figure 18
below depicts the flight deck configuration that will be implemented in our aircraft.
Figure 18. Flight deck of 737 to be used in the flight deck of the Manta.

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WEIGHT
The weight sizing process consisted of two major sizes, the initial sizing and the refined sizing.
The former was guided by Raymer’s method in Ch. 3 that consisted of weight-fraction estimates
and culminated with a take-off gross weight (TOGW) calculation. The resulting assumptions and
computations drove the initial performance analyses that would in part inform, along with
Raymer’s Ch. 6 and Ch. 15, the refined weight sizing.
Before running numbers to satisfy the mission requirements for this specific new plane design, we
tested the fidelity of a newly written MATLAB script following Raymer’s Ch. 3 weight calculation
process. This was done by using current-day planes such as the B-2, the A330-743L, and the B777-
200 as test-subjects for the script and then comparing calculated values to the real-life ones. In
addition to closely matching TOGW calculations, the MATLAB script was also compared to
another script written within the team using Nicolai’s Ch. 5 weight-calculation method. Both
scripts yielded similar results, so the team decided to confidently move forward with Raymer’s
initial sizing estimates.
The initial sizing took several mission requirements into consideration, while also making various
assumptions. The mission requirements that were addressed include the weight of the 400
passengers, the 10 crew members, and the 30 lb luggage carried by each. Also, a total nautical mile
range of 3700 was assumed in order to account for the extra 200 n.mi. that an emergency mission
would require. Additionally, the script accounts for 30 extra minutes of endurance and 6% reserve
fuel, including 1% of trapped fuel that will never be used but inevitably contributes to the fuel
weight. The rest of the assumptions were handled by utilizing the given tables and graphs in the
chapter. Thus, we arrived at an SPC of 0.5 for cruise and 0.4 for loitering by matching Manta’s
propulsion system to a high-bypass turbofan, weight fractions for historical mission segments, an
empty weight fraction based on a statistical curve-fit equation that used constants for a jet transport,
and an L/D_max that best matched the BWB configuration discussed earlier. These numbers and
an initial weight guess participated in the iteration process to determine the team’s first TOGW.
Initial Sizing and Weight Estimates
According to Nicolai, the general weight of an aircraft is deduced into three primary branches:
𝑊𝐹𝑢𝑒𝑙 , WFixed, and WEmpty. All three of these variables, once summed together, result in the initial
takeoff weight of the aircraft, WTO. Sorting through each of those three sizing quantities, it is seen
that they have their own respective categories. WFuel, for example, is a fairly simple calculation,
where the formula outputs the amount of fuel needed to complete the flight path:
However, the formula to calculate the fixed weight, WFixed, is rather complicated, as it contains the
weight of the crew, their equipment, the passengers, the food, and the drinks. With all this being
said, the value of the WFixed is as follows:

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The final component of the Takeoff Weight, the empty weight, WEmpty, has numerous inputs: the
structure of the aircraft, the propulsion system, the subsystems within the aircraft, the avionics
system, and all the other instruments on-board.
Once all three of these values, the WFuel, WFixed, and WEmpty, have been calculated, they can all be
inputted into the following equation, to determine the final output, WTO:
After further analysis by the rest of the team, the scripts were adjusted with numbers that better fit
the mission requirements as met by the specific plane design. Changes in the script included cruise
velocity, L/Dmax, and SPCimprovements. Furthermore, the Ch. 3 mission segment weight fractions
based on historical trends were refined using Raymer’s Ch. 6. These modifications led to further
convergence among analyses performed throughout the team.
The next step became to use the refined weight sizing method communicated in Raymer’s Ch. 15.
This process involved calculating the weights of different components within the empty weight
using statistical weight equations. Since the equations take in a design TOGW, iterations had to be
performed. This was done by replacing the statistical empty weight fraction equation of the
previously used MATLAB script with the quotient of the new empty weight calculation and the
TOGW. As a result, the script came up with a more accurate figure for TOGW that takes individual
empty weight components into account. This final result became the impetus for the last major
performance analyses.
Figure 19. The logical structure of an iteration flow chart loop.

21
Material Selection
The determination of materials became a major assumption that affected the weight estimates,
with an additional influence on the cost of the aircraft. Composites, namely, have seen rapid
growth in the aerospace industry due to many benefits such as their high specific strength ratio (a
measure of strength to weight). The Boeing 787 illuminated the team’s path toward greater
composite usage by having an increased structural weight percentage dedicated to composites
compared to older aircraft. The other driving material, steel, represents the other spectrum of
weight selection. Since iter8 Manta will have folding wings, the hinges will need to be reinforced
greatly to secure a safely fixed broader wing section. Figure 20 below presents a rough scheme
of the airframe structure, split up by structural components with their selected material. Figure
21 displays the effect of material weight on the overall takeoff gross weight. In this plot, the
Fudge Factor implies the material selection by solely considering weight. A Fudge Factor of 1
represents the baseline structural weight for a cargo or transport aircraft, as given by the
appropriate Ch. 15 equations [Raymer]. Increasing the multiplier amounts to selecting a heavier
material such as steel, while decreasing the factor below 1 indicates the selection of a lighter
material such as composites. The strongest influence on TOGW, as presented by the figure,
ended up being the weight of the wing. Thus, most of the wing is composed of composites, with
only a small fraction of wing weight dedicated to reinforcing the hinge.
Figure 20. Material structure of Manta

22
Figure 21. Trade study of how changes in the “fudge factor” affect TOGW where heavy
materials (steel) have a fudge factor greater than one and light materials (composites) have a
fudge factor of less than one.
PROPULSION SELECTION
After outlining the project requirements, it was clear that the propulsion team had three primary
duties: selection of the powerplant, selection of the avionics instrumentation, and assuring
compliance with the Code of Federal Regulations. Initially, the task with the highest priority was
the determination of how to power the aircraft. At the beginning of the project the propulsion team
had very little to work with in terms of the configuration and aerodynamics of the aircraft. So,
some of the first steps were producing initial estimates that would help ballpark the power required.
The majority of these estimates were based off of aircraft already in service, most notably the
Boeing 737. Published values of the 737, were adapted to satisfy the mission requirements. This
included accounting for a change in the number of passengers and amount of cargo onboard the
aircraft. The first estimate called for roughly ten thousand horsepower provided by the engines.
As the characteristics of the aircraft continued to be updated, the team utilized these values to
iterate on the necessary power required.
Two methods to select an engine for an aircraft were used. The first was to take the plane’s weight,
aerodynamic characteristics, and desired flight plans, and use those numbers to compute ideal
engine characteristics. These characteristics would be passed to a company like General Electric

23
or Rolls Royce to design a brand-new engine for the parameters. Alternatively, one can look at the
characteristics of existing engines and pick the one that best fits mission criteria to save money
and time. One benefit of using pre-existing engines is shorter development time, as they’ve already
been tested. At an early stage, it was unknown how many engines would be on the aircraft. Early
research showed 31 pertinent engines to use. A script was created to compute performance
characteristics for every available engine and every number of engines per configuration up to 5,
and then ranked results according to a cost function.
In addition to a list of engines, the script was provided with initial estimates of aircraft
characteristics (𝐶 𝑑𝑜, 𝐶𝑙𝑚𝑎𝑥, S, b, etc.), mission requirements (takeoff and landing lengths, range,
payload, etc.), and a set of priorities.
Table 3. Initial Priorities When Running Engine Selection Script -- note that there’s nothing
special about the exact values. All that matters is their size relative to one another.
Table 3 displays the prioritization of design parameters in different runs of engine data. While
running, the script keeps track of which configurations meet mission requirements. After
computing values for all 155 engine & engine count configurations, each set of values is divided
by its maximum. For example, if the possible cruise speeds range from 130 to 260 m/s, every value
is divided by 260 m/s. This makes the values dimensionless. They are then centered on 0 by
subtracting out the mean. Finally, multiplying by the corresponding priority produces a score.
Scores in distinct categories (cost, cruise speed, payload, etc.) are summed to obtain an overall
configuration score.
Priority Run A Run B Run C Run D
Low Cost 0.5 0.5 0.0 0.5
High Cruise Speed 0.1 0.1 0.5 0.5
High Payload 0.1 0.5 0.1 0.0
High Range 0.1 0.5 0.2 0.0
Short Runway 0.1 0.1 0.1 0.0
Tight Turn 0.1 0.1 0.1 0.0
Low cj 0.5 0.1 0.5 0.5

24
Table 4. Results Obtained using initial priorities and initial aircraft characteristics (Ranked by
overall score)
Results for Runs A through D (whose priorities are listed in Table 3), are available in Table 4. In
the color coding of Table 4, a few engines showed up repeatedly, regardless of which priorities
were entered. This made them prime candidates for the design the GEnx-2B67B. In order to
minimize cost, maximize fuel efficiency, and maximize cruise speed, as seen prioritized in Table
2, the two-engine configuration was highest ranked.
Using Priority Set D and refined aircraft characteristics from the weights and aero teams, the final
engine choice was 2 GEnx2B67B engines. Using Priority Set D and refined aircraft characteristics
from the weights and aero teams, the final engine choice was 2 GEnx2B67B engines. All possible
configurations consisting of 1 or 2 engines are plotted against the nondimensionalized criteria
selected in Priority Set D (cost, range, and efficiency) in Figure 22.
Ranking Run A Run B Run C Run D
1 3 x GEnx-2B67B 3 x GEnx-2B67B 5 x GEnx-2B67B 2 x GEnx-2B67B
2 4 x GEnx-2B67B 4 x GEnx-2B67B 4 x Rolls Royce Trent
XWB-97
1 x GE9X
3 2 x Rolls Royce Trent
XWB-97
2 x Rolls Royce Trent
XWB-97
4 x GEnx-2B67B 1 x Rolls Royce Trent
XWB-97
4 3 x GEnx-1B70 3 x GEnx-1B70 5 x GEnx-1B70 1 x Rolls Royce Trent
XWB-84
5 3 x Rolls Royce Trent
XWB-97
XWB-97
XWB-84
3 x GEnx-2B67B

25
Figure 22. On the left graph, configurations which met or exceeded the mission requirements
are marked by a star, while others are simply dots. The graph on the right is the same graph with
only the starred options, of which there are 5.
Using the initial estimates, there were dozens of engine configurations that fit the mission criteria.
However, after refining the weights and aerodynamics of the design, there were only 5 engine
configurations that could fulfill the mission. Of those 5, 4 are single engine options. For safety the
2 GEnx-2B67B (Figure 23) design was selected. The pair of engines would be mounted on the
topside of the aircraft towards the rear. Although there were singular engines that were powerful
enough to meet the requirements of the design, a minimum of two engines was necessary. Since
safety and reliability is the number one priority, we came to this decision to avoid disaster in the
event of a bird-strike or potential engine failure.
Figure 23. The GEnx-2B67B engine on a Boeing 747-8
The GEnx-2B67b engines are already extremely common in the industry of commercial aviation.
These engines are the primary power plants for both the Boeing 747-8 and the 787 Dreamliner.
One of the primary benefits of the General Electric engines is the high bypass ratio of 9.6. A high
bypass ratio is extremely beneficial, especially for a blended wing body aircraft. Due to the
location of the engines, noise pollution inside of the cabin is much more of a concern than for a
typical aircraft with engines mounted on the wings. The excess air flowing around the core of the
turbofan engine significantly hinders the noise pollution and provides a more comforting
experience for the passengers. Another benefit of the GEnx-2B67b engines is the remarkable fuel
efficiency. Compared to General Electric’s previous engine the CF6, which was the primary
power plant for numerous Airbus airliners, these engines provide a 15% increase in fuel
efficiency. The only downfall of the propulsion system pertains to the location of attachment for
the engines. In Figure 24, the location attachment of the engines is depicted. Although placing
the engines on the rear topside maximizes aerodynamic efficiency, it creates a potential danger.
In the event of engine detachment, the engine could potentially thrust itself forward and land

26
directly on top of the fuselage. This risk threatens not only the structure of the aircraft but the
passengers as well. As a result of this risk, extra safety precautions will be taken in regard to the
engine mounts to avoid this from happening at all costs.
Figure 24. Manta with engine placement
Combined with the aerodynamic characteristics of the Manta, the GEnx-2B67b allowed for every
mission requirement to be met. This included range, takeoff, and landing requirements.
FLIGHT RESULTS/PRELIMINARY FINDINGS
Final Weight
The following table demonstrates the refined weight sizing results, where components are grouped
into three main subsystems: empty, fuel, and payloads weight. The empty weight is further divided
into structural, propulsion, and equipment weight. Noting the steady increase of empty weight over
an aircraft’s lifespan, a 2% empty weight allowance was accounted for in these calculations as
recommended by Raymer [Raymer]. Additionally, the components weights are presented in the
middle column as results from the bare equations in Ch. 15 and in the right column as the same
results enhanced by so-called “fudge factors” that consider material weight [Raymer].

27
Table 5. Weight breakdown of all aircraft components and payloads.
Volumes
The volumes were calculated using the measuring tools within CAD software and design
requirements based on number of passengers as seen in Table 5.
Table 6
Component Volume [m3]
Fuel Tank 54.6
Entire Fuselage 1825.4
Each Mid-Wing Section 142.3
Pressurized Cabin 485
Drag

28
Once airfoil selection and design parameters were finalized, the complete drag profile for the
Manta was developed. After surveying a range of possibilities for estimating the parasite drag
coefficient, CD0, it was determined that the method outlined in Commercial Airplane Design
Principles seemed to be one of the more conservative approaches and explicitly described the
calculation parameters for supercritical airfoils [Sforza]. Expressed in individual components of
form factor, K, coefficient of skin friction for turbulent flow, CF , wetted surface area , Swet, and
planform area, S, the parasite drag coefficient was analytically expressed.
Where the ith iteration is the component.
Since most of the literature regarding parasite drag coefficients include a conventional fuselage
design and a wing with a fixed thickness to chord ratio, it was recommended to calculate the
parasite drag coefficients of the body using a weighted function of the wetted surface area. The
body of the aircraft was broken up into three distinct regions as shown in Figure 7. Each region
of the aircraft is categorized by thickness to chord ratio and uses its own mean aerodynamic
chord when calculating parasite drag parameters.
Where CD,Body represents the parasite drag coefficient of the body, and the subscripts F,M,W
represent the regions illustrated in Figure 7. According to Sforza, the form factor K varies with
airfoils but can be expressed as ( 1 + 4.2(t/c)max) for supercritical wings. Alternatively, the form
factor for the vertical tails and nacelles are represented by 1.25(1.2(t/c)max + 100(t/c)4
max) [Sforza].
Lastly, since CF is dependent on Reynold’s number, it is worth mentioning that Re calculations
were estimated at an altitude of 12,000 m, a free stream velocity of 250 m/s, and using each
component’s respective cMAC. Table 7 conveniently summarizes the individual components of the
entire aircraft.
Table 7. Break down for parasite drag by body, nacelle, vertical tail, and auxiliary unit
contributions. * Note that the body contribution is a weighted function of the three regions that
make up the body as shown in Figure 7. The individual regions are not directly considered in the
final summation.
Component K CF SWET (m2
) CD
Fuselage/Cabin 1.756 .0021 1041.5 .00789
Mid-Wing 1.420 .0023 607.5 .00725
End-Wing 1.252 .0027 388.9 .00739
Body* - - - .00759
Nacelles 1.12 .0028 72.9 .000484

29
Vertical Tails 1.28 .0030 35.1 .000284
Auxiliary Units - - - .000412
Σ CD
.00877
As previously mentioned, BWB aircraft often have a reduced aspect ratio as a consequence of their
design and realistic wingspan limitations. Although a cornerstone of the design process was
reducing drag, the final design called for a compromise between the parasite and induced drag. It
can be shown in Equation 3 that the induced drag coefficient is inversely proportional to aspect
ratio and goes to show how an increase of 43% can greatly improve performance.
As a compromise between induced and parasite drag, the benefits of a lower parasite drag
coefficient at typical cruising speed outweigh the consequences of a lower aspect ratio [Anderson].
Mach
Since the aircraft will also be operating in the lower bounds of the transonic region, compressibility
effects and wave drag begin to influence performance. Supercritical airfoils offer superior
performance when approaching Mach 1 and are exclusively used on the Manta. Due to its
inherently difficult nature, a series of equations and methods were used to determine the
compressibility drag coefficient. The drag divergence number was estimated via the Korn equation
as described in Commercial Airplane Design Principles and Configuration Aerodynamics as
[Sforza].
Although this equation represents the drag divergence number for an airfoil using a 2D lift
coefficient, cl, it can be used to reasonably determine the critical Mach number and wave drag for
a wing using the quarter-chord sweep angle and thickness to chord ratio [Mason].
These equations represent the critical Mach number and wave drag coefficient, respectively.
Although compressibility effects are expressed in terms of 2D parameters, results were validated
against a Boeing 747-100 with reasonable accuracy [Mason]. They also imply that the
approximation must be broken into regions based on relative thickness. Like the approach for
calculating CD0, a ratio between area spanwise strips, Sstrips, and the entire planform area is used.

30
Using the graphical and curve fitting method shown in Fundamentals of Flight as a method of
validating results, it can be seen in Figure 25 that the results from the Korn method converge with
the anticipated results. Although the aircraft faces an abrupt change in drag due to compressibility
effects, it still has the potential to operate near .8 Mach and retains its competitive edge as a
commercial transport vehicle. Unfortunately, the tapering of the aircraft caused a decreased effect
in the sweep back angle but was done so to improve drag performance at more reasonable Mach
values.
Figure 25. A comparison between the Korn and Shevell methods for Determining Wave Drag.
The aircraft was assumed to be operating under initial vehicle estimations at an altitude of
10,000 m.
𝑪 𝑳 Values
The equation for CLCruise utilizes the Carson speed and the equation
𝐶 𝐿𝑐𝑟𝑢𝑖𝑠𝑒 = √(
1
3
) ∗ 𝜋 ∗ 𝑒 ∗ 𝐴𝑅 ∗ 𝑐 𝑑𝑜
And yields a value of 0.21.
According to Raymer, CL,MAX for most aircraft with a moderate quarter-chord sweep angle can be
approximated as
Where cl and Λc/4 represent the 2D cl,MAX and quarter chord sweep angle [Raymer]. By utilizing
data shown in previous sections, it can be shown that the CL,MAX for the given wing
configuration is approximately 2.2. However, factoring in safety, uncertainty, and unaccounted
3D effects, the Manta is more likely to operate with an actual CL,MAX of1.7. This adjustment also
seems to support other early estimations in literature surrounding BWB designs [Liebeck].

31
By using the low speed cl data gathered from XFLR5 and applying the Prandtl-Glauert rule, to
correct for compressibility, the CLMAX equation can be rewritten as
where cl,0 is the low speed airfoil data gathered using XFLR5, M is the free stream Mach number,
and 𝚲c/4 is the quarter-chord sweep back angle of the wing [Anderson].
Figure 26. Estimated CL vs 𝛼 curve for the entire wing. This graph includes a .8CL reduction to
account for safety. The trend of increasing Mach number and increasing CL is identified.
Velocities

32
Figure 27. Three axis graph that shows altitude and velocity versus time into flight.
The first two minutes of flight in Figure 27 consist of takeoff and flight in the traffic pattern, which
is 304m above ground level. Assume ground level to be sea level for simplicity. Takeoff speed is
defined as 1.2*Vstall,which for Manta is 84.55 m/s. The aircraft then climbs from runway to pattern
altitude (1000ft or 304m) at Vtakeoff, which is 1.2*Vstall. In the second region of the graph (~t=2 to
t=14), the aircraft flies with horizontal speed VPRmin and climbs at a rate of 17.52 m/s. This
represents the most economical rate of climb. Once reaching 12.5km, the aircraft stops climbing
and cruises at V=220m/s -- this is the large, flat section of the graph.
Around t=219, the aircraft begins its descent. There are a number of conditions under which an
aircraft might descend and land, so the numbers in the plot represent one possibility, as opposed
to a prescribed, optimal plan. For example, air traffic controllers might have a plane descend
sooner in order to follow another plane into an airport, or later if there is bad weather. Similarly,
they may fly straight into the runway, or need to perform an S-shaped traffic pattern. In this case,
we designed the flight plan such that it would take 20 minutes to descend from cruise altitude to
pattern altitude, and we expect to remain in the pattern for three runway lengths (3*2438m), plus
the arc length required to perform the turns of an S shaped traffic pattern.
On the graph, descent occurs in two stages. As mentioned, this wouldn’t have to be the case --
the pilot has many options -- but here velocity decreases rapidly until the point that further
decreases would require negative thrust. Afterwards, the plane’s acceleration is limited, so it
slows down more slowly.

33
Finally, after descending to pattern altitude, the plane begins flying at its approach speed, defined
as 1.3*Vstall, or 38 m/s. It performs the S-shaped traffic pattern and lands.
A list of critical velocities can be seen in Table 8. In addition, the takeoff and landing distances
calculated using these velocities are also in Table 8.
Table 8. Critical velocity values through flight
Takeoff Velocity 84.55 m/s
Landing Velocity (Approach Speed) 38.00 m/s
Takeoff Distance (Dry or Wet Pavement) 180 m
Landing Distance (Dry Pavement) 267 m
Landing Distance (Wet Pavement) 284 m
Flight Time (for 700nmi design mission) 1.97 hours
Balanced Field Takeoff Distance 400 m
Flight Envelope
In Figure 28, the flight envelope for the Manta is shown and displays how Vstall and Vcruise vary
with height. Anywhere inside of the curve is flyable.
Figure 28. Flight envelope for Manta for up to the ceiling of 17.5km.

34
Turning Flight
According to FAR regulations, the limit load factor of a commercial transportation aircraft is
determined by the equation
whichever is greater. W represents the maximum gross takeoff weight and must be given in units
of pounds [Sforza]. Assuming that the aircraft can be manufactured to meet minimum
requirements, and using the preliminary design parameters, the anticipated turning envelope is
represented in Figure 29 and Figure 30.
Figure 29. Turning envelope for Manta Ray assuming level and unaccelerated flight.
Figure 30. Load factor vs velocity

35
Although a dynamic loading analysis was not performed, the substantial leading edge sweep angle
and relatively low aspect ratio offer a level of protection against gust loads [ATPL]. These design
parameters result in smaller increase in wing loading in the event of a vertical gust. However, the
effects may potentially be offset by a low wing loading. Due to an inherently larger wing area,
BWB aircraft are more susceptible to strong gusts and may require additional structural
reinforcement. Thus, further analysis on the offsetting effects of low aspect ratio and low wing
loading are required to understand the effects of vertical gust on BWB aircraft.
Stability/Control
In order to find the center of gravity of the Manta Ray, a series of values were extracted from
research: the individual weights of all the components on-board the aircraft, Mn[N], along with
their individual distances from the nose of the aircraft, Xn[m]. The former values were charted in
Table 9, while the latter values are charted below. The moment for each component, MnXn, was
then calculated and implemented the equation below to calculate the location of the aircraft’s
gravitational center.
Table 9. Center of gravity breakdown for all aircraft components and payloads.
Component Distance from nose (m) Moment (Nm)
Empty
Structure
Wing 22.11 6.83E+06
Landing gear 20.34 4.24E+05
Air Induction 44.45 1.38E+06
Empennage 35.5 1.98E+05
Fuselage 15.3 4.99E+05
Propulsion
Nacelle group 35 1.00E+06
Fuel System 26 6.43E+04
Engine controls 15.17 5990.796836
Equipment
APU 15.5625 67189.38188
Instruments 16.0932 4.10E+04
Surface controls (includes hydraulics and pneumatics) 13.47685 1.45E+05
Electrical 10.05085 3.68E+04
Avionics 3.0127 2.26E+04
Air conditioning 8.05085 2.51E+05
Anti-icing 20.375 4.94E+04

36
Furnishing 7.32 2.30E+05
Handling group 5.25 1910.574896
Fuel
Fuel (total) 19 7.08E+06
Payloads
Crew 13.9 248087.5
Crew Baggage 13.5 36045
Passengers 13 9923500
Passenger Baggage 15 1441800
Total Moment (Nm) 29973861.8
x_cg (m) 18.2
Some values, such as the weight of the passengers, had to be distributed over a mesh - as they are
not going to be located in one central location during the entire flight time. Since they are spread
out over a given distance, their weight and distribution distance was averaged to find their most
central location. Afterwards, a summation equation was then used to calculate the center of gravity
of the aircraft.
In the equation above, the center of mass (XCM), which is the output, is generated by inputting
the mass of all the components, Mn, and the location of each component from the nose of the
aircraft, Xn within. This yields a center of mass in the x-axis, with a value of 18.2 m from the
nose of the aircraft. The cg of the aircraft changes as the gross weight changes in flight as seen by
Figure 31

37
Figure 31. Cg Envelope where the plane starts of with a TOGW and lands with empty fuel tanks
FAULT TOLERANCE
One of the key pillars of passenger aircraft design is safety. With a design as unconventional as
the Manta, several concerns arise with respect to ensuring a secure and stable aerodynamic
structure and passenger experience.
With any aircraft, there exists the potential for severe impact to performance due to engine
malfunction and/or blowout. On a typical tube-and-wing aircraft, engines are most often placed on
the underside of the wings or attached to either side of the aft end of the fuselage. On the Manta,
however, these configurations were impossible due to the lack of a proper horizontal stabilizer in
addition to the critical hinge function of the wings. Instead, engines were embedded in the vertical
stabilizers at the aft of the fuselage on the upper face of the body, as shown in Figure 32. With
such a placement, one concern is an engine escaping its nacelle and inertially sailing forward and
colliding with the passenger cabins and/or the cockpit. In order to mitigate such a scenario, the
engines will be housed within an electronic elevator that will allow them to be tilted at a skyward
angle. With a simple kinematic projectile analysis, it was determined that at the Carson cruise
velocity, 1.316*VDmin, the engine will dislodge from the aircraft, causing a backwards impulse.
Inertial momentum will then carry the engine upwards and forwards, allowing time for the aircraft
to glide out of the impact zone before it descends. Unfortunately, the landing zone of the engine is

38
dependent on the location where it becomes dislodged and is completely unpredictable, however
the chance of such an event occurring is low.
The hinge design on the Manta is quite rudimentary but resolute. Inspired and justified by the
uncanny success of the Boeing 777X’s revolutionary folding wingtips, the wing attachments will
be made from the same carbon fiber composite as the fuselage nacelle, with aluminum lining the
leading edge to alleviate some of the aerodynamic loading. The hinge consists of robust auto-
locking steel pins to ensure stability inflight, operated electronically. Boeing has not made public
any relevant operational information regarding the function of the wingtips on the 777X, and as
such ample fault tolerance specific to the hinge mechanism was unable to be performed.
Figure 32. The folding wingtips on the Boeing 777X, the inspiration and proof of concept for the
folding wings on the Manta.
In the event that one or both engines blow out or become incapable of producing thrust,
unpowered glide calculations from the cruising altitude and the landing-loitering altitude of the
aircraft were performed and were tabulated below in Table 10, including values accounting for
failure of one or both inboard wing hinges. One large safety benefit of a blended-wing-body such
as the Manta is its capacity to generate lift and mitigate drag as it has a notably large wetted area
and its profile consists entirely of supercritical airfoils; as a result, its maximum lift-to-drag ratio
is exceptionally high, granting it the ability to fall in style for considerable distances. Analysis
for the unpowered accelerate-stop and accelerate-go distances are included. Glide distance is
calculated as:

39
Table 10. Unpowered gliding distances for several hinge failure scenarios of the iter8 Manta.
Fuel leakage is another concern. Onboard the Manta, fuel is to be stored in isolated tanks on either
side of the passenger cabins within a second isolated section of the fuselage. This mitigates the
concern for fuel to enter the breathing space of passengers, and in addition ensures there are two
formidable defensive barriers—the tank housing and the fuselage housing—between the fuel
supply and open air. In the case of an emergency landing requiring a fuel dump, there will be a
valve which will allow fuel to flow from the tank out the aft clamshell door on the underbelly of
the fuselage.
The Manta is equipped with a swath of electronic equipment, from avionics to virtual reality
window paneling. Much of this equipment is not essential to safe flight, however given the recent
failures of the Boeing 737 MAX and the uniqueness of the flight deck and procedures of flying an
aircraft of this nature including hands-on stabilization and extreme caution in turning flight to
avoid exposing passengers to undue g-force, Manta pilots must be carefully trained and certified
to fly.
LOGISTICS
CFR Compliance
The iter8 Manta is constructed, as the 787, with 50% advanced carbon fiber composites by volume
[Fraga]. While a much more expensive structural choice, this provides security with respect to
several of the CFR part 25 rules regarding fireproofing and fire isolation. Each component of the
aircraft involving a flammable or toxic gas is to be isolated within its own steel vessel, each
insulated with a layer of composite, which is a corrosion and flame retardant material. Each cabin
is to be equipped with fire extinguishing materials, with one fire extinguisher stored in each galley.
Each engine and its fuel supply is housed entirely separately from the other, allowing the pilot to
reserve the ability to stop and start either or both engine mid-flight. Each is symmetrically placed
latitudinally with respect to the center of gravity of the aircraft. The flight deck and instrumentation
package includes all CFR-required instruments, including but not limited to an altimeter,
machmeter, bank and pitch indicator, fuel pressure and quantity indicators, etc. All externally
mounted instruments such as pitot tubes will be attached to the aircraft on the underside of the non
folding wing area, allowing them to be uninhibitedly exposed to the true flight conditions and
easily wired and/or connected to their various sensors through the vast empty volume of the aircraft
fuselage.

40
On Boarding
The passenger experience onboard the Manta will be unparalleled in the current market. As access
to virtual reality technology exponentially increases, traditional entertainment—including
windows—will become proportionally obsolete onboard the aircraft of the future, a class to which
the Manta belongs. Imagine a flight where every passenger had a window seat; this is a feasible
reality for the future of air travel. This aircraft will feature high-quality virtual reality paneling
across its cabins’ walls and integrated interactive seat-back screens to ensure the passengers feel
safe and at home while this daunting experimental layout boasts a cutting-edge windowless frame
as seen in Figure 33.
Figure 33. An artist’s rendering of what the VR interior of a futuristic aircraft might look like.
[Santus]
As with most commercial aircraft, the Manta possesses a robust APU which powers air
conditioning units above every seat in addition to a wide variety of lights, from safety lights lining
the cabins to the VR paneling to the position, navigation, anti-collision, landing, etc. lighting
systems on the exterior of the fuselage. The two-by-three cabin stack layout requires two lavatories
per cabin section, one in business class and one in economy class. Each cabin will have one aisle
and be attended by two crewmembers. The composite materials comprising the fuselage nacelle
will yield improved pressurization efficiency, just as pioneered in the Boeing 787 Dreamliner.
There will be passenger doors connecting to the top deck of the passenger cabins on either side of
the fuselage, however only one will be utilized if the jetway at the airport’s disposal does not have
the capacity to diverge or if multiple jetways are not available. Both exits serve as emergency exits.
Boarding will likely be slightly longer than that of the average flight given the sheer volume of
passengers and the multi-level traffic flow, however this process could certainly be streamlined by
innovative airlines. There will be equipment for full food and beverage service available, but at
the discretion of the operating airline.
COST

41
The iter8 Manta is not a budget aircraft—state of the art innovation designed to provide a premier
passenger experience in addition to affordability and efficiency has a large price tag. Using the
cost model outlined in Table 11, the raw unit material cost of the iter8 Manta was estimated at a
steep $100 million, barring engineering, testing, certification, and manufacturing labor costs.
While this may be a staggering figure, keep in mind that the technology aboard the Manta is cutting
edge, unlike anything on the market. It has the capacity to carry 400 passengers safely and
comfortably for a range of up to ~9,000km, reduce noise pollution with its top-housed high bypass
turbofan engines, reduce passenger discomfort and fatigue as well as mitigate aerodynamic loading
damage with its sleek carbon fiber composite shell, promote sustainability with a supercritical
airfoil profile which reduces drag and therefore fuel consumption, all while fitting comfortably in
domestic gates worldwide with a folded-wings fuselage span of 45m; this is unprecedented.
Table 11. Estimates for the raw unit materials cost for the aircraft. Excludes testing,
manufacturing labor, marketing, and engineering costs. [Arnot]
Item Cost ($Mil.)
Carbon composite fuselage nacelle* 23.8
Glassware 0.026
Aluminum skeleton* 1.46
Steel & Titanium reinforcements* 3.78
Landing Gear 0.80
Tires (8) 0.048
Avionics/Flight Deck 1.5
APU/Pressure Systems 0.35
Seats (economy, business class) 1.05, 0.25
Seatbelts 0.0252
Seatback Entertainment systems 8.0

42
Virtual Reality windows** 3.0
Engines (2 GEnx-2B67B) 51.2
Lavatories (12) 3.0
Paint Job 0.1
Total 97.1
*Based on $85/kg aero grade composites, $0.52/kg steel, $4/kg titanium, $13/kg aluminum
figure [Fraga]
**No accurate estimate for this type of technology; not on the market in 2020 but likely in 2029
The operating cost per flight-hour to break even on the average flight of the Manta is outlined in
Table 12, assuming a fully sold flight. Cost was determined using the flight profile derived and
displayed in Figure 27. In order to conserve as much fuel as possible in accordance with iter8’s
mission to reduce the carbon footprint of the modern airliner as well as foster efficiency and
affordability, the most economic climb conditions were selected. This condition occurs at
minimum power required, i.e.
Where CL,PRmin is the lift coefficient at the conditions for minimum power required, Vclimb is the
most economic climb velocity, and (
𝑅
𝐶
) 𝑒𝑐𝑜 is its corresponding rate of climb. This value was then
iterated through altitudes from sea level up to cruise at 12,500m altitude in order to determine fuel
burned in the climb using:
Where cj takeoff is the specific fuel consumption in
𝑁
𝑁∗ℎ𝑟
at takeoff. These values were then indexed
and averaged in order to determine the average rate of fuel burn as well as the average rate of
climb. These averages were used to calculate the climbing time, tclimb, and the cost of fuel burned
during takeoff:

43
The climbing time was determined to be 11.41 minutes, and Costfuel, per kg= $0.9726 per kg fuel.
These values were then iterated over Newtons of fuel burned, from Nfuel, takeoffaverage to the full fuel
weight. Cost of cruising fuel burn was calculated using the following equation:
Range and endurance were tabulated below in Table 12. Endurance was multiplied by the number
of pilots and crew members and their hourly wages (estimated at $100 and $50, respectively) to
determine crew salary costs. The costs for fuel and crew were added together and a proportion
created according to the cost dividends outlined in Figure 34to estimate total operating cost for
the iterative endurance value and was divided by the Manta’s maximum passenger capacity.
Figure 34. Modern breakdown of total airline operating costs. [ICAO]
Table 12. Operating cost breakdown for the Manta with respect to Newtons of fuel burned.
Time in Flight Range, km Fuel burned, N Operating Cost, $ Cost/pax, $
0:12 85.7 7,479 2,090 5.23
0:30 897 21,479 5,648 14.12
1:00* 1,804 35,479 9,201 23.00
2:00 3,624 63,479 16,302 40.76
3:30* 6,667 94,479 27,820 69.55

44
9:00*** 18,351 259,479 69,081 172.70
* = Completes required 700nmi trip
** = Completes maximum required 3,500nmi trip
*** = Range for full fuel burn
As is evident from these estimates, there is ample room to set competitive pricing with other fleets
of aircraft while still providing a hefty profit margin. Comparison with competing models is
performed in Table 13.
Table 13. Operating cost analysis for the Manta versus its most similar competitors [Wynma]
Make &
Model
SFC at cruise,
N/N-hr
Operating Cost/flight-
hr, $
#
Seats
Cost/pax/flight-hr,
$
iter8 Manta 0.40 9,034 400 23.00
Boeing 787 0.64 8,007 296 27.05
Boeing 777 0.57 11,146 384 29.03
Airbus A320 0.6 8,964 277 32.36
With this pricing model, if charging $500 per passenger from JFK to Heathrow—a flight that
typically goes for $700/passenger—breakeven on the raw unit cost occurs after just 563 flights
(which only take three hours; 4-6 of them can take place on a given day and 563 flights is reached
after just 94-141 days, well under 6 months). Of course, there are many other factors for the airline
including marketing, reservations, staff, scheduling, etc., but it is evident that the Manta is an
extremely efficient investment for any airline.
FUTURE iTER8TIONS
This BWB configuration shows great promise as an alternative aircraft design for the traditional
tube-and-wing that has the potential to serve myriad purposes. The Manta is one such example,
however the body of this plane can be sized up, down, and sideways to make it a suitable candidate
for a jumbo-jet luxury passenger vessel (Albatross), a long-range low-capacity transporter for
expedited, inexpensive, and more environmentally sustainable transcontinental travel (Stingray),
and a high-payload cargo aircraft (Pelican), to name a few. iter8 is proud to offer innovative
solutions to air travel’s most prominent problems, in hopes of providing an efficient, sustainable,
affordable, safe, and uniquely exceptional passenger experience.

45
REFERENCES
[1] Raymer, D.P.,“Aerodynamics,”Aircraft Design: A Conceptual Approach, 2nd ed., American
Institute of Aeronautics and Astronautics, Inc, 1992, pp 270 - 282.
[2] Sforza, P.M, “Drag Estimation,” Commercial Airplane Design Principles, 1st ed. , Elsevier,
Massachusetts,2014, pp 361 - 392.
[3] Liebeck, R. H., “Design for the Blended Wing Body Subsonic Transport,” Journal of Aircraft,Vol.41,
No. 1 January - February 2004, pp 10-24.
[4] Kothia, D., Singh, J. and Dadhich, A., “Design, Analysis, and Optimization of Folding Wing
Mechanism for Effective Utilization of Air Side Area,” International Journal of Aerospace and
Mechanical Engineering, Vol. 3, No. 5, September 2016, pp 22 - 28.
[5] Mason, W.H.,“Transonic Aerodynamics of Airfoils and Wings,” Configuration Aerodynamics, 2006.
[6] U.S. Government Accountability Office, “Airport Infrastructure : Information on Funding and
Financing for Planned Projects,” U.S. GAO, GAO-20-298, Februrary 2020.
[7] “Global Market Forecast : Flying by Numbers”, Airbus, 2019.
[8] Ko, A. Et Al., “MDO of a Blended-Wing-Body Transport Aircraft with Distributed Propulsion,”
AIAA’s 3rd Annual Aviation Technology, Integration, and Operations Technical Forum,. AIAA,2003.
pp 9 - 10.
[9] Anderson, J.D.,“Airfoils, WIngs, and Other Aerodynamic Shapes,” Introduction to Flight, 8th ed.,
McGraw Hill, New York, 2016, pp 356 - 377.
[10] Chen, Z. Et Al “ Assessment on Critical Technologies for Conceptual Design of Blended-Wing-
Body Civil Aircraft,” Chineses Journal of Aeronautics,2019.
[11] Staelens, Y.D.,“Study of Belly-Flaps to Enhance Lift and Pitching Moment Coefficient of a
Blended-Wing-Body Airplane in Landing and Takeoff Configuration”, December 2017.
[12] “ATPL Training/Principles of Flight #69 Limitations - Gust Loads,” Aviation Training Network,
September 2018.https://www.youtube.com/watch?v=BqJO-7z4eRk

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