Conceptual Design Review
May 7, 2010
Team 3
Team XG Intl. Inc
Concept Design Review
Gijung Bae
Joe Blake
Jung Hoon Choi
Jack Geerer
Alison Jean Gong
Sang Jin Kim
Mike McCarthy
Nick Oschman
Bryce Peterson
Lawrence
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ge
Hwan Song
May 2010
Conceptual Design Review
May 7, 2010
Team 3
Team XG Intl. Inc
I Contents
II Executive summary .......................................................................................................... 4
- III Mission Statement .................................................................................................... 6
- IV Design requirements ................................................................................................ 6
- V. Selected “best” aircraft concept ............................................................................... 7
Walk-around ................................................................................................................... 7
Values of major design parameters ................................................................................ 9
- VI. Results of aircraft sizing and carpet plots .............................................................. 10
Brief description of sizing code ..................................................................................... 10
Validation & Fudge Factor ............................................................................................ 14
Fixed design parameter values ..................................................................................... 16
Drag Prediction ............................................................................................................. 16
Results of engine modeling........................................................................................... 17
Mission modeling .......................................................................................................... 20
- VII. Major design trade-offs ....................................................................................... 21
Trade off overview ........................................................................................................ 21
Summary of carpet plot studies ................................................................................ 23
 - VIII. Aircraft description............................................................................................ 24
Dimensioned three-view, to scale ................................................................................ 24
Representative internal layout ..................................................................................... 26
 - IX. Aerodynamic design details / justification .......................................................... 28
Drag Polar...................................................................................................................... 29
- X. Performance ............................................................................................................ 30
- XI. Propulsion .............................................................................................................. 33
- XII Structures ............................................................................................................... 34
Configuration layout and highlights.............................................................................. 36
Important Load Paths ............................................................................................... 36
Wing-fuselage intersection ....................................................................................... 38
Engine mounts and landing gear integration ........................................................... 39
Materials ....................................................................................................................... 41
- XIII. Weights and balance ............................................................................................ 42
Overview discussion on prediction methods................................................................ 43
Center of gravity ........................................................................................................... 45
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- XIV. Stability and control ............................................................................................. 47
Preparations for malfunctions and extreme situations ................................................ 49
Engine failure ............................................................................................................ 49
Extreme crosswind .................................................................................................... 50
- XV. Noise ..................................................................................................................... 50
- XVI. Cost ...................................................................................................................... 51
- XVII. Summary ............................................................................................................ 54
APPENDIX A: Bibliography ................................................................................................ 55
APPENDIX B: Sizing code ................................................................................................... 56
APENDIX C: Compliance Matrix ........................................................................................ 57
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II Executive summary
Team XG International, Inc. sought to develop an environmentally-sensitive aircraft
which will provide our customers with a 21st-century transportation system that
combines speed, comfort, and convenience while meeting NASA’s N+2 criteria.




reduce 40% fuel consumption
reduce noise by 42 dB
reduce NOx emission
reduce takeoff field length
The XG Endeavour’s range of 3,700 nautical miles enables it to make almost any
continental flight, as well as many intercontinental flights, and the Mach 0.80 cruising
speed makes the trip in a hurry. The convertible cabin layout option also allows the
customer to add their own personal touch. While XG International sought to create a
fast and convenient air transportation system, they also required an environmentally
sensitive aircraft. Therefore, the XG Endeavour offers considerable reductions in noise
and exhaust emissions, fuel consumption, and runway field length.
The power of three GE Honda Aero HF120 variant turbofan engines, capable of
outputting a combined 10,800 pounds of thrust, lifts the XG Endeavour into the sky. At
the cruising altitude of 45,000 feet, fuel consumption is greatly reduced when one
internal turbofan is turned off, while the two remaining engines, along with the sleek,
swept back wings, and cruciform tail allow the XG Endeavour to soar to its destination.
During daylight hours, solar film on the upper surfaces of the fuselage and wings collect
energy and provide power for in-flight systems.
XG International engineers designed an all-new sizing code for the Endeavour project.
The jet’s body, geometry, weight, airfoil, design mission, and many other factors were
incorporated in the development of this new code.
In order to compete in the business jet market, the advantages and disadvantages of
every design aspect needed to be weighed against one another. While the three
turbofan engines added weight and considerable maintenance costs, having all three
engines would be safer in an engine-out situation, and it would also be fuel efficient
when turning one off during cruise. The cruciform tail allows for an aft fuselage engine.
Solar film without adding significant weight is possible power conservator. The size of
the fuselage was also taken into consideration. Inside the cabin, there is ample space to
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comfortably seat nine people and the two required crew members, though the final
layout will ultimately depend upon the customer’s specifications.
Formers, ribs, stringers and longerons made of a titanium aluminum alloy and
composite materials make the XG Endeavour reinforced structure to withstand the
forces experienced during flight.
The two horizontal stabilizer-mounted engines have been designed to meet CFR36 Stage
4 noise requirements. Also, more than ample takeoff thrust allows for a decreased time
to climb which reduces ground signature during the flyover stage of takeoff. After
considering these and many more features, the team has decided upon the present
design for the XG Endeavour.
Through careful cost analysis, and using the Rand DAPCA IV cost model, XG International
estimates that the Research Test Development & Evaluation and Flyaway cost be about
$2.4 billion. With a profit per aircraft around $750,000, the breakeven point is at 20
aircraft, which makes this a very practical business plan for the future. These
estimations include considerations for depreciation, insurance, crew, fuel and
maintenance.
The XG Endeavour has been expertly designed to lead the way into the future of noise,
exhaust, gas, and field length reduction while providing a fast and convenient mode of
air transportation.
- Team XG
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 - III Mission Statement
The XG team strives to develop ambitious advancements in green performance and
technology by fulfilling the promise outlined in our mission statement:
“Develop an environmentally-sensitive aircraft which will
provide our customers with a 21st century transportation
system that combines speed, comfort, and convenience while
exceeding NASA’s N+2 criteria.”iv
An aircraft can damage the environment in many ways; from the consumption of finite
fossil fuels, to harmful emissions, and noise; all of which must be considered in a
complete green project. For the aircraft must be environmentally-sensitive, the
manufacturing details such as material choice and engine details; and the operating
lifespan of the final product must be considered equally.
Though it was expected that the Endeavor would be designed with several elite
performance capabilities in mind. Tight constraints forced several parameters to be
changed from the previous document. The aircraft will still strive to incorporate a
shorter takeoff length than similar sized competitors’ aircraft while still making an effort
to minimize the sacrifice in speed and range.
 - IV Design requirements
The design requirements of XG Endeavour were mainly
guided by the NASA’s N+2 goals. In order to be one of
the main competitors of the business jet field, the
aircraft had to have the green image, and NASA’s
requirements state them very well. A table of NASA’s
N+2 goals for a twin aisle aircraft are copied and
displayed to the right.
The most important design requirement that NASA
proposed is to reduce the fuel consumption of the
aircraft by 40%. It was unlikely that any one component
Figure IV.1: NASA’s N+2 Design Requirements
of any aircraft could make a 40% impact by itself, so
major analysis was conducted to improve efficiency in multiple areas such as reducing
areas responsible for drag, or through weight reduction. Cumulatively, it is possible to
achieve the future standards. This bodes especially true if engine technology will follow
an improvement trend.
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NASA goals also state that the noise emitted during flight has to be reduced by a total of
42 dB. In order to meet this requirement, various changes in the aircraft configurations
must be considered. The main consideration of this was the placement of the engines.
Depending on engine location, noise can be reduced during certain conditions of flight.
For example, when the engines are placed on top of the wing, the sound wave bounce
off the wings and and back into to the sky, which in turn reduces the noise detected by
the sensors located on the ground, which means that the noise emission during takeoff
will be dramatically reduced.
The third requirement was to reduce the NOx emission by 75%. This requirement can be
met by reducing the fuel consumption, and possibly by using a new technology that
may be developed by the year 2020 that will emit the much less NOx.
One of the final NASA requirements set by the N+2 goals stated above was to shorten
takeoff field length by 50%. But because the afore stated NASA N+2 goal are intended to
be applied to large twin isle aircrafts that are limited by their longer take off need, this
particular goal was less practical for a small aircrafts that can already operate in
relatively small airport. Because XG Endeavour is not a large twin aisle aircraft, the
airport field length is short enough to consider this requirement as lower priority in
relationship to the others.
it is clear that all these requirements have at least and indirect relationship to one
another. For example, if the takeoff field length is reduced, the aircraft reduces the rate
of fuel consumption as well as resulting in a reduction of NOx emission. In addition, an
improvement in take off distance enables the aircraft to rise off of the ground in a
shorter period of time, reducing the time of which sampled noise will also be detected.
All things considered, The team has concluded that the foremost priority is to reduce
the fuel consumption, because this will enable the aircraft to emit much less NOx;
reduce the takeoff gross weight; and eventually reduce the takeoff distance.
 - V. Selected “best” aircraft concept
Walk-around
The three engine configuration is designed to provide more fuel efficient flight at cruise
condition, the point at which the most of the fuel is used. The total fuel consumption of
the flight will be reduced by using engines optimized for cruise condition.
The closeable duct, which feeds air to the center engine, will decrease the drag when
the aft fuselage mounted engine is not in use. The engine inlet duct will be deployed out
of fuselage when the engine is in use, and when the engine is not being used, the duct
can fold into the fuselage. Because it does not coincide with the pressurized section of
the fuselage, empty weight added to incorporate this feature will not be significant; the
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team thinks that the benefit from the reduction of drag during the long cruise will be
greater than the increased empty weight to close the ducts.
The 2 engines that are on the horizontal stabilizers will produce 2000 lbs of thrust each,
producing maximum thrust of 4000 lbs during cruise. These engines will operate at the
most efficient throttle setting during cruise to maximize fuel efficiency.
The aft fuselage engine will produce 6800lbs of thrust. This is needed for takeoff and
climb sections of the flight. This engine is designed to be turned off during cruise section
of the flight mission.
The sizing code calculated the static margin of the airplane. It depends heavily on the
location of the main wing because fuel and landing gear will be housed in the wings. The
calculated static margin was 16% for the final design. The CG travel graph was also
generated using the sizing code.
Below (Figure V.1) are two rendered views of the XG Endeavour. The image indicates
examples of the concepts mentioned above.
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Figure V.1: (top) rendered top view of the airplane in flight (bottom) bottom view displaying third engine.
Values of major design parameters
The aircraft needed some of the input variables to start design.
Table VI.1: Design parameter constants
Design
Parameter
Value
t/c


CLmax
AR
Wo /S
TSL/Wo
0.14
28.13°
0.5
1.6
7.8
92.6
0.33
The design parameters for the airplane were set according to historical values or
calculated in the sizing code. Wing loading and thrust-to-weight ratio at the sea level
were found in the current constraint diagrams for the final concept. Aspect ratio for the
concept was chosen based on the size of similar sized business jets. The sweep angle
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was chosen so that the center of lift is located behind the center of gravity. Thickness to
chord ratio was calculated based on the NACA model. In this instance, CLmax is assumed
based on the historical data and the shape of the XG Endeavour.
 - VI. Results of aircraft sizing and carpet plots
Brief description of sizing code
The sizing code was constructed using the Microsoft Excel spreadsheet. Six separate
sheets were combined to calculate the empty weight, fuel weight, static margin, drag,
flight time, and the shape of the concept aircraft. The ‘Main page’ is where all the major
input and output values are displayed. Below is a screen shot of our main page for the
final design.
Figure VI.1: Sample image of the ‘Main’ worksheet of the Sizing Code
The length, diameter and locations of fuselage, nose, and empennage were input first,
then the leading edge angle, taper ratio, aspect ratio of the various lifting surfaces were
input. The area of the main wing was calculated by team the weight of the aircraft, so it
was not an input value in the main page.
The locations of the engines were input. Because there were two separate types of
engines, the calculations for thrust were divided for analysis. The thrust of engine used
in cruise was first calculated then the aft fuselage mounted engine was used to carry
rest of the thrust the aircraft needs to takeoff. With above information, the sizing code
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displayed a rough sketch of the aircraft in a graph format. (See figure above) This
allowed us to quickly notice any major mistakes with location of the major components.
Geometry worksheet is where the aircraft’s general shape is calculated from the input
values that were input in the main worksheet. Three-view of the airplane is displayed to
easily view if any structural conflict exists. Below is an example of our geometry work
sheet.
In the wing geometry calculation both canard and yield; were inputted in case we need
extra lifting or control surfaces: but because we did not home canard # Div/0 appeared
for canard configurations.
Figure VI.2: Sample image of the ‘Geometry’ worksheet of Sizing Code
The sizes of the engines were calculated based on the series of similar turbofan engines
previously used in the business aircrafts. The diameter and the length of the engines
were calculated from the amount of trust each engine needs to generate. Also, some of
the drag coefficients related to the geometry of the aircraft were also calculated in
geometry worksheet, and the values were then used in mission details worksheet for a
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more detailed calculation of the drag. The finalized 3-view of the aircraft was then
transferred to the main page as an output.
The constraint diagram worksheet is generated based on the constraint diagram tools
provided. The input value of the cruise Mach number from the main page was used in
this worksheet to generate 5 constraint lines. These lines were used to calculate the
wing loading and thrust to weight ratio. These values were also copied to the main page.
The thrust to weight ratio used in the finalized concept was 0.36 and wing loading of
92.6 lb/ft2 was used.
Figure VI.3: Sample image of the ‘Weight’ worksheet of the sizing code
The weight worksheet is where empty weight, fuel weight and the total weight of the
aircraft are calculated. The initial guess of total gross weight is used to calculate the
calculated gross weight of the concept. When the guessed total gross weight is different
from the calculated value, the iteration process is used to match the guessed value to
match the calculated value of weight. The equations from Raymer’s Aircraft Design: A
Conceptual Approach Fourth Edition were used to calculate the various components’
weights. Equations for Cargo/ Transport weights were used; however, because these
equations are outdated, validation of the equations was necessary. The validation of the
sizing code will be discussed in the later part of the report. The center of the lift of the
aircraft was also calculated in the weight worksheet. Miscellaneous components
weights were found from the various data sources. For instance, weight of APU,
passenger seats, were found from various manufacturers’ websites and historical data.
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Figure VI.4: Sample image of the ‘Airfoil’ worksheet
The airfoil worksheet was used to define the geometry of the airfoil according to the
NACA 4-digit airfoil series. The original intention was to generate drag polar of the airfoil
within the sizing code; however, due to time constraint, the sizing code was limited to
calculate the thickness to cord ratio and the cross sectional view of chosen airfoil.
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Figure VI.5: Sample image of the ‘mission’ detail of the sizing code.
Mission detail worksheet is where drag, weight, and center of gravity were calculated at
the specific points of the mission. The mission was divided into 11 sections which were
initial, taxi, takeoff, climb, cruise, loiter, attempt to land, climb, divert, loiter, and land.
The atmospheric conditions at different mission altitudes were also calculated. The
velocities were inputted mostly from historical data. The original intention was to find
the best rate of climb for the climb condition. But the calculation part was not
completed hence the inputted value of the climb velocity was used.
Validation & Fudge Factor
For CoDR Report, validation and fudge factor of XG Endeavour were done by inputting
the data of existing aircraft and compare the empty weight. The aircraft that was used
for the validation must have similar features compared to XG Endeavour. To fulfill the
requirements, Bombardier Challenger 300 was used to validate the code. Basic
dimensions such as lengths of the fuselage, height and width of the cabin and other
50+features of the aircraft could be found from Bombardier website. Few specific
dimensions were estimated by using the picture because of the lack of information that
can be obtained.
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After the validation was completed, it was possible to find the fudge factor by using the
following equation,
𝐹𝑢𝑑𝑔𝑒 𝐹𝑎𝑐𝑡𝑜𝑟 =
𝐴𝑐𝑡𝑢𝑎𝑙 𝐸𝑚𝑝𝑡𝑦 𝑊𝑒𝑖𝑔ℎ𝑡 𝑜𝑓 𝐶ℎ𝑎𝑙𝑙𝑒𝑛𝑔𝑒𝑟 300
𝐸𝑠𝑡𝑖𝑚𝑎𝑡𝑒𝑑 𝐸𝑚𝑝𝑡𝑦 𝑊𝑒𝑖𝑔ℎ𝑡 𝑓𝑟𝑜𝑚 𝑡ℎ𝑒 𝑆𝑖𝑧𝑖𝑛𝑔 𝐶𝑜𝑑𝑒
Fudge factor could be inputted in the XG Endeavour to better predict the empty weight.
Once the validation was completed, empty weight based on the sizing code was
predicted 17500lb, while the actual empty weight of the Challenger 300 was 18500lb.
Based on these data, fudge factor of XG Endeavour was calculated using the following
equation
18500
𝐹𝑢𝑑𝑔𝑒 𝐹𝑎𝑐𝑡𝑜𝑟 =
= 1.05
17500
For the reference, specific details of dimensions of the Challenger 300 are shown on the
following page.
Figure VI.6: Code constants worksheet as seen in the sizing code.
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Also, the picture shown below is the comparison between the drawing of Challenger
300 from Bombardier website and drawing from the sizing code.
Figure VI.7: Sample image of the plane sketch from the sizing code
Fixed design parameter values
Drag Prediction
Drag was predicted by using the subsonic drag build-up methods. Three different types
of drags had to be determined in order to determine the drag as a whole: induced drag,
parasite drag, and wave drag. To calculate the estimated parasite drag, all component
drag values were added up by using the equation on the following page (equation VI.1).
# 𝐜𝐨𝐦𝐩
𝐂𝐃𝐏 = ∑𝒊=𝟏
𝑲𝒊 𝑸𝒊 𝑪𝒇𝒊 𝑺𝒘𝒆𝒕𝒊
𝑺𝒓𝒆𝒇
Equation VI.1
where
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Sref
Cf
K
Q
Team 3
Team XG Intl. Inc
is the reference area
is the skin friction coefficient
is the form factor
is the interference factor
The net surface wetted area was found by summing up all the surface areas using the
equations stated in the lecture notes given. In addition, form factor and interference
factor were calculated and assumed respectively. Because the atmospheric conditions
are different due to the different altitude the aircraft flies at, the drag was calculated for
each of the mission segments. The drag was calculated for each mission segments
separately because atmospheric conditions vary due to altitude. Below is the table
containing the values of drag calculated for each of the segments of the flight mission:
Table VI.2. Values of Drag for the aircraft designs
Taxi
Take off
Climb
Cruise
Loiter
Attempt to Land
Climb
Divert
Loiter
Land
Total CD
0.158
0.051
0.070
0.067
0.053
0.047
0.049
0.059
0.111
0.086
Drag
3220
6381
3397
3709
5273
8049
7239
4366
2753
2528
Results of engine modeling
Engines were selected then analyzed for the project. Details on the actual engine
selection can be found in the later ‘XI. Propulsion’ section of the report. Some details for
the two different engine sizes were identified.
The 2000 pound thrust model was not scaled because data is already available for this
model. The 2000 pound thrust model has a bypass ratio of 2.9, takeoff thrust of 2050
pounds, and a compressor pressure ratio of 24. In order to calculate the data for the
6800 pound thrust model, an Excel sizing routine was used to scale the engine for the
appropriate thrust level. The 6800 pound thrust model has a bypass ratio of 3.2, can
produce 6800 pounds of thrust, and has a compressor pressure ratio of 26. In order to
scale the engine, several assumptions were made. First, technological improvement
factors were taken into consideration to determine what engine performance would be
like in the year 2020. Since the 2000 pound model was already the correct size, no
efficiencies were needed and the larger engine was simply scaled from the smaller one.
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Several sources were used to generate thrust available and thrust required for the flight
of the final designs. These plots show velocities will be attainable in certain
configurations at specific altitudes. In order to calculate the thrust available, a basic
cycle analysis with the atmospheric conditions at various altitudes were inputted. The
results of the calculations are shown below for takeoff, several altitudes that a typical
mission climbs through, and for the cruise condition:
12,000
10,000
Thrust (lb)
8,000
6,000
4,000
Thrust Available
Thrust Required
2,000
0
0
20
40
60
80
100
120
140
Velocity (mph)
Figure VI.9: Thrust vs. velocity at take off.
The takeoff diagram shows that thrust available exceeds thrust required up to the
takeoff speed of 120 mph. Therefore the propulsion system is sufficient for takeoff.
12000
10000
Thrust (lb)
8000
6000
4000
Thrust Available
2000
Thrust Required
0
0
100
200
300
400
500
600
Velocity (mph)
Figure VI.10: Thrust vs. velocity at 15,000 ft segment climb.
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12000
10000
Thrust (lb)
8000
6000
4000
Thrust Available
Thrust Required
2000
0
0
100
200
300
400
500
600
Velocity (mph)
Figure VI.11: Thrust vs. velocity at 25,000 ft segment climb.
12000
10000
Thrust (lb)
8000
6000
4000
Thrust Available
2000
Thrust Required
0
0
100
200
300
400
500
600
Velocity (mph)
Figure VI.12: Thrust vs. velocity at 35,000 ft segment climb.
.
The above three diagrams are for the climbing configuration at three different altitudes.
All show that thrust available exceeds thrust required up to about 480 mph. Therefore,
the maximum climb speed is approximately 480 mph.
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12000
10000
Thrust (lb)
8000
Thrust Available
Thrust Required
6000
Thrust Available (2 engines)
4000
2000
0
0
200
400
600
800
Velocity (mph)
Figure VI.13: Thrust vs. velocity at 45,000 cruise.
At the cruise condition, thrust available always exceeds thrust required with three
engines running. After shutting down the big engine to conserve fuel, the thrust
available drops significantly. With two engines running there is enough thrust available
to enable the aircraft to travel up to 530 mph which exceeds the design cruise speed of
484 mph.
Mission modeling
55000
35000
15000
-5000
-500
0
Mission
Altitude(ft)
Accumulative Distance(nmi)
A/C Weight(lb)
Fuel Weight(lb)
500
Taxi
0
0.758
29093
9837
1000
Takeoff
10000
2.758
28945
9689
1500
Climb
40000
32.76
26544
7288
2000
Cruise
45000
3443
26463
7207
2500
Loiter
16000
3463
26247
6991
3000
Attempt
to Land
2000
3464
26142
6886
3500
Climb
10000
3480
26003
6747
4000
Divert
40000
3680
25958
6702
Loiter
16000
3700
25925
6669
4500
Land
2000
3700
23732
4476
The aircraft will take off at sea level, and climb to cruise altitude of 40,000 ft. when it
reaches 40,000 ft, aft fuselage engine will shut down to save fuel. The thrust will be
sufficient with 2 smaller engines mounted on horizontal stabilizer for the cruise segment
of the mission. As the aircraft reaches the vicinity of the destination, it will descent to
loitering altitude, then attempt to land. In case of possible situation when aircraft
cannot land at the primary airport, aircraft will climb and cruise to the secondary airport
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with most efficient fashion. The aircraft is designed for maximum range of 3700 nmi
with maximum payload of 9 passengers and their baggage.
 - VII. Major design trade-offs
Trade off overview
Table VII.1: Concept options (overhead view)
Previous concepts
Final settled design
The most important design requirement that NASA proposed is to reduce the fuel
consumption of the aircraft by 40%. This requirement forced the team to work very hard
in changing various configurations from modern aircraft. The group thought of some
possible methods to save fuel efficiency. The hybrid engine configuration is one of the
few design ideas of which the team explored during this process. In the SDR report, the
team considered applying one turbofan engine and two turbo prop engines, turbofan
being used during takeoff and turboprops for cruise. Ultimately, this idea was not used
due to its significant loss in loss of the cruise speed and little significant gain in fuel
efficiency. Instead, three turbofan engines of different sizes were implemented in the
final design. Two smaller engines will be used during cruise segment of the mission, and
one bigger engine will be used for takeoff. In this way, fuel consumption may be
reduced by approximately 25% to 30% even with today’s technology. In addition, with
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the constantly advancing and improving trend of engine technology, the team is
confident that 40% reduction of fuel consumption can be achieved by 2020.
Turbo prop concept was under consideration until the last minute of the design process.
It was very attractive concept because of its fuel efficiency compared to the turbofan
engines; however, because of its slow cruise speed of Mach 0.6, it will not be able to
perform as fast as competition aircrafts. Also the fuel saving will not be great compared
to other turbo prop engine mounted engine aircraft such as Beechcraft Kingair 350i.
The unducted fan engine idea was not used in the final concept because of increase of
noise. The supersonically spinning fan blades will create shock wave which will increase
the noise and possible exceed the minimum noise requirement. Also because there
were not enough data we were able to obtain, enough engine specification data,
unducted fan engine was less attractive despite the fuel efficiency.
The issue of not being able to turn the engine after once it is shut off has to be
addressed. However, by 2020, the engines will be reliable so that engine can be shut
one and off as it is needed. We believe it will be able to operate similar to the gasoline
engines in the hybrid cars where gasoline engines are not turned on while battery
powered motor is running.
The NOx emission will be decreased from less fuel used during the mission. NOx is a
byproduct of combustion of fossil fuel. Therefore, with close to 40% fuel reduction, NOx
will be successfully reduced. Previously we were considering employing catalytic
reduction technology on aircraft for the first time. However, because of weight increase
and reduction of thrust, the concept of employing catalytic reduction was not
implemented in the final concept.
The emergency door was removed in the final concept because, for the size of the
aircraft, emergency door is unnecessary and will increase structural weight needed for
reinforcing the pressurized cabin.
The crucifix form horizontal stabilizer will have small engines mounted on the because
of empennage engine. T-tail was used for one of the concepts for the previous report;
however, because we have engines mounted on the horizontal stabilizers, T-tail
mounted engines would create nose pitching down moment and increase trim drag
during the cruise.
Canard also was used for concept design, however because we believe aircraft will be
successfully controlled with current size of horizontal stabilizer, extra canard control
surface will increase empty weight of the aircraft as well as drag during the flight.
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Summary of carpet plot studies
The carpet plot was another way to represent the constraints of XG Endeavour. Based
on the current constraint diagram and the weight prediction from the sizing code, it was
possible to make different plots and construct the carpet plot.
Considering the important features of XG Endeavour, there were four main constraint
plots:
Estimated Gross Weight vs. Wing loading at;
a) Different thrust to weight ratio at the sea level
b) 2g maneuver
c) Takeoff ground roll
d) Landing ground roll.
For the Estimated Gross weight vs. Wing Loading, thrust to weight ratio at the sea level
and the wing loading were 0.36 and 92.625, respectively. In order to make a carpet plot,
both thrust to weight ratio at the sea level and wing loading varied by around 20%.
For easier calculation, thrust to weight ratio at the sea level (T/W) was estimated at 0.4,
while wing loading was estimated at 90.
The final result of carpet plot of XG Endeavour is shown below.
FigureVII.1: Gross Weight vs. Wing Loading based on various thrust to weight ratio at the sea level.
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For the next step, estimated weights at takeoff ground roll and landing ground roll were
calculated. These calculations were made from first taking exponential equations based
on the takeoff ground roll of 4500ft and landing round roll of 2500ft, then followed by
using those equations to substitute the values of wing loading. A new graph was
generated by plotting these new results with the with the previous carpet plot above.
The ultimate carpet plot of XG Endeavour is shown below.
FigureVII.2: Ultimate Carpet Plot of XG Endeavour.
From the carpet plot above, it was possible to ensure that estimated gross weight
calculation was reasonable. Also, it is possible to conduct the off design.
 - VIII. Aircraft description
Dimensioned three-view, to scale
Below is the 3-view documentation of the XG Endeavor modeled generated in Catia.
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FigureVIII.1: Scaled 3 views diagram of the XG aircraft with retractable ducts open.
The exact values of surface location and center of gravity can be found in the sizing code
section of the report. The following figures are just brief overviews of the sized
representation of the airplane
FigureVIII.2: Overhead dimensions of the XG Endeavour
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FigureVIII.3: Front view dimensions of the XG Endeavour.
Representative internal layout
To maximize cabin headroom, the floor of the cabin
was lowered (See figure right). In order to contain all
the equipment required for a comfortable flight; spaces
in the nose of the craft, the aft of the craft and the
wings had to be optimized. On the following page are
the available images representing space availability.
FigureVIII.4: View of the cabin
cross section to demonstrate
depressed floor.
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Not shown:
The fuel tank extends to the small
area under the cabin.
FigureVIII.5: Representational internal lay out overhead.
Engines had to be supplied with both fuel and a steady
free stream of air when operated. For all three engines,
fuel would be pumped from the fuel reserves in the wings,
tail and lower aft storage of the cabin into the engines.
Those engines mounted on the tail were located in
positions with uncontested access to air. The center
engine, however, had to have a duct designed to feed the
air. To create laminar flow, the ducts had to be sized and
contoured appropriately. Raymer (pg 458) has an example
of a duct for an air-fed breathing engine as well as explains
the methodology and requirements of duct sizing. Because
the duct demands specific dimensions, restrictions for the useable space are then
imposed (see left). In Figure VIII.6, The image shows the aft engine, the air duct, area
for fuel reserve and fuel pump, and area for other components in red, green, salmon
and purple respectively. Careful arrangement of all other equipment (APU, air
conditioner, motor to close and open air ducts, etc.) must be fitted within these
constraints.
FigureVIII.6: Generated 3D view
of the available space in the aft
of the aircraft.
In previous business modeling reports, it was stated that the internal layout of the
aircraft was customizable. As a result, the internal cabin layout could be altered,
provided that the center of gravity of the entire internal layout remained between 22-24
ft from the nose of the craft and within the weight tolerance of 300-350 lbs (See
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Appendix B: the ‘Center of gravity’ section ‘Sizing code’ for details). Below is the layout
of the aircraft test case. In this case, the moment of each individual seats and other
equipment for validation. For the test case, the cabin center of gravity was located
22.75 ft from the nose of the craft, and the weight of the interior totaled 306 lb.
Below is the floor plan of the cabin interior generated in Catia for the test case.
Figure VIII.7: Cabin floor plan used for project analysis. Dimensions are in feet.
As a professional business jet, the number of expected passengers falls within the range
of 6-10 people. Above, in figure 6, is a possible 8 passenger, 1 crew floor plan of the
cabin layout. During taxi and take off, the lavatory seat can double over as seat.
The lavatory location and exit locations are fixed and cannot be moved, but any other
aspect of the system is changeable.
 - IX. Aerodynamic design details / justification
The airfoil was selected from historical data. Although it is not the optimized airfoil
selection, the aircraft was designed with NACA 2414 airfoil. Technical data of NACA 2414
airfoil is attached in the drag polar section below. During the takeoff and landing, the
aircraft requires CL= 1.6 which is more than NACA 2414 can produce. Therefore, high
lifting device such as flaps and slats are considered to achieve required CL.
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Figure IX.1: High lifting devices
Drag Polar
Drag polar is a way to represent the coefficient of lift and coefficient of drag of the
airplane. For XG Endeavour, NACA 2414 was used for the main airfoil and the drag polar
is shown below.
Figure IX.2. Drag Polar for NACA 2414.
Compare to XG Endeavour’s Reynolds number, which is around 1.02e6, purple line of
the drag polar was used for the drag polar of XG Endeavour. For the landing, parasite
drag and induced drag at the landing were added to the cruise drag, while for the take
off, parasite drag and induced drag at the takeoff were added to the cruise drag.
Resulting drag polar at takeoff, cruise and landing are shown below.
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Drag Polar
1.5
Coefficient of Lift
1
Takeoff
0.5
Cruise
0
0
0.02
0.04
0.06
0.08
0.1
0.12
0.14
Landing
-0.5
-1
Coefficient of Drag
Figure IX.3. Drag Polar of XG Endeavour
 - X. Performance
A very important aspect of an aircraft’s performance is the aircraft’s maneuverability at
various flight speeds, and the aerodynamic loading that is applied to the aircraft during
various points in the design mission. The aircraft’s abilities and limitations regarding its
performance provide for a set of parameters that comprise what is known as the flight
envelope of an aircraft. One technique that allows for a better understanding of the
aerodynamic loading on the aircraft is the V-n (also known as V-g) diagram. The V-n
diagram better illustrates the performance characteristics of an aircraft during various
different maneuvers, and different flight speeds and aircraft orientations. A standard
V-n diagram of an unknown aircraft is shown below in Figure X.1.
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Figure X.1: A standard V-n (or V-g) diagram is shown to provide a visual understanding of aircraft loading at various
flight speeds.
The V-n diagram reveals much about an aircraft. The region between the red lines,
including the green and yellow ranges, is a typical flight envelope. The envelope is
bounded on the left by a pair of curves. These curves constitute the stall velocities at
various load factors. At these speeds, any load factor above the corresponding value is
aerodynamically unavailable. The point marked maneuvering speed is the speed at
which the aircraft may perform a maneuver that results in the maximum possible load
factor. Any maneuvers made to the left of the curve result in a stall of the aircraft. This
holds true for both the upper and lower sections of the curves that make up the left
bound of the envelope. The envelope is bounded at the upper limit of the envelope by
the maximum positive load factor (which in this case, is approximately 4.5g, according
to the figure). It follows then, that the lower bound of the flight envelope is the
negative maximum load factor (in this case, nearly -2g). Between these load factors, the
stall speeds, and the maximum structural cruise speed is the normal operating range.
Unless the aircraft comes under extreme aerodynamic duress for one reason or another,
the entire flight design mission should be contained within the normal operating range.
If the aircraft’s maneuvering results in a load factor that exceeds either the positive or
negative limit shown in the V-n diagram, structural damage is possible. Subsequently,
further disregard of the load factor limits and maneuvers that exceed the upper and
lower bounds of the acceptable operating ranges will eventually lead to structural
failure in one or more critical components of the aircraft, although the flight crew and
passengers will most likely experience black-out, or red-out before structural failure
occurs, depending on whether the upper or lower limit is exceeded. So, exceeding
these parameters may lead to the loss of the aircraft, or even human life, for a number
of reasons. This is what makes the V-n diagram and its parameters so important. The
right side of the flight envelope is bounded by the maximum flight speed of the aircraft,
which may also be considered the aircraft’s dive speed. Even steady level flight at
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speeds greater than the maximum structural cruise speed but less than the dive speed is
not advised. Flight at these speeds cannot be sustained for extended periods of time
without eventual structural damage. When the aircraft exceeds the right-most limit of
the flight envelope, structural failure will occur. It is crucial that the aircraft not exceed
any of the boundaries of a V-n diagram.
Figure X.2 shows the V-n diagram
for Team XG’s flagship business jet.
V-n Diagram
4n
The V-n diagram allows for several
performance specifications to be
3
illustrated, and allow for a visual
2
interpretation of the aircraft’s
limitations. For steady level flight
1
at exactly 1g, the aircraft’s normal
V
0
stall speed occurs at 220 ft/s. At
0
200
400
600
800
1000 1200
the upper limit of the stall curve,
-1
the maneuvering speed can be
-2
determined from the graph.
Maneuvers creating a maximum
load factor of +3.333g may be Figure X.2: Team XG's V-n diagram is shown to allow for the
performed once the aircraft visualization of the aircraft's flight envelope.
reaches the maneuvering speed of approximately 400 ft/s. The stall curve intersects the
lower bound at the same speed as the normal stall speed, as the maximum negative
load factor of the aircraft is -1.000g. Between these curves and our design cruise speed
of 710 ft/s is the normal operating range for the Team XG aircraft. Between the upper
and lower maximum load factor limits, the design cruise speed, and the aircraft’s
maximum dive speed of 1020 ft/s is the caution range for our aircraft.
These values were determined using a number of other predetermined performance
specifications. Some of those specifications and the aforementioned flight
characteristics determined from the V-n diagram are tabulated in Table X.1. In Table
X.1, the aircraft’s normal stall speed, dive speed, cruise speed, and maneuvering speed
(denoted as the stall speed at maximum positive load factor in Table X.1) are shown,
courtesy of the V-n diagram. However, Table X.1 also provides the take off distance,
landing distance, best endurance velocity, and best range of our aircraft. The takeoff
and landing distances of 4000 ft and 2500 ft, respectively, are important to note as they
will determine which airports our aircraft will be able to access. Also very important to
note are the best range and best endurance velocity.
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Table X.1: Performance specifications for Team XG's business/personal aircraft are
tabulated to provide clarity.
Performance Specifications
Takeoff Distance
Landing Distance
Best Range
Best Endurance Velocity
Cruise Speed
Dive Speed
Normal Stall Speed
Stall Speed at Maximum (+) Load
Factor
Values (units)
4000 (ft)
2500 (ft)
3700 (nmi)
710 (ft/s)
710 (ft/s)
1020 (ft/s)
220 (ft/s)
400 (ft/s)
The best range of 3700 nmi for the aircraft is, coincidentally, the maximum design
mission range of our aircraft, which can be achieved at the best endurance velocity.
Similarly, the best endurance velocity is our aircraft’s cruise speed. The best endurance
velocity of 710 ft/s was chosen for the design cruise speed for the obvious reason that it
is the most fuel efficient cruise speed, and at the same time, a timely arrival at our
customers’ destination is also ensured. These parameters were some of the focal points
of Team XG’s performance goals.
 - XI. Propulsion
The final design utilizes three turbofan engines. Two of these turbofans can produce up
to 2000 pounds of thrust, while the remaining one can produce up to 6800 pounds of
thrust. To incorporate these engines into the final design, they were modeled after an
existing turbofan engine. The baseline engine was the HF 120 turbofan, which is
currently manufactured by GE Honda Aero Engines. Below are a picture and a
schematic of the 2000 pound thrust version of the engine and the dimensions of that
model:
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Figure XI.1: (left) Photograph of the GE Honda Aero Engine, the HF 120 turbofan and (right) a simple schematic of
the same engine
What makes this engine stand out as our final engine choice is environmentally friendly
design which minimizes noise and emissions. In addition, it is very fuel efficient sporting
a low SFC, and it has a very high thrust to weight ratio.
It was very important to generate thrust available and thrust required versus velocity at
various altitudes for the final design. These plots are crucial because they show what
velocities will be attainable in certain configurations at specific altitudes. The data for
these plots was generated from several different places. First, the thrust required was
available from the team’s sizing code. In order to get the thrust available, some
calculations were necessary. Using a basic cycle analysis with the atmospheric
conditions at various altitudes inputted, it was simple to calculate the maximum thrust
the engines could produce at a given velocity. The numerical results of the analysis are
shown in the “VI. Sizing Code; results of engine modeling” section of the report.
Overall, the propulsion system for the aircraft is very effective. It meets the thrust
required criteria at all altitudes and configurations. Stability and control-related
propulsion considerations, including emergency conditions, are found in section XIV
below. Also, the three engine configuration allows for unnecessary thrust to be cut out
while cruising therefore saving fuel. The propulsion system also meets all noise and
emissions requirements.
 - XII Structures
Structural integrity is a major design factor in aircraft construction. But before the
structure can be designed, one must determine the types of loads that would be
imposed on the aircraft. In general, each part of the aircraft is subject to many different
loads. In the final design structure, there might be thousands of loading conditions. Of
them, hundreds could prove to be critical for parts of the airplane.{cite 1}
Generally, the strain allowances on aircrafts are defined by load factor, n. In simple
terms, load factor is defined as the component of force on a structure divided by the
aerodynamic weight. The factor of safety, which is the ultimate load over the intended
load, is also a crucial consideration in aircraft design.
The FAA establishes two kinds of load conditions:{Site 1}

Limit loads are the maximum loads expected in service. FAR Part 25
(and most other regulations) specifies that there be no permanent
deformation of the structure at limit load.
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Ultimate loads are defined as the limit loads times a safety factor. In
Part 25 the safety factor is specified as 1.5. For some research or
military aircraft the safety factor is as low as 1.20, while composite
sailplane manufacturers may use 1.75. The structure must be able to
withstand the ultimate load for at least 3 seconds without failure.
In addition to these strength factors, designers have to also deal with fail-safety, fatigue,
corrosion, maintenance and inspect-ability, as well as production value. Modern
airplane structures are usually designed as load carrying shells that are reinforced by a
series of frames, formers, longerons, and other ridged shapes that make up the airplane
skeleton. Skin stringers, spars, ribs and boxes then enforce the thin airplane skin.
To the right is a diagram of how
a wing’s structural support can
be
arranged
to
contain
components necessary for the
aircrafts operation. Making the
most of negative space in
airplane control surfaces is
crucial to increase aerodynamic
efficiency as well as allow more
room for additional freedom.
Below (Figure XII.2) is an
example of a typical airplane
airframe. Notice that the Figure XII.1: Overhead view of a wing showing ribs, spars,
windows, the doors and the cavities for appliance storage.
cockpit glass have a different method of re-enforcement than the rest of the craft.
and
Figure XII.2: The structural skeleton of a similar type aircraft
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A typical real world airplane will have several areas of which calculations are crucial. For
the sake of the design project, however, these calculations will be generalized.
Configuration layout and highlights
In order to maintain a safe and light weight aircraft, the airplane’s skeletal structure, or
airframe, has to be built with the loads in consideration. Such loads not only include the
internal mass forces (such as payload, fuel load, equipment and passengers etc.) but
also include flight forces, and ground forces. Ground forces induced by taxiing and
landing can be easily estimated, given the airplane’s weight and performance
parameters. Flight forces, such as propulsion thrust, lift, drag maneuver and wind can be
mathematically estimated and tested digitally or in a wind tunnel.
Important Load Paths
Because very few details are known about the particular structural components of the
Endeavour, a general sketch was created to visualize the airframe of the craft. Below are
the estimated locations of the ground forces on the aircraft from an overhead view.
Figure XII.3: Force distribution sketch and estimated locations of major stress factors.
Here, ‘areas of concern’ are marked. These areas were considered important to focus on
because of their unusual load situation. Moments formed on the joints or connectors of
the wings and stabilizing surfaces of the aircrafts (they are marked with tan hatching)
must be addressed through re-enforcement of the main bulkheads at the wing roots
and areas like the joints for the landing gear, where unusually high of pressure is
applied, has to be resting on framework that is design to properly distribute the entire
craft’s weight. In general, moment M is a function of force, and distance. For a
cantilever, the moment created on the attached
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𝑀 = ∫ 𝐹 ′ (𝑥)𝑑𝑥
Equation XII.1
Where F is the cross product of force per unit length, or, F(x) = force x distance. When
designing locations where airplane features intersect, this equation is invaluable.
With these forces in consideration, a skeleton of the major ribs, stringers, formers and
frames was then sketched. The results of the analysis for the theoretical load paths and
the locations of the applied forces have been overlaid the aircraft diagram below.
Figure XII.4: Major paths of load and the location of crucial formers, ribs, longerons and strings
To further reinforce the airframe, other frames and formers are needed. Ideally the
entire plane will be manufactured by the lightest and strongest, durable material such
as composite structure. Below is the example of the framework of the aircraft.
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Figure XII.5: Generated 3D model of the formers and stringers as they would be located on the actual craft
Wing-fuselage intersection
The aircraft wings are mounted under the fuselage and secured through two of the main
supporting formers. Mounting the wings low on the aircraft yields several benefits. One
such benefit is the ability to reinforce the joints connecting the wing to the fuselage
without disrupting the pressurized portion
of the aircraft. The structures must be
strong enough to hold the wings and the
aircraft together, despite the forces
generated by the lift of the wing at the wing
root.
Most aircraft built in this formation today
use a carry-through method of securing the
craft to the airplane similar to that shown in
the image to the right (Figure XII.6).
Figure XII.6: Cross section demonstrating how a low
wing can be affixed to the aircraft.
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Engine mounts and landing gear integration
Because of the locations of the engines, creative measures were made when addressing
the engine mount options. As stated in the previous section of the paper, the engine
collection currently consists of three separate turbo-fan engines. There is one central
turbofan engine which can produce 6800 lbf of thrust, and two cruise fans that can
produce a thrust of 2000 lbf.
The two smaller engines are mounted through the horizontal stabilizers. Options on
how an engine can be mounted through the control surface are demonstrated in the
image below.
Figure XII.7: Side view of the horizontal stabilizers with engines.
With both, the engines are mounted relatively close to the fuselage to reduce moment.
However, unless mounted exactly on the centerline, moment about the c.g. can never
be completely eliminated. Further analysis would be needed to determine the most
efficient method to keep the engine secure. When more details (such as control surface
sizing requirements) of the aircraft become available a more definitive decision will be
made between the two options. General analysis tactics can give an estimation of the
structural stability required to resist the force caused by the thrust of the engine. Below
is an example of a simple truss-type structural option of re-enforcing the stabilizers of
the tails with the locations of the engines currently consideration.
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Figure XII.8: A simple single plane truss formation
The center engine will be mounted using ring supports to allow for modular operation
and maintenance. The engine flow duct must be accommodated for. In Raymer (pg 315),
dimensions of the duct are stated, the structure then has to work around the required
lengths for laminar airflow.
Similarly, the structural requirements of the landing gear will be measured. To reduce
moment, increase stability and increase the use of negative fuselage space, inwardlyretracting wheels (i.e. toward the centerline of the aircraft) were used. By having wheels
that swing outward during landing (like that shown in the image below, far left) you
increase the yaw stability of the aircraft while taxing by creating a stable triangular base.
FigureXII.9: examples of inwardly retracting landing gear (right-hand figure not to scale)
The rear landing gear is located on the intersection of a main wing and a fuselage. The
stringer itself is slightly located on the frame of the aircraft of which the empty weight
center of gravity is. This way, the center of gravity is forward of the rear landing gear at
all times. Below are some example diagrams of the situation and methodology on how
to solve for their location.
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Figure XII.10: (top) Wheel loading geometry taken from Raymer page, (bottom left) side view and (bottom right)
bottom view landing gear locations in relationship to the center of gravity.
Materials
Historically, airplane airframes have been designed from materials chosen for their
strength. But as material production evolves, it becomes more safe and affordable to
make modern craft out of materials with several other properties in mind such as
memory-shape or thermal-expansion capability. When making material choices, it is
important to pay attention to more than just the strength to weight ratio. On-board
materials often have unique applicability, and designers will often select them based on
their:
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




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Fracture toughness
Crack propagation rate
Notch sensitivity
Stress corrosion resistance
Exfoliation corrosion resistance
Gamma Titanium Aluminides (GTA), a new class of lightweight alloys has been in
development over the past few decades, and are said to sustain temperatures higher
than 600oC. According to InnovaTiAl, a European commissioned project which addresses
“NanoTechnologies and nano-sciences, … materials and new production process…”,
gamma titanium aluminides have proven to meet the requirements for aeroengines and
could easily be applied for use in airfoils, wheels, pylons and other types of supports.
GE’s announcement in 2005 to use intermetallics in the GEnx engine underscores the
appeal of such materials. Although the cost to implement such materials have been a
major hurdle for existing aircrafts, it is likely that this material will be common in crafts
of the future, including the Endeavour.
Other materials that would prove useful for the future of this craft include the more
common composites. Among the many materials available, below are a few possible
options and their use.
Table XII.1: Material selections for structural use.
Material type/name
Fiber Glass
Composites
Thermoplastics
Aluminum based alloys
Use
Control surfaces such as wings and ailerons and rudders
Wings and fuselage skin.
Rib and frame structures in the craft
Engine composition and general supporting structures.
Currently, the design team is weighing heavily on the idea of using aluminum-based
alloys, particularly the GTA type materials described above. Currently, making the entire
aircraft from GTA material is impractical largely because of the manufacturing cost and
available data. If, in the future, the availability and practicality of using the GTA type
materials becomes feasible, then the option to make a more considerable portion of the
superstructure of the aircraft from this material will be reevaluated.
 - XIII. Weights and balance
The estimation of the weight of a conceptual design aircraft is a critical part of the
design process. The group used the sizing code and equations in Chapter 15 in Raymer’s
book to find the best weight of the major components. The aircraft was divided into 3
major sections: fixed structures, inner structures, and payloads. Fixed and inner
structures were catalogued as structures that do not move or change weight during the
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flight. Payload and fuel were of major concern for computing c.g. travel information.
The summation of these groups provides the total empty weight. The fixed components
section contains the wing, horizontal tail, vertical tail, fuselage, main landing gear, and
nose landing gear. The total weight of the fixed components is 13,125 lb. The inner
structures group contains avionics, seats, and compartmental fixtures; it weighs 660 lb.
The payload group consists of passengers, crew, and luggage. The total weight of the
payload group is 2,223 lb. Total empty weight is generated by summing the weight of
structures, propulsion, and equipment groups; it sums to 16,800 lb, which is comparably
low to other current-generation aircraft of similar size including the Gulfstream G250,
at 24,150 lb, and the Bombardier Challenger 300, at23,349 lb. The gross weight of the
aircraft is 29,800 lb.
Overview discussion on prediction methods
The weight was calculated using 22 equations with 90 constants in Raymer to compute
the best estimate of the empty weight. The component weight equations are put into
the sizing code and provide an estimated empty weight.
Fixed components
Table XIII.1: Component chart of location, weight and moment of each component
Name
#
XLocation(ft)
Weight(lb)
Moment From
Zero Point
Wwing
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
2
2
33.141
40.725
39.809
44.5
36.141
11
43
43
43
35.141
11
43
11
33.141
33.141
7.3333
33.141
40.725
1157.3
465.71
379.68
4149.4
798.17
214.15
0
40.4
84.658
283.58
456.01
198
105.51
102.64
586.26
1155
269.55
211.33
38355
18966
15115
184649
28847
2355.7
0
1737.2
3640.3
9965.4
5016.2
8514
1160.6
3401.5
19429
8470.2
8933.4
8606.4
Whorizontaltail
Wverticaltail
Wfuselage
Wmain landing gear
Wnose landing gear
Wnacelle group
Wengine controls
Wstarter(pneumatic)
Wfuel system
Wflight controls
Wapu installed
Winstruments
Whydraulics
Welectrical
Wavionics
Wfurnishing
Wairconditioning
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Payload
Inner Structure
Conceptual Design Review
May 7, 2010
Total
Wanti-ice
Whandling gear
Wsolarfilm
Wengine
Avionics
Seat1
Seat2
Seat3
Seat4
Seat5
Seat6
Sofa
Laboratory
Galley
Environmental
Control
Pilot
Passanger1
Passanger2
Passanger3
Passanger4
Passanger5
Passanger6
Passanger7
Passanger8
Passanger9
Luggage
Team 3
Team XG Intl. Inc
2
2
1
3
1
1
1
1
1
1
1
1
1
1
1
33.141
35.141
44.5
43
20
16.7
17.7
25.46
30.335
15.63
30.335
21.65
33.9
13
59.612
8.9418
300
1861.5
234
34
34
34
34
34
34
34
34
34
1975.6
314.23
13350
80044
4680
567.8
601.8
865.64
1031.4
531.42
1031.4
736.1
1152.6
442
20
30
600
2
1
1
1
1
1
1
1
1
1
2
5
16.7
17.7
25.46
30.335
15.63
30.335
21.65
21.65
21.65
11.7
32.312
234
195
195
195
195
195
195
195
195
195
234
15680
1170
3256.5
3451.5
4964.7
5915.3
3047.9
5915.3
4221.8
4221.8
4221.8
2737.8
518210
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Team XG Intl. Inc
Center of gravity
Location (ft)
Weight (lb)
Figure XIII.1: X-location CG travel diagram plotted in relationship to the airplane sketch
Above is the visual representation of the center of gravity, the location of which changes
according to fuel consumption, as well as payload and crew allocation. Since the
location of center of gravity (c.g.) shifts during different stages of the mission, it was
important to mark the path of the variations, known as the CG travel diagram (Figure
XIII.1). As shown the figure above, the location of center of gravity ranges from 32.6 to
34.5 ft from the nose. This travel distance was calculated by locating various weights of
components, which is important to know in order to ensure that the center of gravity is
always in front of the neutral point when in flight.
Because the XG Endeavour engines are placed so far the aft of the aircraft, the initial
location of center of gravity is placed further from the nose of the aircraft than on
convention jets of similar sizes with a more traditional configuration. In this case, the
center of gravity was calculated to be 32 feet from the nose. The initial location of the
c.g was calculated by calculating the moment created by various components.
To simplify the problem, the aircraft was divided into 3 major sections, as implied in
table XIII.1:
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


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Team XG Intl. Inc
Fixed structures
Inner structures
Payloads
Fixed and inner structures were structures that do not move or change weight during
the flight. Payload and fuel were major concern for computing center of gravity travel
information. The weight of the aircraft will change constantly during a mission and so
important points were identified to find location of center of gravity.
The mission was divided into 12 major segments:
 Take off
 Gear up
 Fuselage tank use
 Wing tank use
 Gear down
 Land
 Reserve fuel
 Passenger disembark
 Crew disembark
 Add fuel
 Crew boarding
 Passenger boarding
During the design mission, the payload and fuel weight changes. First, the initial center
of gravity was located at 32 feet from the nose, when the aircraft was at its maximum
gross weight of 29000 lb. During the take off, the fraction of fuel is used from the tank
installed in fuselage under the cabin. During the climb and cruise, more of the fuel
stored in the wing box is consumed. Because these fuel locations are behind the initial
location of the center of gravity, the c.g. will travel forward as the weight of the aircraft
decreases. Therefore, as the travel progresses the aircraft becomes more stable.
Because more stability is advantageous during the cruise, this situation is very favorable.
With more stability, the passengers will then experience a smoother, more comfortable
ride. When passengers and pilot disembark the weight is at its minimum, with only
trapped fuel and oil. At this point, the weight of the aircraft is approximately 21500 lb.
The fuel is added, and passengers and crew board for the next flight.
Below are numerical representations of the X-location of the Center of gravity, and the
total weight of the aircraft and its contents.
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Table XIII.2: The values of CG, and Weight for separate segements of the mission
Wight Segment
Take off
Gear Up
Fuselage Tank
Wing Tank
Gear Down
Land
Reserve Fuel
Passenger Off
Crew Off
Add Fuel
Add Crew
Add passenger
CG (ft), from the nose
32.1
32.1
32.0
32.0
32.0
32.0
31.9
33.0
33.4
33.4
33.0
32.1
Weight (lb)
29163
29011
26468
25951
25916
25916
23669
21680
21446
26940
27174
29163
 - XIV. Stability and control
The basic idea behind providing static stability for the aircraft is to come up with a
design which keeps it from yawing, rolling, or changing pitch too easily as forces from
the surrounding fluid.
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Figure XIV.1: Static stability diagrams for equilibrium flight, unstable flight and neutral static stability.
Much of this has to do with the plane’s static margin, which is a percentage of the
distance between c.g. and neutral point versus the mean aerodynamic chord (6.6 ft)
that tells how far back the neutral point is from the center gravity (32 ft down from the
nose). Historically, we found that the static margin is between 4 and 16 percent. Lower
static margins makes the plane more responsive to the pilot’s control, while higher
static margins result in a plane which is more stable, but slow to respond.
For low altitude flights (during takeoff and landing) responsive controls are desired, and
for cruise a comfortable, ‘smooth ride’, more stability is wanted. Because of this, a static
margin of 18% was chosen, thus putting the neutral point 1.6 ft behind the center of
gravity.
Figure XIV.2: Wing and horizontal stabilizers diagram and areas of concern
To size the elevators, we used conventional business jet values found in Dan Raymer’s
textbook. Each elevator is 90% of the tail span and 32% of the tail chord; thus, each
elevator has a chord length of 1.43 ft and a span of 10 ft. this gives them a planform
area 14.3 ft2. We were also able to find trim diagrams for elevators with these (percentwise) dimensions, shown below:
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Figure XIV.3: These trim diagrams for NACA 1410 were taken from historical data
Preparations for malfunctions and extreme situations
It is very important to consider potential issues that may occur during flight when
working the plane’s stability.
Engine failure
In a situation where one of the exterior engines fails resulting in only one of the wing
engines proving thrust, a large moment which makes the plane unstable in terms of
yaw. This particular problem can be directly resolved by restarting the center turbofan.
This eliminates the stress of single outboard engine yaw on the structure of the aircraft.
In the event of the center engine failure, the outboard engines will need to provide
adequate thrust to gain maneuverable altitude and land, as this would be the only
phase of flight affected by such a failure.
Propulsion analysis in the “Results of engine modeling” section of the “VI. Results of
aircraft sizing” portion of the report (above) discovered that this engine alone could
produce the thrust required for safe controlled flight long enough to make an
emergency landing. In the event that the center engine is to lose thrust, the two wing
engines could also perform all maneuverability requirements of the airplane. In this
case, assuming the aircraft would need to rise to a height of 25,000 ft (in the event of a
missed landing, for example) the smaller engines provide the necessary thrust to do so.
Even though the fuel consumption and time to climb would be increased, the aircraft
would be able to adequately perform.
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Extreme crosswind
Another major concern was that of ‘Crosswind landing’. In this situation, a crosswind
which alters the aircraft’s orientation is applied. Such an impact makes yaw control
much more difficult, which is especially a problem when the jet prepares for landing.
The control surfaces must be sized such as to be able to adequately combat the induced
sideslip, and responsive enough to make the final adjustments to align properly with the
runway
Historical values taken from Dan Raymer’s text book were used to size the rudder and
ailerons. The rudder has a span that is 70% of the vertical stabilizer’s and a mean chord
length that is 30% of vertical stabilizer’s. Both are 7.07 ft, giving the rudder a span of
4.95 ft and a chord of 2.12 ft. The planform area is 10.5 ft2.
The total span is 51.82 ft, and the ailerons are sized such that they comprise 37% of the
span. They were sized to have chords that are 20% the length of the wings’ mean
aerodynamic chord, which, as stated earlier, is 6.6 ft. As a result, the ailerons are
designed to have 9.6 ft spans and 1.33 ft chords. They each have planform areas of
12.73 ft2 and aspect ratios of 7.22.
 - XV. Noise
Aircraft noise is produced by aircraft during all phases of the flight; on the ground when
parked, taxiing, during takeoff, landing, and while en route. One of the primary goals of
the design is to reduce this noise pollution as much as possible.
As the engines will undoubtedly be the largest noise-producing fixtures on the aircraft,
special consideration was given to their type (propfan, turbofan, or turboprop), their
shrouding, and their orientation on the aircraft. The two stock HF120 turbofans are
rated as individually compliant with the Stage IV guidelines, and it is expected that the
single 6000 lb thrust-scaled variant will operate in a decibel band within a reasonable
margin of those same guidelines. The shrouds for all three engines is intended to muffle
a considerable portion of the noise produced, though at the higher-frequencies
produced at takeoff they are likely to be less effective.
The EPNL noise recording for any airborne platform is measured at three distinct points
to reflect three important phases of flight: sideline, takeoff flyover, and landing flyover.
Reducing takeoff distance and time to climb considerably decreases the Endeavour’s
takeoff flyover noise signature, while the housings for all three engines are designed to
reduce sideline noise. The overall noise signature estimate of the aircraft would be
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Team XG Intl. Inc
expected to fall between 200 and 220 decibels (cumulative). Specific data is available
only on the two stock HF120 engines themselves and then scaled for the larger variant,
plus aerodynamic noise attributable to the landing gear, flaps, and control surfaces
accounted for by similarly-sized competitors and historical estimates.
The two smaller engines are mounted rearward on the horizontal stabilizer to reduce
pressure-wave noise in the cabin while also allowing for a possible future retrofit with
fuel-saving propfan engines. A significant drawback of the propfan engine is its highamplitude dynamic frequency vibration, and this is one of the reasons it was not
pursued for development in conjunction with the Endeavour project. The placement of
the largest engine in the baffles of the aircraft acts to reduce sideline and ground noise
during takeoff, where its noise output will be highest, but this compounds the issue of
internal vibration and noise.
To reduce the amount of oscillatory motion transferred to the cabin and passengers, the
aircraft is outfittable with an Active Vibration Control System, placed on the structural
engine mounts in the case of the internal engine, and at the root of the horizontal
stabilizer in the case of the two external engines. Such a system operates by producing
noise/vibration at opposing frequencies to its input, effectively dampening up to 90
percent of the vibration and noise given off by the engine. This system will be especially
necessary in tandem with the propfan system but adds weight and was deemed a luxury
in the face of industry data which indicated passengers expect some measure of engine
noise during all stages of flight. A notable feature of the AVC system is that, in addition
to the overall dampening it is capable of providing, it also is tunable in that a specific
region or compartment of the aircraft can be effectively quieted to near zero vibration
at the expense of weaker performance in the other regions of the aircraft. This will allow
for absolutely quiet video-conferences and maximally comfortable VIP accommodations
in-flight.
 - XVI. Cost
To estimate the cost to develop and manufacture the proposed aircraft, the Rand
Corporation’s “Development and Procurement Costs of Aircraft model” (DAPCA IV)
model was used from Raymer’s textbook. This model has been used primarily on
fighters, bombers, and transports but can apply to a broader class of aircraft. The
DAPCA IV model guesses the hours required for Research, Development, Test, and
Evaluation (RDT&E) and production by the engineering, manufacturing, tooling, and
quality control groups. This takes into account everything from technology research to
ground testing.
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The four above control groups are multiplied by the hourly rates to yield costs. Table
XVI.1 below shows a table summary of those hourly costs that are used.
Table XVI.1: Hours Estimated for DAPCA IV.
Engineering hourly rate
$86
Tooling hourly rate
$88
Quality control hourly rate
$81
Manufacturing hourly rate
$73
Given equations were used to calculate the four control groups along with the following:
 development support cost,
 flight test cost,
 manufacturing materials cost,
 engine production cost
 avionics cost.
Together, these costs form the RDT&E + flyaway cost. This total cost includes all aircraft
being produced. Based on the definition given in the textbook and how many aircraft
could actually be produced in 5 years, 160 aircraft were determined to be produced in
that 5 year period. A summary of the results is shown in Table XVI.2 below.
Table XVI.2: Cost Prediction Summary
Type
Price
RTD&E + Flyaway Cost
Production Cost
Profit Per Aircraft
Breakeven Point
Production Run
$2.4 Billion
$15 Million
$750,000
20 aircraft
160 units in 5 years
The RTD&E + flyaway cost was determined to be $2.4 billion. This is the cost for the
entire fleet of 160 aircraft to be built in a five year period. This translates to a
production cost of nearly $15 million per aircraft built. Keep in mind; this doesn’t
include new technology factors that may need to be added into development.
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Team XG Intl. Inc
The breakeven point along with the profit produced per aircraft was where the DAPCA
IV seemed to struggle to capture the correct values. The equation used to calculate this
was the production cost divided by the profit per aircraft. Or simply;
𝐵𝑟𝑒𝑎𝑘𝑒𝑣𝑒𝑛 𝑝𝑜𝑖𝑛𝑡 =
𝐶𝑜𝑠𝑡 𝑜𝑓 𝑃𝑟𝑜𝑑𝑢𝑐𝑡𝑖𝑜𝑛
𝑃𝑟𝑜𝑓𝑖𝑡 𝑝𝑒𝑟 𝑎𝑖𝑟𝑐𝑟𝑎𝑓𝑡 𝑠𝑜𝑙𝑑
Equation XVI.2
If the profit was raised then the breakeven point became a seemingly unreasonable
value.
From the presentation, Gulfstream mentioned that normal business aircraft turn a 30%
profit while our aircraft was near 5%. To get the profit to 30%, the breakeven point
would become absurdly low which doesn’t make reasonable sense. This could be due to
the DAPCA IV model using empty weight as a main input into the equations and the
aircraft designed has a low empty weight. After all, the DAPCA IV is just a model so it
need not apply to all aircraft. This is the drawback from using given equations and
historical data. Because of these issues, the profit was determined to be $750,000 an
aircraft producing a good breakeven point of 20 aircraft.
Also applying to the cost spectrum is the variable cost and fixed costs of the aircraft.
Table XVI.3 shows the variable operating cost. Of special note, a two man crew
equation was applied from the textbook and the current fuel/oil costs were factored in
to produce the results. The direct operating cost per seat-mile was found to be $.22.
This seems reasonable as the business jet is not meant to load up on passengers. A
business jet is more for traveling in comfort and on a strict time schedule without
commercial airline delays.
Table XVI.3: Variable operating cost
Rates
Depreciation
Insurance
Crew
Fuel/Oil
Maintenance
DOC
DOC/Seat-Mile
Values
6.6% / year
$30,000 / year
$230 / block hour
$1158/flight hour
$764 / flight hour
$2274 / flight hour
$0.22
Some of the determined fixed costs are shown below in Table XVI.4. These costs are
going to be on the consumer that purchases the aircraft. The costs shown below are
educated guesses on how much it will be to put the aircraft up in a hangar along with
training and landing fee per certain airports which is based on weight. This list is not all
inclusive but it gives a good idea of what is to be expected. Overall, the determined
costs seemed reasonable when applied to the business aircraft model.
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Team XG Intl. Inc
Table XVI.4: Fixed Operating Costs
Type
Hangar
Training
Landing
Cost
$80,000 / year
$40,000 / year
$386 / landing
 - XVII. Summary
XG International, Inc. sought to create a cutting edge, environmentally sensitive yet fast
and convenient air transportation system. The smartest minds available working with
state-of-the-art technologies have answered with the XG Endeavour. This business jet
maximizes proven designs with some of the latest advancements in noise and fuel
consumption reduction and power conservation to offer the ideal, small-size, mid-range
business jet. The compliance matrix can be found in appendix c.
Through much research and heavy consideration of aircraft designs and features, XG
International decided upon a standard body aircraft with swept back, low wings, and a
cruciform horizontal stabilizer. The XG Endeavour’s three GE Aero Honda HF120 variant
turbofan engines, two of which are located on the ends of the horizontal stabilizer, and a
third up-scaled model located inside aft of the aircraft body, enable flight speed of Mach
0.8 and a range of 3,700 nautical miles. The majority of the cabin layout of the aircraft
will be set to the customer’s specifications, which makes the XG Endeavour an elite form
of sky transportation.
The relatively small HF120 turbofans and the aerodynamic design make the XG
Endeavour more than compliant with NASA N+2 requirements. The XG Endeavour is
able to reduce fuel consumption by using all three of its turbofan engines during first
and second segment climb, then turning the central engine off and relying solely on its
two external engines for the cruise. The use of solar film furthers the XG Endeavour’s
environmental friendliness. Though the XG Endeavour does not meet its target takeoff
distance of 3,400 ft., its actual takeoff distance of 4,000 ft. allows access to nearly any
small airport that the customer may desire.
The XG Endeavour boasts a cruising speed of Mach 0.80, which will cover its 3,700
nautical mile range with haste. Add that to a takeoff distance of a mere 4,000 feet, and
this business jet is as versatile as any other jet in its class. The two horizontal stabilizermounted engines are the GE Honda Aero HF120 variant turbofans. Each of these
engines is capable of producing 2,000 pounds of thrust. The aft fuselage engine is an
up-scaled model of the HF120, capable of 6,800 pounds of thrust. All three of these
engines will be used for takeoff and the climb to the XG Endeavour’s cruising altitude,
where the
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Team XG Intl. Inc
APPENDIX A: Bibliography
Frei, Stefan. Urban Mader. “Technology Speed of Civil Jet Engines”, Techzoom
publications. 2006
Findlay, S.J., N.D. Harrison. “Why Aircraft Fail.” Material Study Magazine. November
2002. QinetiQ Ltd. Famborogh, Hampshire. pp 18-25.
Heet, Erika. ‘FrontRunners: Above and Beyond.’ Robb Report. Ed. April 01, 2010. New
York, NY. pp 23-24
Megson, T.H.G. “Aircraft Structures for Engineering Students.” Third edition. ‘Part II:
Aircraft Structures’. Butterworth Heinemann, Burlington, MA. 1999.
Raymer, Daniel P. “Aircraft Design: A conceptual Approach” Fourth Edition, American
Institute of Aeronautics and Astronautics, Inc. Virginia, 2006.
Tamura, Marcelo Satoru. “A Simplified Geometric Method for Wing Loads Estimation”
2009 Brizilian Symposium on Aerpsoace Eng. & Applications. September 14-16,
2009. S. J. Campos, SP, Brazil.
Tien, John K., and Caulfield, Thomas, eds. “Superalloys, Supercomposites and
Superceramics”. New York: Academic Press, 1989.
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APPENDIX B: Sizing code
See attached file.
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APENDIX C: Compliance Matrix
Requirement
Target Threshold
XG
Endevour
Compliant
Maximum Mach Number
0.85
0.8
0.82
No
Empty Weight (lb)
18,500
20,000
16,000
Yes
Gross Weight (lb)
28,000
32,000
29,100
No
Takeoff Distance (ft)
3,400
3,800
4,000
No
Maximum Range (nmi)
3,700
3,600
3,700
Yes
Design Mission Range (nmi)
3,700
3,600
3,700
Yes
Noise (dB)
42
50
>42
Yes
Seats
10
8
9
Yes
Volume Per Passenger (ft3)
65
60
60
Yes
TSFC (% of avg)
55
65
65
Yes
N0X Emissions (% of avg.)
25
50
10
Yes
Charge Time - 220V 80A* (hr)
2
4
1.5
Yes
Charge Time - 125V 15A**
(hr)
3
5
4
Yes
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